NASDA-HDBK-1007D CB H-IIA User’s Manual
Second Edition December 2001
Published by NATIONAL SPACE DEVELOPMENT AGENCY OF JAPAN PREFACE
This H-IIA User’s Manual presents information regarding the H-IIA launch vehicle and its related systems and launch services. This document contains information for launch services including mission performance capability, environmental conditions, spacecraft and launch vehicle interface conditions, launch operations and interface management. A brief description of the H-IIA launch vehicles and the launch facilities of Tanegashima Space Center is also included. As the H-IIA program is progressing, this document is subject to change and will be revised periodically. Requests for further information or inquiries related to this manual or interfaces between spacecraft and the H-IIA launch system should be addressed to:
National Space Development Agency of Japan
Mission Operations Department Office of Space Transportation Systems
World Trade Center Building 26F 2-4-1 Hamamatsu-cho, Minato-ku Tokyo, 105-8060 Japan Telephone + 81-3-3438-6465 Fax + 81-3-5402-6527
© NASDA This manual cannot be copied, duplicated, or quoted in part or whole (including drawings and photographs) without permission from NASDA.
0-1 H-IIA User’s Manual revision control sheet
Revision Date Revised page Approved A F7irst J—une. 199 ———— ———
All pages are redesigned. Major revised items : S7econd Dec. 199 - 4.6.9 RF Link Interface - Table A2.5.1
Tsechnicals Other
Contents 0-1 ~ 0-15 0-17 etc. 0-18 0-19 0-21 Tsechnicals Other 1-2 1-3 1-3 1-4 1-4 1-7 Chapter 1 1-5 1-9 1-8 1-10 1-11 1-13 1-16 ~ 1-20 1-15 1-22 1-22 Second Dec. 1998 1-23 Revision A Tsechnicals Other 2-1 ~ 2-12 2-1 2-15 2-3 Chapter 2 2-17 ~ 2-23 2-4 2-8 2-9 2-13 Tsechnicals Other 3-3 3-1 3-5 3-2 Chapter 3 3-11 3-3 3-12 3-5 ~ 3-11 3-13 3-21 3-23 3-24 3-24 3-25
0-2 Revision Date Revised page Approved A Tsechnicals Other 4-6 ~ 4-18 4-1 4-26 ~ 4-33 4-2 4-38 ~ 4-43 4-41 4-48 4-53 4-52 4-54 Chapter 4 4-56 4-55 4-59 4-57 4-60 4-58 4-61 4-63 4-64 ~ 4-68 4-69 4-71 4-74 4-72 4-73 Tsechnicals Other Second Dec. 1998 Revision A 5-1 5-1 5-2 5-3 5-4 5-5 5-7 5-6 5-10 5-7 5-12 5-11 5-14 5-18 Chapter 5 5-15 5-19 5-20 5-21 5-23 5-25 5-26 5-28 5-29 5-31 5-32
0-2-10-3 Revision Date Revised page Approved A Tsechnicals Other 6-1 6-3 6-2 6-4 Chapter 6 6-10 ~ 6-14 6-5 6-19 6-9 6-20 6-14 6-19 6-22 Tsechnicals Other Appendix 1 A1-2
Tsechnicals Other A2-9 A2-4 A2-10 A2-5 A2-15 A2-6 Second Dec. 1998 A2-16 A2-12 Revision A A2-17 A2-13 Appendix 2 A2-20 ~ A2-21 A2-17 A2-24 A2-18 A2-27 A2-21 A2-22 A2-23 A2-25 A2-26 A2-28 ~ A2-32 Tsechnicals Other All Appendix 3.3 (except the A3.5-2 Appendix 3 pages A3.5-3 mentioned in A3.5-6 ~ 9 the right table) Appendix 3.8 Appendix 3.9
0-2-20-4 Revision Date Revised page Approved B Tsechnicals Other Cover 0-2-3 0-3 Contents 0-10 0-11 etc. 0-12 0-14 0-16 0-17 0-18 Tsechnicals Other
Chapter 1 1-4 1-13 1-8 1-15 1-19 Tsechnicals Other 2-2 2-8 Second Dec. 1999 2-5 2-15 Revision B 2-7 2-16 2-10 2-24 2-11 2-25 2-12 Chapter 2 2-13 2-14 2-15 2-16 2-18 2-19 2-22 2-23 Tsechnicals Other 3-5 3-1 3-12 3-2 Chapter 3 3-14 3-9 3-10 3-13 3-15
0-2-30-5 Revision Date Revised page Approved B Tsechnicals Other 4-3 4-36 4-24 4-56 4-38 4-39 4-40 4-41 4-42 4-53 Chapter 4 4-54 4-55 4-57 4-63 4-64 4-65 4-67 4-68 4-69 Second Dec. 1999 4-70 Revision B Tsechnicals Other 5-10 5-18 Chapter 5 5-20 5-11 5-31 5-32 Tsechnicals Other Chapter 6 6-9 6-10 Tsechnicals Other Appendix 1 A1-1 A1-2 Tsechnicals Other
Appendix 2 A2-12 A2-22 A2-26
0-60-2-4 Revision Date Revised page Approved B Tsechnicals Other A3.1-2 A3.1-1 A3.1-7 A3.1-6 A3.2-1 A3.2-6 A3.2-2 A3.3-1 A3.2-3 A3.4-6 A3.2-7 A3.5-1 A3.4-1 A3.5-9 A3.4-2 A3.6-1 A3.4-3 A3.7-1 Second D3ec. 1999 Appendix A3.4-7 A3.8-1 Revision B A3.5-10 A3.9-1 A3.6-2 A3.6-5 A3.6-6 A3.6-9 A3.7-2 A3.7-5 A3.7-6 A3.7-9 A3.9-1 A3.9-2
0-2-50-7 Revision Date Revised page Approved C Tsechnicals Other Contents etc. 0-3~6 0-10~19 Tsechnicals Other
Chapter 1 1-4~7 1-1~3 1-18~19 1-13 1-20~21 Tsechnicals Other 2-1 Chapter 2 2-7 21-14~2 Tsechnicals Other 3-12 Chapter 3 3-16~18 3-21~22 Second 3-26 Dec. 2001 Revision C Tsechnicals Other 4-4~8 4-24~28 4-21~23 4-30~36 Chapter 4 4-29 4-46~49 4-37~45 4-54~56 4-50~51 4-58~59 4-71 Tsechnicals Other 5-1 Chapter 5 5-4 5-7 5-17 5-19 Tsechnicals Other Chapter 6 6-10 6-3 6-12 6-9
0-2-60-8 C Revision Date Revised page Approved Tsechnicals Other Appendix 1 A11-2 A1- Tsechnicals Other
Appendix 2 A62-25 A2- A2-11 A2-20 Tsechnicals Other A3.1-1~2 A3.1-3~5 A3.1-6 A3.2-4~5 A3.2-1~3 A3.4-2 Second Dec. 2001 A3.2-6~7 A3.4-4~5 Revision C A3.3-1 A3.4-7 A3.3-3 A3.5-4 A3.3-5~7 A3.6-2 Appendix 3 A3.4-1~3 A3.6-5 A3.4-6 A3.7-2 A3.5-1 A3.5-4 A3.6-1 A3.7-1 A3.8-1 A3.9-1 A3.10-1~10 A3.11-1~4
0-2-70-9 CONTENTS
CHAPTER 1. INTRODUCTION 1.1 Purpose of the User’s Manual ...... 1-1 1.2 H-IIA Launch System ...... 1-1 1.2.1 H-IIA launch vehicle ...... 1-1 1.2.2 Launch facilities ...... 1-4 1.2.3 Payload accommodations ...... 1-5 1.2.4 Users / NASDA relationship ...... 1-5 1.2.5 Advantages of H-IIA ...... 1-5 CA 1.3 H-IIA Launch System Related Documents ...... 1-22 1.3.1 H-IIA Payload-Related Facilities and GSE Manual (in preparation) .... 1-22 1.3.2 Launch Vehicle Payload Safety Requirements ...... 1-22 1.4 Definition of terms ...... 1-22
CHAPTER 2. MISSION PERFORMANCE 2.1 General ...... 2-1 2.1.1 Mission profile ...... 2-1 2.1.1.1 Booster and first stage phase ...... 2-1 2.1.1.2 Second stage phase ...... 2-2 A 2.2 Performance Ground Rules ...... 2-7 2.2.1 Payload mass definition ...... 2-7 A 2.2.2 Launch vehicle configurations ...... 2-7 2.2.3 Launch vehicle performance confidence levels ...... 2-7 2.3 Geostationary Transfer Orbit (GTO) Mission ...... 2-7 2.3.1 Payload capability for single launch...... 2-7 2.3.2 Payload capability for dual (GTO and GTO) launch ...... 2-8 2.3.3 Typical orbital parameters ...... 2-8 2.3.4 Injection accuracies ...... 2-8 2.3.5 Typical sequence of events...... 2-8 2.3.6 Typical trajectory ...... 2-8 B 2.3.7 Typical flight parameters ...... 2-9 2.4 Sun-Synchronous Orbit (SSO) Mission ...... 2-16 C 2.4.1 Payload capability ...... 2-16 2.4.2 Typical orbital parameters ...... 2-16 2.4.3 Injection accuracies ...... 2-16 2.4.4 Typical trajectory ...... 2-16 2.5 Low Earth Orbit (LEO) Mission ...... 2-18 2.5.1 Payload capability ...... 2-18 2.6 Earth Escape Mission ...... 2-21 2.6.1 Payload capability ...... 2-21
0-3 2.7 Spacecraft Orientation and Separation ...... 2-22 2.7.1 General description ...... 2-22 2.7.2 Separation sequence ...... 2-22 2.7.3 Spin-up performance ...... 2-22 2.7.4 Pointing accuracy ...... 2-22 2.7.5 Relative separation velocity...... 2-23 2.7.6 Separation tip-off rate ...... 2-23 2.7.7 Dual launch sequence ...... 2-23
CHAPTER 3. ENVIRONMENTS 3.1 General ...... 3-1 3.2 Mechanical Environments ...... 3-1 3.2.1 General ...... 3-1 3.2.2 Combined load factor ...... 3-1 A 3.2.3 Sinusoidal vibration ...... 3-2 3.2.4 Random vibration ...... 3-2 3.2.5 Acoustics ...... 3-2 C 3.2.6 Shock ...... 3-2 3.3 Thermal Environment ...... 3-11 3.3.1 General ...... 3-11 3.3.2 Prelaunch environment ...... 3-11 3.3.3 Launch and flight environment ...... 3-12 3.4 Fairing Internal Pressure Environment ...... 3-18 C 3.5 Contamination and Cleanliness...... 3-23 3.5.1 Prelaunch contamination and cleanliness ...... 3-23 3.5.2 Flight contamination control ...... 3-23 3.6 Radiation and Electromagnetic ...... 3-26 3.6.1 Launch vehicle generated radio environment ...... 3-26 3.6.2 LV generated electromagnetic environment ...... 3-26 3.7 Spacecraft Compatibility Test Requirements ...... 3-26
CHAPTER 4. INTERFACE REQUIREMENTS 4.1 General ...... 4-1 4.2 Frequency Requirements...... 4-1 4.2.1 General ...... 4-1 4.2.2 Fundamental frequencies ...... 4-1 4.3 Balance Requirements...... 4-3 4.3.1 General ...... 4-3 4.3.2 Height limit of the center of gravity ...... 4-3 4.3.3 Balance requirements ...... 4-3
0-4 4.3.3.1 Static balance ...... 4-3 4.3.3.2 Dynamic balance...... 4-3 4.4 Payload Fairing ...... 4-3 4.4.1 Fairing types ...... 4-4 4.4.1.1 Model 4S ...... 4-4 4.4.1.2 Model 5S ...... 4-4 4.4.1.3 Model 4/4D-LS...... 4-4 4.4.1.4 Model 4/4D-LC ...... 4-5 4.4.1.5 Model 5/4D ...... 4-5 4.4.1.6 Model 5S-H...... 4-5 C 4.4.2 Stay out zone around the payload adapter ...... 4-5 4.4.3 Large door ...... 4-5 4.4.4 Mission modification ...... 4-6 C 4.4.4.1 Access door ...... 4-6 4.4.4.2 Umbilical connectors ...... 4-6 4.4.4.3 Transparent window ...... 4-6 4.4.4.4 Internal antenna...... 4-7 4.4.4.5 Separate air conditioning ...... 4-7 4.4.4.6 Acoustic blankets ...... 4-7 4.5 Payload Adapter ...... 4-49 C 4.5.1 Configuration of payload adapter ...... 4-49 4.5.1.1 Separation mechanism ...... 4-49 4.5.1.2 Ejection mechanism ...... 4-49 4.5.1.3 Spacecraft separation monitoring switches ...... 4-49 4.5.1.4 Nomenclature of payload adapter ...... 4-50 4.5.2 Adapter types ...... 4-50 4.5.3 Mission modifications ...... 4-50 4.5.3.1 Separation springs...... 4-50 4.5.3.2 Umbilical connectors ...... 4-50 4.6 Electrical and RF Interface ...... 4-52 4.6.1 General...... 4-52 4.6.2 Electrical grounding ...... 4-52 4.6.3 Umbilical interface ...... 4-52 4.6.3.1 Umbilical lines for single launch mission ...... 4-52 4.6.3.2 Umbilical lines for dual launch mission ...... 4-52 4.6.4 Command and power interface ...... 4-53 4.6.4.1 Pyrotechnic command ...... 4-53 4.6.4.2 Electrical command (discrete signal) ...... 4-54 4.6.4.3 Dry loop command ...... 4-54 4.6.4.4 Power supply ...... 4-55 4.6.5 In-flight telemetry ...... 4-55
0-5 4.6.5.1 Separation status transmission ...... 4-55 C 4.6.5.2 Dynamic environments data transmission ...... 4-55 4.6.6 Interface connectors between spacecraft and launch vehicle ...... 4-56 4.6.6.1 Interface connectors procurement responsibility ...... 4-56 4.6.6.2 Interface connectors for single launch ...... 4-56 4.6.6.3 Interface connectors for dual launch ...... 4-56 4.6.6.4 Standard interface connector specifications ...... 4-56 4.6.6.5 Other interface connector characteristics ...... 4-57 4.6.7 RF constraints ...... 4-58 4.6.7.1 Fairing transparency for spacecraft RF communications ...... 4-58 4.6.7.2 Operating constraints ...... 4-58 4.6.8 Electrical and RF requirements for launch phase ...... 4-58 4.6.8.1 Electrical requirements ...... 4-58 4.6.8.2 RF requirements ...... 4-58 4.6.9 RF link interface ...... 4-59 4.6.9.1 General ...... 4-59 4.6.9.2 RF link with ML / STA2 ...... 4-59 4.6.9.3 RF link with MTCS ...... 4-60 4.7 Other Ground Equipment Interfaces ...... 4-75 4.7.1 Power ...... 4-75 4.7.2 Liquids and gases ...... 4-75 4.7.3 Propellant / gas sampling and analyzing ...... 4-75 4.7.4 Filling equipment room ...... 4-75
CHAPTER 5. LAUNCH OPERATIONS 5.1 General ...... 5-1 5.1.1 Scope ...... 5-1 5.2 Overview of the launch-related organizations ...... 5-1 5.3 Launch Operations Requirements ...... 5-4 A 5.3.1 Safety requirements ...... 5-4 5.3.2 Launch operations interface requirements ...... 5-4 5.4 Responsibility and Organization ...... 5-4 5.4.1 Launch operations organization ...... 5-4 5.4.2 Responsibility ...... 5-5 5.5 Restrictions ...... 5-8 5.5.1 Restrictions on the ground ...... 5-8 5.5.2 Restrictions on launching ...... 5-8 5.5.2.1 Launch window ...... 5-8 5.5.2.2 Launch postponement ...... 5-9 5.6 TNSC Facilities and GSE Related to Launch Operations of Spacecraft ...... 5-11
0-6 5.7 Launch Operations...... 5-13 5.7.1 Spacecraft-related operations and programs ...... 5-13 5.7.2 Phase 1 (spacecraft preparation and functional test) ...... 5-13 5.7.3 Phase 2 (hazardous operations for spacecraft) ...... 5-14 5.7.3.1 Preparing and assembling pyrotechnics and solid motor ...... 5-14 5.7.3.2 Spacecraft operations...... 5-15 5.7.3.3 Final spacecraft assembly ...... 5-15 5.7.4 Phase 3 (joint operations by spacecraft and launch vehicle organizations)...... 5-16 5.7.4.1 Encapsulation into the payload fairing ...... 5-16 5.7.4.2 Encapsulation for single launch ...... 5-16 5.7.4.3 Encapsulation for dual launch ...... 5-17 5.7.4.4 Transportation of encapsulated spacecraft...... 5-17 5.7.4.5 Mating with launch vehicle...... 5-18 5.7.4.6 Spacecraft inspection after installation ...... 5-18 5.7.4.7 Y-3 operation ...... 5-18 A 5.7.4.8 Y-2 ~ Y-0 operation ...... 5-18 5.7.4.9 Terminal countdown...... 5-19 5.7.4.10 Recycle operations (Launch postponement) ...... 5-20
CHAPTER 6. INTERFACE MANAGEMENT 6.1 General ...... 6-1 6.1.1 Launch service organization...... 6-1 6.1.2 Interface management document...... 6-1 6.1.2.1 Documents to be submitted by the spacecraft organization ...... 6-1 6.1.2.2 Document to be submitted by NASDA ...... 6-2 6.2 Interface Work with Spacecraft Organization ...... 6-2 6.2.1 Interface schedule / Interface items ...... 6-2 6.2.1.1 Standard mission ...... 6-2 6.2.2 Mission analysis ...... 6-3 6.2.2.1 Trajectory analysis ...... 6-3 6.2.2.2 Orbit dispersion analysis ...... 6-3 6.2.2.3 Sun angle analysis ...... 6-3 6.2.2.4 Spacecraft separation analysis...... 6-3 6.2.2.5 Relative orbit analysis of vehicle and spacecraft ...... 6-4 6.2.2.6 Spacecraft coupled loads analysis ...... 6-4 6.2.2.7 Radio frequency compatibility study ...... 6-4 6.2.2.8 Integrated thermal analysis ...... 6-4 A 6.2.3 Interface test...... 6-4 6.2.3.1 Fit check of spacecraft and PLA ...... 6-4 6.2.3.2 Umbilical connector disconnection test ...... 6-5
0-7 6.2.3.3 Separation shock test ...... 6-5 6.2.4 Mission modification ...... 6-5 6.2.4.1 Payload fairing mission modification ...... 6-5 6.2.4.2 Payload adapter mission modification ...... 6-6 6.2.5 Standard services and optional services ...... 6-6 6.2.5.1 Distinction between standard and optional service items ...... 6-6 6.2.5.2 Notes ...... 6-6 6.3 Spacecraft / H-IIA and Spacecraft / Launch Operations Interface Control ...... 6-15 A 6.3.1 Interface control document ...... 6-15 6.3.2 Coordination items and timing ...... 6-15 6.4 Mission Analysis...... 6-17 6.4.1 General...... 6-17 6.4.2 Reference planning phase ...... 6-17 6.4.3 Detailed planning phase ...... 6-18 6.4.4 Final planning phase ...... 6-18 6.4.5 Spacecraft coupled loads analysis (CLA)...... 6-18 6.4.6 Integrated thermal analysis ...... 6-19 A 6.5 Safety Reviews ...... 6-19 6.5.1 Payload safety requirements ...... 6-19 6.5.2 Requirements on the safety program plan of the spacecraft organization ...... 6-19 6.5.3 Safety review data package ...... 6-20 6.5.4 Outline of NASDA safety review ...... 6-20 6.5.5 Interface ...... 6-20 6.6 Reviews and Other Meetings ...... 6-20 6.6.1 Reviews before launch operation ...... 6-20 6.6.1.1 Mission readiness review (MRR) ...... 6-20 6.6.1.2 Spacecraft interface confirmation review (SIC) ...... 6-20 6.6.2 Reviews for launch operation ...... 6-21 6.6.2.1 Launch vehicle readiness review (LVRR) ...... 6-21 6.6.2.2 Spacecraft readiness review (SCRR) ...... 6-21 6.6.2.3 Flight readiness review (FRR) ...... 6-21 6.6.3 Safety review ...... 6-21 6.6.4 Meetings ...... 6-22 6.6.4.1 Interface meeting ...... 6-22 6.6.4.2 Daily meeting ...... 6-22 6.6.4.3 Launch vehicle readiness meeting ...... 6-22 6.6.4.4 Precountdown coordination meeting
0-8 APPENDIX 1. HISTORY OF NASDA LAUNCH VEHICLES A A1.1 General ...... A1-1 A1.2 Abstract of The H-I Launch Vehicle ...... A1-1 A1.3 Abstract of The H-II Launch Vehicle ...... A1-1
APPENDIX 2. OUTLINE OF TANEGASHIMA ISLAND AND TANEGASHIMA SPACE CENTER A2.1 Tanegashima Island ...... A2-1 A2.1.1 Location and topography ...... A2-1 A2.1.2 Climate ...... A2-1 A2.1.3 Traffic ...... A2-1 A2.2 The Tanegashima Space Center ...... A2-4 A2.2.1 General ...... A2-4 A2.2.2 Osaki Launch Range ...... A2-4 A2.2.2.1 Yoshinobu Vehicle Assembly Building (VAB) ...... A2-4 A2.2.2.2 Movable Launcher (ML) ...... A2-5 A A2.2.2.3 Block House (B/H)...... A2-5 A2.2.2.4 No. 2 Spacecraft Test and Assembly Building (STA2)...... A2-5 A2.2.2.5 Spacecraft and Fairing Assembly Building (SFA)...... A2-5 A2.2.2.6 No. 1 Spacecraft Test and Assembly Building (STA1)...... A2-6 A2.2.2.7 Third stage and Spacecraft Assembly Building (TSA) ...... A2-6 A2.2.2.8 Non-Destructive Test Facility (NDTF)...... A2-6 A2.2.2.9 Nakanoyama Telemetry Command Station...... A2-6 A2.2.2.10 Other facilities ...... A2-6 A2.2.3 Takesaki Area ...... A2-7 A2.2.3.1 Administration Building (AB)...... A2-7 A2.2.3.2 Takesaki Range Control Center (RCC) ...... A2-7 A2.2.3.3 Takesaki Observation Stand ...... A2-8 A2.2.3.4 Takesaki Space Exhibition Hall ...... A2-8 A2.3 No. 2 Spacecraft Test and Assembly Building (STA2) ...... A2-13 A2.3.1 General ...... A2-13 A2.3.2 Main Structure and Functions of STA2 ...... A2-13 A2.4 Spacecraft and Fairing Assembly Building (SFA)...... A2-17 A2.4.1 General ...... A2-17 A2.4.2 Main Structure and Functions of SFA ...... A2-17 A2.5 Third Stage and Spacecraft Assembly Building (TSA)...... A2-20-1 A A2.5.1 General ...... A2-20-1 A2.5.2 Main Structure and Functions of TSA ...... A2-20-1 A2.6 Yoshinobu Vehicle Assembly Building (VAB) ...... A2-21 A2.6.1 General ...... A2-21
0-9 A2.6.2 Main Structure and Function of VAB ...... A2-21 A A2.7 Movable Launcher (ML) ...... A2-25 A2.7.1 General ...... A2-25 A2.7.2 Main structure and functions of ML ...... A2-25 A2.8 Takesaki Range Control Center (RCC)...... A2-28 A2.8.1 General ...... A2-28 A2.8.2 Main structure and functions of Takesaki RCC ...... A2-28 A2.9 Down range stations...... A2-30 A2.9.1 Ogasawara downrange station ...... A2-30 A2.9.2 Christmas downrange station ...... A2-30
APPENDIX 3. PAYLOAD ADAPTER APPENDIX 3.1 937M ADAPTER ...... A3.1-1 B APPENDIX 3.2 937MH ADAPTER...... A3.2-1 APPENDIX 3.3 937M-SPIN ADAPTER...... A3.3-1 APPENDIX 3.4 937M-SPIN-A ADAPTER ...... A3.4-1 APPENDIX 3.5 1194M ADAPTER ...... A3.5-1 APPENDIX 3.6 1666M ADAPTER ...... A3.6-1 APPENDIX 3.7 1666S ADAPTER ...... A3.7-1 APPENDIX 3.8 2360S ADAPTER ...... A3.8-1 APPENDIX 3.9 3470S ADAPTER ...... A3.9-1 APPENDIX 3.10 1666MA ADAPTER ...... A3.10-1 C APPENDIX 3.11 239M ADAPTER ...... A3.11-1
0-10 FIGURES
Figure 1.2.1 H-IIA Launch Vehicle Family ...... 1-6 C Figure 1.2.2 H-IIA (H2A202) launch vehicle configuration ...... 1-8 A Figure 1.2.3 H-IIA (H2A212) launch vehicle configuration ...... 1-9 Figure 1.2.4 Location of major facilities in Osaki Launch Range ...... 1-10 Figure 1.2.5 No. 2 Spacecraft test and assembly building (STA2) ...... 1-11 Figure 1.2.6 Spacecraft and fairing assembly building (SFA)...... 1-11 Figure 1.2.7 Vehicle assembly building (VAB) ...... 1-12 A Figure 1.2.8 Overview of New Yoshinobu Launch Complex ...... 1-13 Figure 1.2.9 H-IIA launch operations process ...... 1-14 Figure 1.2.10 Payload fairings for single launch with 1194M adapter ...... 1-15 Figure 1.2.11 (1/2) Payload fairing for dual launch with 1194M adapter ...... 1-16 A Figure 1.2.11 (2/2) Payload fairings for dual launch with 1194M adapter ...... 1-17 Figure 1.2.12 Payload fairing for single launch ...... 1-18 Figure 1.2.13 Payload adapters (example) ...... 1-19 Figure 1.2.14 Concept of users / NASDA relationship ...... 1-20 Figure 1.2.15 Concept of users / NASDA relationship after establishment of NASDA launch operations team ...... 1-21
Figure 2.1.1 Typical GTO mission profile for H2A2024 ...... 2-3 A Figure 2.1.2 Typical GTO mission profile for H2A212 ...... 2-4 Figure 2.3.1 Typical flight parameters for GTO mission (H2A202 with 4S fairing) ...... 2-10 Figure 2.3.2 Typical flight trajectory for GTO mission (H2A202 with 4S fairing) ...... 2-10 Figure 2.3.3 Typical flight parameters for GTO mission (H2A2022 with 4S fairing) ...... 2-11 Figure 2.3.4 Typical flight trajectory for GTO mission (H2A2022 with 4S fairing) ...... 2-11 Figure 2.3.5 Typical flight parameters for GTO mission (H2A2024 with 4S fairing) ...... 2-12 Figure 2.3.6 Typical flight trajectory for GTO mission (H2A2024 with 4S fairing) ...... 2-12 Figure 2.3.7 Typical flight parameters for GTO mission (H2A212 with 5S fairing) ...... 2-13 Figure 2.3.8 Flight trajectory for GTO mission (H2A212 with 5S fairing) ...... 2-13 Figure 2.3.9 Payload capability for GTO mission (H2A202 with 4S fairing) C
0-11 Figure 2.3.11 Payload capability for GTO mission (H2A2024 with 4S fairing) C
Figure 3.2.1 Typical longitudinal static acceleration for GTO (H2A202) ...... 3-4 Figure 3.2.2 Sound pressure level inside fairing with acoustic blanket (H2A202, H2A2022, H2A2024) ...... 3-6 A Figure 3.2.3 Sound pressure level inside fairing with acoustic blanket (H2A212) ...... 3-7 A Figure 3.2.4 Typical spacecraft separation shock spectrum with 1194M adapter ...... 3-9 A Figure 3.2.5 Typical spacecraft separation shock spectrum with 2360S adapter ...... 3-10 A
0-12 Figure 3.3.1 Internal surface temperature profiles of 4S or 4/4D upper fairing for GTO mission ...... 3-15 Figure 3.3.2 Internal surface temperature profiles of 5S or 5/4D upper fairing for GTO mission ...... 3-15 Figure 3.3.3 Internal surface temperature profiles of 5S-H fairing C for HTV mission ...... 3-16 Figure 3.3.4 Aerothermalheatluxofthefreemolecularflow (GTO mission) ...... 3-17 Figure 3.4.1 Typicalinthemapressureprofilefor GTO mission (Model 4S fairing) ...... 3-19 C Figure 3.4.2 Typical pressure decay rate profile for GTO mission (Model 4S fairing) ...... 3-20 C Figure 3.4.3 Typical internal pressure profile for SSO mission (Model 4S fairing) ...... 3-21 C Figure 3.4.4 Typical pressure decay rate profile for SSO mission (Model 4S fairing) ...... 3-22 C
Figure 4.2.1 H-IIA launch vehicle axes ...... 4-2 A Figure 4.4.1 Model 4S ...... 4-9 Figure 4.4.2 Usable volume of model 4S with 1194M adapter ...... 4-10 Figure 4.4.3 Model 5S ...... 4-11 Figure 4.4.4 Usable volume of model 5S with 1194M adapter ...... 4-12 Figure 4.4.5 Model 4/4D-LS ...... 4-13 Figure 4.4.6 Usable volume of model 4/4D long upper fairing with 1194M adapter ...... 4-14 Figure 4.4.7 Usable volume of model 4/4D short lower fairing with 1194M adapter ...... 4-15 Figure 4.4.8 Model 4/4D-LC ...... 4-16 Figure 4.4.9 Usable volume of model 4/4D clamshell lower fairing with 1194M adapter ...... 4-17 Figure 4.4.10 Model 5/4D ...... 4-18 Figure 4.4.11 Usable volume of model 5/4D lower fairing with 1194M adapter ...... 4-19 Figure 4.4.12 Usable volume of model 5/4D lower fairing with 1194M adapter ...... 4-20 Figure 4.4.13 Model 5S-H ...... 4-21 C Figure 4.4.14 (1/2) Usable volume of the model 5S-H #1 ...... 4-22 Figure 4.4.14 (2/2) Usable volume of the model 5S-H #2 ...... 4-23 Figure 4.4.15-1(1/3) General configuration of the 1194M adapter ...... 4-24 Figure 4.4.15-1(2/3) Stay-out zone around the 1194M adapter (I / III -axis ±20°) ...... 4-25
0-13 Figure 4.4.15-1(3/3) Stay-out zone around the 1194M adapter (II / IV -axis ±70°) ...... 4-26 C Figure 4.4.16 Large door ...... 4-27 Figure 4.4.17 ¿ 450 access door ...... 4-28 Figure 4.4.18 ¿ 600 access door ...... 4-29 Figure 4.4.19 Allowable areas of ¿ 450 access door on model 4S fairing ...... 4-30 Figure 4.4.20 Allowable areas of ¿ 450 access door on model 5S fairing ...... 4-31 Figure 4.4.21 Allowable areas of ¿ 450 access door on model 4/4D long upper fairing...... 4-32 Figure 4.4.22 Allowable areas of ¿ 450 access door on model 4/4D short lower fairing ...... 4-33 Figure 4.4.23 Allowable areas of ¿ 450 access door on model 4/4D clamshell lower fairing ...... 4-34 Figure 4.4.24 Allowable areas of ¿ 450 access door on model 5/4D upper fairing ...... 4-35 Figure 4.4.25 Allowable areas of ¿ 450 access door on model 5/4D lower fairing ...... 4-36 Figure 4.4.26 Allowable areas of ¿ 450 access door on model 5S-H fairing .... 4-37 Figure 4.4.27 Allowable areas of ¿ 600 access door on model 4S fairing ...... 4-38 Figure 4.4.28 Allowable areas of ¿ 600 access door on model 5S fairing ...... 4-39 Figure 4.4.29 Allowable areas of ¿ 600 access door on model 4/4D long upper fairing ...... 4-40 Figure 4.4.30 Allowable areas of ¿ 600 access door on model 4/4D short lower fairing ...... 4-41 Figure 4.4.31 Allowable areas of ¿ 600 access door on model 4/4D clamshell lower fairing ...... 4-42 Figure 4.4.32 Allowable areas of ¿ 600 access door on model 5/4D upper fairingr ...... 4-43 Figure 4.4.33 Allowable areas of ¿ 600 access door on model 5/4D lower fairing ...... 4-44 Figure 4.4.34 Allowable areas of ¿ 600 access door on model 5S-H fairing .... 4-45 Figure 4.4.35 Transparent window ...... 4-46 Figure 4.4.36 Typical Installation of internal antenna ...... 4-47 Figure 4.4.37 Typical Configuration of acoustic blanket ...... 4-48 Figure 4.6.1 Umbilical interfaces for single launch ...... 4-61 Figure 4.6.2 Umbilical interfaces for dual launch ...... 4-63 Figure 4.6.3 Pyrotechnic command wiring diagram ...... 4-64 Figure 4.6.4 Electrical command wiring diagram...... 4-65
0-14 Figure 4.6.5 Dry loop command wiring diagram ...... 4-66 C Figure 4.6.6 Interface connectors for single launch ...... 4-67 Figure 4.6.7 (1/2) Interface connectors for dual launch (via payload adapter) ...... 4-68 Figure 4.6.7 (2/2) Interface connectors for dual launch (via payload fairing) ...... 4-69 Figure 4.6.8 Acceptable spurious radiation levels ...... 4-71 Figure 4.6.9 VAB RF link schematic (for dual launch) ...... 4-72 Figure 4.6.10 RF link schematic during ML transfer (for dual launch) ...... 4-73 Figure 4.6.11 Launch pad RF link schematic (for dual launch) ...... 4-74
Figure 5.2.1 Launch operations organization of NASDA ...... 5-2 Figure 5.2.2 Relationship between the user and NASDA A after establishment of NASDA launch operations team ...... 5-3 Figure 5.4.1 Organization chart for launch operations (except for Y-2 ~ Y-0)...... 5-6 A Figure 5.4.2 Organization chart for Y-2 ~ Y-0 ...... 5-7 Figure 5.5.1 Typical launch postponement schedule ...... 5-10 Figure 5.6.1 Location of spacecraft-related buildings in TNSC’s Osaki Launch Range ...... 5-12 Figure 5.7.1 Typical spacecraft launch operations schedule ...... 5-21 Figure 5.7.2 Typical operations flow diagram for spacecraft preparation at TNSC ...... 5-22 Figure 5.7.3 Typical phase 1 operations flow diagram ...... 5-23 Figure 5.7.4 Typical phase 2 operations flow diagram ...... 5-24 Figure 5.7.5 Typical phase 3 operations flow diagram for single launch (for 4S fairing) ...... 5-25 Figure 5.7.6 Typical phase 3 operations flow diagram for dual launch ...... 5-26 Figure 5.7.7 Typical encapsulation sequence for single launch (for 4S fairing) ...... 5-27 Figure 5.7.8 Installation sequence for dual launch ...... 5-28 Figure 5.7.9 Transportation sequence of the encapsulation spacecraft ...... 5-29 Figure 5.7.10 Encapsulated spacecraft and launch vehicle mating ...... 5-30 Figure 5.7.11 Typical countdown schedule ...... 5-31 Figure 5.7.12 Typical launch vehicle system countdown schedule ...... 5-32
Figure 6.2.1 Typical spacecraft / H-IIA launch vehicle interface schedule for standard mission ...... 6-8
0-15 Figure A1.1.1 Configuration summary of NASDA launch vehicles ...... A1-3
Figure A2.1.1 Location of Tanegashima space center...... A2-3 Figure A2.2.1 Location of Facilities of Tanegashima Space Center ...... A2-9 Figure A2.2.2 Location of major facilities in Osaki Launch Range ...... A2-10 Figure A2.2.3 Overview of New Yoshinobu Launch Complex ...... A2-11 C Figure A2.2.4 H-IIA launch operations process ...... A2-12 A Figure A2.3.1 No. 2 Spacecraft Test and Assembly Building (STA2)...... A2-14 Figure A2.4.1 Spacecraft and Fairing Assembly Building (SFA) ...... A2-19 A Figure A2.5.1 Third Stage and Spacecraft Assembly Building (TSA) ...... A2-20-2 Figure A2.6.1 Yoshinobu Vehicle Assembly Building (VAB) ...... A2-22 Figure A2.6.2 Yoshinobu Vehicle Assembly Building (VAB, sectional view)...... A2-23 Figure A2.7.1 Movable Launcher (ML) ...... A2-26 Figure A2.8.1 Takesaki Range Control Center (RCC) ...... A2-29 Figure A2.9.1 Ogasawara downrange station...... A2-32 Figure A2.9.2 Christmas downrange station ...... A2-32
Figure A3.1.1 Photograph of the 937M adapter ...... A3.1-1 C Figure A3.1.2 General view of the 937M adapter...... A3.1-2 Figure A3.1.3 Details of the 937M adapter...... A3.1-3 Figure A3.1.4 Stay-out zone around the 937M adapter ...... A3.1-4 Figure A3.1.5 Limit load of the 937M adapter ...... A3.1-5 Figure A3.1.6 Spacecraft separation shock spectrum with the 937M adapter ...... A3.1-6 B Figure A3.1.7 Limit loads at separation plane of the 937M adapter ...... A3.1-7 C Figure A3.2.1 Photograph of the 937MH adapter ...... A3.2-1 Figure A3.2.2 General view of the 937MH adapter ...... A3.2-2 Figure A3.2.3 Details of the 937MH adapter ...... A3.2-3 Figure A3.2.4 Stay-out zone around the 937MH adapter...... A3.2-4 Figure A3.2.5 Limit load of the 937MH adapter...... A3.2-5 Figure A3.2.6 Spacecraft separation shock spectrum with the 937MH adapter ...... A3.2-6 Figure A3.2.7 Limit loads at separation plane of the 937MH adapter ...... A3.2-7 Figure A3.3.1 Photograph of the 937M-SPIN adapter ...... A3.3-1 A Figure A3.3.2 General view of 937M-SPIN adapter ...... A3.3-2 Figure A3.3.3 Details of 937M-SPIN adapter #1 ...... A3.3-3 Figure A3.3.4 Details of 937M-SPIN adapter #2 ...... A3.3-4 Figure A3.3.5 Stay-out zone around the 937M-SPIN adapter...... A3.3-5
0-16 Figure A3.3.6 Limit load of the 937M-SPIN adapter...... A3.3-6 A Figure A3.3.7 Spacecraft separation shock spectrum with the 937M-SPIN adapter ...... A3.3-7 Figure A3.4.1 Photograph of the 937M-SPIN-A adapter ...... A3.4-1 C Figure A3.4.2 General view of 937M-SPIN-A adapter...... A3.4-2 Figure A3.4.3 Details of 937M-SPIN-A adapter...... A3.4-3 Figure A3.4.4 Stay-out zone around the 937M-SPIN-A adapter ...... A3.4-4 Figure A3.4.5 Limit load of the 937M-SPIN-A adapter ...... A3.4-5 Figure A3.4.6 Spacecraft separation shock spectrum with the 937M-SPIN-A adapter ...... A3.4-6 Figure A3.4.7 Limit loads at separation plane of the 937M-SPIN-A adapter ... A3.4-7 Figure A3.5.1 Photograph of the 1194M adapter ...... A3.5-1 A Figure A3.5.2 General view of 1194M adapter ...... A3.5-2 Figure A3.5.3 Details of 1194M adapter #1 (Details of separation plane) ... A3.5-3 Figure A3.5.4 Details of 1194M adapter #2 (Cross section of frames)...... A3.5-4 Figure A3.5.5 Details of 1194M adapter #3 (Details of frames) ...... A3.5-5 Figure A3.5.6 Stay-out zone around the 1194M adapter (I/III) ...... A3.5-6 Figure A3.5.7 Stay-out zone around the 1194M adapter (II/IV)...... A3.5-7 Figure A3.5.8 Limit load of the 1194M adapter ...... A3.5-8 Figure A3.5.9 Spacecraft separation shock spectrum of the 1194M adapter ...... A3.5-9 Figure A3.5.10 Limit loads at separation plane of the 1194M adapter ...... A3.5-10 B Figure A3.6.1 General view of the 1666M adapter...... A3.6-2 A Figure A3.6.2 Details of the 1666M adapter #1...... A3.6-3 Figure A3.6.3 Details of the 1666M adapter #2...... A3.6-4 Figure A3.6.4 Details of the 1666M adapter #3 (Spacecraft frame) ...... A3.6-5 Figure A3.6.5 Stay-out zone around the 1666M adapter ...... A3.6-6 Figure A3.6.6 Limit load of the 1666M adapter ...... A3.6-7 Figure A3.6.7 Spacecraft separation shock spectrum of the 1666M adapter ...... A3.6-8 Figure A3.6.8 Limit loads at separation plane of the 1666M adapter ...... A3.6-9 B Figure A3.7.1 General view of the 1666S adapter ...... A3.7-2 Figure A3.7.2 Details of the 1666S adapter #1 ...... A3.7-3 Figure A3.7.3 Details of the 1666S adapter #2 ...... A3.7-4 Figure A3.7.4 Stay-out zone around the 1666S adapter #1 ...... A3.7-5 B Figure A3.7.5 Stay-out zone around the 1666S adapter #2 ...... A3.7-6 Figure A3.7.6 Limit load of the 1666S adapter ...... A3.7-7 A Figure A3.7.7 Spacecraft separation shock spectrum of the 1666S adapter ...... A3.7-8 Figure A3.7.8 Limit loads at separation plane of the 1666S adapter ...... A3.7-9 B
0-17 Figure A3.8.1 Photograph of the 2360S adapter...... A3.8-1 A Figure A3.8.2 General view of 2360S adapter ...... A3.8-2 Figure A3.8.3 Details of 2360S adapter #1 ...... A3.8-3 Figure A3.8.4 Details of 2360S adapter #2 ...... A3.8-4 Figure A3.8.5 Details of 2360S adapter #3 ...... A3.8-5 Figure A3.9.1 General view of 3470S adapter ...... A3.9-2 C Figure A3.9.2 Details of 3470S adapter #1 ...... A3.9-3 Figure A3.9.3 Details of 3470S adapter #2 ...... A3.9-4 Figure A3.9.4 Details of 3470S adapter #3 ...... A3.9-5 Figure A3.10.1 Photograph of the 1666MA adapter ...... A3.10-1 Figure A3.10.2 General view of the 1666MA adapter ...... A3.10-2 Figure A3.10.3 Details of the 1666MA adapter #1 ...... A3.10-3 Figure A3.10.4 Details of the 1666MA adapter #2 ...... A3.10-4 Figure A3.10.5 Details of the 1666MA adapter #3 ...... A3.10-5 Figure A3.10.6 Stay-out zone around the 1666MA adapter (+X / -X) ...... A3.10-6 Figure A3.10.7 Stay-out zone around the 1666MA adapter (+Y / -Y) ...... A3.10-7 Figure A3.10.8 Limit loads of the 1666MA adapter ...... A3.10-8 Figure A3.10.9 Spacecraft separation shock spectrum of the 1666MA adapter ...... A3.10-9 Figure A3.10.10 Limit loads at separation plane of the 1666MA adapter ..... A3.10-10 Figure A3.11.1 General view of the 239M adapter ...... A3.11-2 Figure A3.11.2 Details of the 239M adapter ...... A3.11-3 Figure A3.11.3 Spacecraft envelope of the piggyback satellite for the 239M adapter ...... A3.11-4
0-18 TABLES
Table 1.2.1 Summary of H-IIA subsystems and characteristics ...... 1-7 AC
Table 2.1.1 Summary of H-IIA launch capabilities ...... 2-5 A Table 2.1.2 Typical sequence of events of H-IIA vehicle family for GTO mission ...... 2-6 A
Table 3.2.1 Combined loads (Limit level) ...... 3-5 Table 3.2.2 Summary of pyrotechnic shock events ...... 3-8 Table 3.3.1 Gas conditioning capabilities ...... 3-13 Table 3.3.2 Maximum temperature and emittance of fairing internal surface ...... 3-14 Table 3.5.1 Source of gaseous contaminant to spacecraft and its installation location ...... 3-25 C Table 3.7.1 Spacecraft structural tests, margin, and duration ...... 3-28 Table 3.7.2 Spacecraft qualification and acceptance tests requirement ...... 3-28
Table 4.4.1 Characteristics of payload fairings ...... 4-8 Table 4.5.1 Characteristics of payload adapters ...... 4-51 C Table 4.6.1 Electrical interfaces ...... 4-63 Table 4.6.2 Standard interface connector specification ...... 4-70
Table 6.2.1 (1/6) Mission Integration WBS ...... 6-9 A Table 6.2.1 (2/6) Mission Integration WBS ...... 6-10 Table 6.2.1 (3/6) Mission Integration WBS ...... 6-11 Table 6.2.1 (4/6) Mission Integration WBS ...... 6-12 Table 6.2.1 (5/6) Mission Integration WBS ...... 6-13 Table 6.2.1 (6/6) Mission Integration WBS ...... 6-14 Table 6.3.1 Standard spacecraft / H-IIA interface control specifications ...... 6-16
Table A1.1.1 Launch results summary of NASDA launch vehicles (from 1975 to 2000) ...... A1-2 BAC
Table A2.3.1 Summary of power supply (STA2) (1 / 2) ...... A2-15 Table A2.3.1 Summary of power supply (STA2) (2 / 2) ...... A2-16 Table A2.4.1 Summary of power supply (SFA) ...... A2-20 Table A2.5.1 Summary of power supply (TSA) ...... A2-20-3 A Table A2.6.1 Summary of power supply (VAB 9th and 10th floor) ...... A2-24 Table A2.7.1 Summary of power supply (ML) ...... A2-27
0-19 ABBREVIATIONS AND DEFINITIONS
AB Administration Building ADEOS Advanced Earth Observation Satellite AGE Aerospace Ground Equipment AH Ampere - Hour BET Best Estimate Trajectory B / H Blockhouse CAM Collision Avoidance Maneuver A CCW Counterclockwise CDR Command Destruct Receiver CG Center of Gravity CLA Coupled Loads Analysis COMETS Communications and Broadcasting Engineering Test Satellite DLF Design Load Factor DOP Dilution of Precision EMC Electromagnetic Compatibility EMI Electromagnetic Interference ESA European Space Agency ETS Engineering Test Satellite FM Flight Model FRR Flight Readiness Reviewer FSO Flight Safety Officer GCC Guidance Control Computer GEO Geostationary Earth Orbit GFRP Glass Fiber Reinforced Plastic GHe Gaseous Helium GMS Geostationary Meteorological Satellite
GN2 Gaseous Nitrogen GOX Gaseous Oxygen GPSR Global Positioning System Receiver C GSE Ground Support Equipment GTO Geostationary Transfer Orbit h Altitude ha Apogee Altitude hp Perigee Altitude HTV H-II Transfer Vehicle C
0-20 ICS Interface Control Specification i Inclination IDF Intermediate Distributing Frame A in. inch INMARSAT International Maritime Satellite Organization ISAS Institute of Space and Astronautical Science LB Launch Building LCDR Launch Conductor LE Liquid Rocket Engine LEO Low Earth Orbit
LH2 Liquid Hydrogen LOX Liquid Oxygen LP Launch Pad LPLF Lower Payload Fairing LRB Liquid Rocket Booster LSOM Launch Site Operations Manager LVRR Launch Vehicle Readiness Reviewer MECO Main Engine Cutoff MECOM Main Engine Cutoff Command ML Movable Launcher A MOD Mission Operations Department A MRR Mission Readiness Reviewer MTCS Masuda Tracking and Communication Station N / A Not Applicable NASA National Aeronautics and Space Administration NASDA National Space Development Agency of Japan NDTF Nondestructive Test Facility NMD NASDA Mission Director NQA NASDA Quality Assurance Monitor NSAFE NASDA Pad Safety Officer NSO NASDA Safety Officer NTO Nitrogen Tetra Oxide OA Overall OIS Operational Intercommunication Telephone System OLR Osaki Launch Range
0-21 OSTS Office of Space Transportation Systems
PC2 Second Stage Propellant Consumption PCD Pitch Circle Diameter PFM Protoflight Model PIF Poly Iso-cyanurate Form pl Place PLA Payload Adapter PLF Payload Fairing PSS Payload Support Structure PST Pad Service Tower A QD Quick Disconnector R Radius RCC Range Control Center RCO Range Control Officer RCS Reaction Control (gas jet) System REF Reference RF Radio Frequency RL Rocket Launcher Q Dynamic Pressure SBB Solid Booster Test Building SC Spacecraft Conductor S / C Spacecraft SCC Satellite Control Center SCRR Spacecraft Readiness Reviewer SDB Sequence Distribution Box SECO Second Engine Cutoff SECOM Second Engine Cutoff Command SECT Section SEIG Second Engine Ignition SEP Separation SFA Spacecraft and Fairing Assembly Building SFU Space Free Flyer Unit SLB Supporting Launch Building Sm3 Standard Cubic Meter
0-22 SOB Strap-on Booster SPL Sound Pressure Level SRB Solid Rocket Booster SSB Solid Strap-on Booster SSO Sun-synchronous Orbit STA Spacecraft Test and Assembly Building STA Station STM Structural Test Model T.B.C. To Be Confirmed T.B.D. To Be Determined T.B.R. To Be Revised
TNSC Tanegashima Space Center A TRMM Tropical Rainfall Measuring Mission TSA Third Stage and Spacecraft Assembly Building TT/C Telemetry, Tracking, and Command A TVC Thrust Vector Control UHF Ultra High Frequency UM Umbilical Mast UPLF Upper Payload Fairing UPS Uninterruptible Power System VAB Vehicle Assembly Building VDC Voltage Direct Current VHS Video Home System VOS Vehicle On Stand α Angle of Attack φ Diameter ω Argument of Perigee Ω Ascending Node
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C H A P T E R 1 .
IN T R O D U C T IO N CHAPTER 1. INTRODUCTION
1.1 Purpose of the User’s Manual This manual provides users with information on the H-IIA launch system, spacecraft / launch vehicle interfaces and the related NASDA launch services. It presents mission performance, launch and flight environments and interface requirements to be considered in the spacecraft design. Launch operations of the spacecraft, interface management and documentation, and outlines of the H- IIA launch vehicle and Tanegashima Space Center (TNSC) are also described. There are two more documents related to H-IIA launch services on launch facilities and range safety. A brief description of these documents is provided in § 1.3.
1.2 H-IIA Launch System The H-IIA launch system, which includes the launch vehicle, launch facilities, etc., is developed by NASDA as an upgraded version of the H-II launch system to answer the various mission demands for many types of payloads, such as the spacecraft of 2-ton class to 3-ton class Geostationary Earth Orbit* (GEO) in the beginning of the 21st century. The H-IIA development policies are shown below: a) To support various demands for many kinds of payloads b) To keep the high reliability and heritage using the technology, experience, personnel and facilities demonstrated in the H-II program c) To improve operability d) To optimize launch opportunities e) To reduce launch costs This section provides a brief description of the H-IIA launch vehicle and launch facilities.
1.2.1 H-IIA launch vehicle (1) General Figure 1.2.1 shows H-IIA launch vehicle configurations, and Table 1.2.1 shows a summary of H-IIA subsystems and characteristics. Discrimination names are used to distinguish a H-IIA launch vehicle configuration as shown in the following: discrimination name format: H2Aabcd, where a : type of stage (single-stage type = 1, two-stage type = 2) C
1-1 b : number of Liquid Rocket Boosters (LRB) c : number of Solid Rocket Booster-As (SRB-A) d : number of Solid Strap-on Boosters (SSB) (if SSB is not used, this number is omitted) The H-IIA launch vehicle is designed with two stages, each powered by engines using liquid hydrogen and liquid oxygen propellant. There are two models, H2A202 Series and H2A212 (or whether LRB is applied or not). Furthermore, in the H2A202 Series, there are three variations with different numbers of SSB. The H2A202 Series H-IIA launch vehicle, which is named H2A202d (d is the number of SSB), consists of the first stage, two SRB-As, the second stage and the payload fairing as shown in Figure 1.2.2. In addition, two or four SSBs are strapped to the side of the first stage according to a user’s requirement. The H2A212 consists of the first stage, two SRB-As, one LRB, the second stage and the payload fairing as shown in Figure 1.2.3. The H2A212 is not equipped with SSBs. The first stage, the SRB-A, the second stage and the payload fairing are commonly used for H2A202 Series and H2A212 class configurations.
(2) First stage The first stage is composed of LOX / LH2 propellant tanks, the engine section, the LE-7A engine, the center body section, the interstage section and the avionics system. In addition, two or four SSBs are attached according to mission requirements. Both propellant tanks are enlarged to increase capacity, using the same aluminum alloy isogrid structures, insulation (PIF) and cryogenic technology as those of the H-II first stage, although spun-formed tank domes are used for both tanks (instead of orange-peel domes) and an attachment ring is added to the LH2 tank cylinder to attach SRB-As and SSBs. The LE-7A engine is a modified model of the LE-7 engine, with improvement in reliability and operability while retaining most of the characteristics and performance (excepting ISP) of the LE-7 engine. The engine section structure, which is an aluminum alloy monocoque, is modified based on the structure of H-II vehicle so that the first stage can stand directly on the launch pad and support two SRB-As, LRBs, and also SSBs. The center body section is an aluminum alloy semi-monocoque structure, and connects the LOX tank and the LH2 tank as in the structure of the H-II vehicle; the structure which straps the LRB can be added to this section. A The interstage section is newly made of CFRP. The aft end of this section is connected to the front end of the first stage LOX tank. On the other hand, its front end is bolted to the aft end of the second stage LH2 tank, where the first and
1-2 second stage separation mechanism which uses a linear shaped charge is installed. The avionics system whose main components are installed in the center body section is newly developed. This system uses the serial data bus system (MIL- A STD-1553B) and is connected electrically to each avionics system of the second stage and the LRB by this bus system. The command signals from the guidance control computer of the second stage (GCC2) are transmitted to the guidance control computer of the first stage (GCC1). The GCC1 sends command signals to each component and controls each nozzle angle of the LE-7A engine and the SRB-A. Two or four SSBs are attached to the engine section and the attachment ring of the first stage LH2 tank. In the case of the H2A2022, two SSBs are ignited at about 50 seconds after lift-off. In the case of the H2A2024, four SSBs are ignited separately in pairs after lift-off. One pair is ignited at about 10 seconds after lift- off and another pair is ignited at about 76 seconds. A
(3) LRB (Liquid Rocket Booster) The LRB has design features similar to those of the first stage except the number of LE-7A engines and application of a nose cone instead of the interstage section. The LRB has two LE-7A engines which are clustered and ignited prior to lift-off. Engine nozzles are controlled by the guidance control computer of the LRB (GCCL). Each engine is cut off separately with time lag to minimize a bias angle of the each engine nozzle.
(4) SRB-A (Solid Rocket Booster) The SRB for the H-IIA, named SRB-A, is redesigned based on the solid rocket technology of the H-II SRB to enhance its reliability and operability. Two SRB-As are strapped on the core stage and augment LE-7A engine thrust from lift-off for about 100 seconds. The motor case of the SRB-A is monolithic and made of filament winding composite material (CFRP) using Thiokol Castor 120 technology. To enhance the launch capability, the propellant capacity is increased by about 6 tons A optimizing the thrust pattern. Further, to enhance reliability, electromechanical thrust vector control (TVC) system for gimbaling the nozzle is applied.
(5) Second stage The second stage is composed of LOX / LH2 propellant tanks, the LE-5B engine, the reaction control (gas jet) system (RCS), and the avionics system. In the H-II, it has the integrated tank with common bulkhead between the LOX C and LH2 tanks. In the H-IIA vehicle, separated LOX and LH2 tanks are adopted to improve operability. The LH2 propellant tank uses the same aluminum alloy isogrid structure,
1-3 insulation (PIF) and cryogenic technology as the H-II, although spun-formed C tank domes are used instead of orange peel domes. The LOX propellant tank is made by welding two elliptical spun-formed tank domes of the same aluminum C alloy as the LOX tank of the H-II vehicle. The LE-5B engine is a modified model of the LE-5A engine with improved reliability and operability, enhancing the thrust from 12.4 tonf to 14 tonf. This engine is able to operate at two thrust levels of 100 % and 60 % for the rated B thrust, and in the idle mode where the thrust level is about 3 % for the rated A thrust. Further, this engine has a multiple restart capability and a long duration coast capability. The RCS is used for attitude control and propellant settling of the second stage C before and after the spacecraft separation. The RCS is mounted under the A component equipment panel of the guidance section and uses hydrazine as its C propellant. Most components of the avionics system are mounted on the component equipment panel. Most of data processing for the equipment mounted on the C second stage and output of the control signal to the equipment are performed by the GCC2. The programs mounted on the GCC1 and GCCL are controlled by the program loaded in the GCC2.
(6) Payload fairings The payload fairing protects the spacecraft from environments to which the spacecraft is exposed from the time of encapsulation through the atmospheric ascent phase. Users are able to choose from 6 types described in § 1.2.3 according to user’s C requirements. Each type is compatible with every H-IIA launch vehicle. Description on payload fairings is provided further in Chap. 4.
(7) Payload adapters / separation system The spacecraft is mounted on the launch vehicle using a payload adapter described in § 1.2.3. Detailed information on the adapters and their separation systems is presented in Chap. 4. Each H-IIA launch vehicle is compatible with each type of payload adapters (including separation systems).
1.2.2 Launch facilities Launch operations are carried out in Osaki Launch Range (OLR) of TNSC which is located at the southeastern end of Tanegashima island. TNSC is the largest rocket range in Japan with the total site area of 8.6 million square meters and accommodates various necessary facilities to launch spacecraft. Necessary facilities for spacecraft launch operations are mostly located in
1-4 OLR. Figure 1.2.4 shows locations of major facilities in OLR. Major facilities related to spacecraft operations are shown in Figures 1.2.5 to 1.2.8. Launch processing in these facilities is normally performed as shown in Figure 1.2.9. If a user requests different processing, the user should coordinate with NASDA.
1.2.3 Payload accommodations NASDA offers 6 types of payload fairings and several types of payload adapters. C Figure 1.2.10 shows payload usable envelops for the model 4S and 5S payload fairings used for a single launch. Figure 1.2.11 shows payload usable envelops for the model 4/4D-LS, 4/4D-LC and 5/4D payload fairing used for dual launch. Figure 1.2.12 shows payload usable envelop for the model 5S-H payload fairing C used for a single launch. Each fairing is compatible with each of the H-IIA launch vehicle configurations. More information of the payload fairing is presented in § 4.4. A Figure 1.2.13 shows example of payload adapters. Detailed information related C to interface with the spacecraft is presented in § 4.5 and APPENDIX 3.
1.2.4 Users / NASDA relationship NASDA offers a full launch service from interface coordination at the A beginning of mission planning through a separation of the spacecraft on the required orbit. For executing this service, NASDA provides a single point of contact, the program manager, for each spacecraft. The program manager shall be responsible for contracts and coordination related to launch. Figure 1.2.14 C shows a concept of users / NASDA relationship and major responsibilities of the program manager. After establishment of the NASDA launch operations team, although the mission director is designated and responsible only for the technical coordination with the user, the program manager commands this coordination. Figure 1.2.15 shows this relationship. C
1.2.5 Advantages of H-IIA The H-IIA launch system and launch services offer the following advantages. a) The H-IIA launch vehicle family provides flexible and various launch capabilities for payloads. b) The H-IIA launch vehicle and launch operations are very simple and reliable. c) Environmental conditions to which the spacecraft is exposed are comparable with or better than those of other launch vehicles. d) The H-IIA launch system is capable of dual launch if a user requires.
1-5 A
H2A202 H2A2022 H2A2024 H2A212 H2A222 (GTO 4.1/3.7 ton) (GTO 4.5/4.1 ton) (GTO 5.0/4.6 ton) (GTO 7.5 ton) (GTO T.B.D. ton) A (*1) (*2) (*1) (*2) (*1) (*2) (*1) C
Current Program Future Program
Figure 1.2.1 H-IIA Launch Vehicle Family (*1):With the LE-7A lower nozzle skirt C (*2):Without the LE-7A lowor nozzle skirt
1-6 Table 1.2.1 Summary of H-IIA subsystems and characteristics A
Item H2A202 H2A2022 H2A2024 H2A212 H2A222 Note Overall length (m) 53 53 53 53 53 C Diameter (m) : Core 4.0 4.0 4.0 4.0 4.0 � : Fairing 4.07 / 5.1 4.07 / 5.1 4.07 / 5.1 4.07 / 5.1 4.07 / 5.1 4S, 4/4D type / 5S, 5/4D type Total weight (ton) : Without Payload 285 316 347 403 520 include payload adapter C Payload weight (ton) : With the LE-7A lower nozzle skirt 4.1 4.5 5.0 7.5 T.B.D. (100 kg) GTO for 4S fairing : Without the LE-7A lower nozzle skirt 3.7 4.1 4.6 (5S fairing for H2A212) Payload fairing : Honeycomb sandwich A 4S,5S type : Clamshell 4/4D-LS, 5/4D type : Upper Clamshell, Lower Tandem 4/4D-LC type : Upper Clamshell, Lower Clamshell 5S-H type : Clamshell Payload adapter C 937M-Spin, 937M-Spin-A : Aluminum, V band, Spin table A 937M, 937MH : CFRP, V band 1666M, 1194M, 1666MA : Aluminum, V band 2360S, 3470S, 1666S : Separation Nuts C First stage 1 1 1 1 1 Propellant :LH2 / LOX 100 ton 100 ton 100 ton 100 ton 100 ton usable weight Tank :LH2 aluminum isogrid : LOX aluminum isogrid Propulsion : LE-7A engine x 1 Thrust 1100 kN in vacuum B : With the LE-7A lower nozzle skirt 1073 kN in vacuum C : Without the LE-7A lower nozzle skirt Isp 440 sec in vacuum : With the LE-7A lower nozzle skirt � 429 sec in vacuum : Without the LE-7A lower nozzle skirt MR 5.9 Burning time 390 sec for roll control : Auxiliary engine time 390 sec
Avionics use data bus system : Guidance Control Computer, Flight (MIL-STD-1553B) Termination, Rate Gyro Package, Lateral Acceleration Unit, VHF LRB Telemetry, Electrical Power N/AN/A N/A 1 2 B Propellant :LH2 / LOX 99.2 ton 99.2 ton usable weight Tank : Same as first stage Propulsion : LE-7A engine x 2 cluster A Burning time 195 sec (no throttling) Avionics : Guidance Control Computer, use data bus system Flight Termination, Electrical Power (MIL-STD-1553B) A SRB-A 22222 Propellant : HTPB composite B (per each) Propellant weight 65.04 ton Thrust (max) 2260 kN in vacuum A Isp 280 sec in vacuum Burning time 100 sec Motor case : Monolithic CFRP Diameter 2.5 m SSB 0 2 4 N / A N / A two SSB are one pair Propellant : HTPB composite (per each) Propellant weight 13.1 ton Thrust (Max) 710 kN in vacuum B Isp (mean) 283 sec in vacuum Burning time 60 sec Motor case : Steel Diameter 1.02 m Second stage 1 1 1 1 1 Propellant :LH2 / LOX 16.6 ton 16.6 ton 16.6 ton 16.6 ton 16.6 ton usable weight Tank :LH2 aluminum isogrid : LOX aluminum Propulsion : LE-5B engine x 1 B Thrust 137 kN A Isp 447 sec MR 5.0 B Throttling 60 % for rated thrust A Idle mode function about 3 % for rated thrust Multi-restart function : Reaction control system (hydrazine) for attitude control : Guidance Control Computer, Avionics Guidance, Navigation, Control Inertial Measurement Unit, Flight and Vehicle Sequencing Termination, UHF Telemetry, use data bus system C-Band Tracking, Range Safety (MIL-STD-1553B) Command, Electrical Power
1-7 1. Payload fairing 1 2. Spacecraft 3. Payload adapter 2 4. Payload support structure 5. Cryogenic He bottles
6. Second stage LH2 tank 3 4 7. Second stage LOX tank
5 8. Avionics Equipment panel A 9. Reaction control system 6 10. Ambient He bottles 11. Second stage engine (LE-5B engine) 7 8 12. Interstage section 9 10 13. First stage LOX tank 12 14. Center body section �11 � 15. First stage LH2 tank 16. Solid rocket booster (SRB-A)
13 17. First stage engine section 18. Auxiliary engine 19. Ambient He bottles 14 20. First stage engine (LE-7A engine) 21. SRB-A movable nozzle A
15
16
17 19 18
21 20
Figure 1.2.2 H-IIA (H2A202) launch vehicle configuration A
1-8 1. Payload fairing 1 2. Spacecraft 3. Payload adapter 2 4. Payload support structure 5. Cryogenic He bottles 3 4 6. Second stage LH2 tank 7. Second stage LOX tank 8. Avionics Equipment panel A 5 9. Reaction control system 6 10. Ambient He bottles 11. Second stage engine (LE-5B engine) 7 A 22 12. Interstage section 8 9 13. First stage / LRB LOX tank 11 10 14. First stage / LRB center body section 12 15. First stage / LRB LH2 tank A 16. Solid rocket booster (SRB-A) 13 13 17. First stage / LRB engine section 18. Auxiliary engine 19. First stage / LRB engine (LE-7A engine) 20. SRB-A movable nozzle 14 14 21. Liquid rocket booster (LRB) 22. LRB nose cone
15 15
21
16
17 17 18
19
20 19
Figure 1.2.3 H-IIA (H2A212) launch vehicle configuration
1-9 A Non-Destructive Test Facility
N No.2 Spacecraft Test to Hirota and Assembly Building
Spacecraft and Fairing Spin Test Building Assembly Building Osaki Power Station Spin Liquid Propellant Measument storage Building Liquid Oxydizer Osaki 5 roads storage crossing gate Third Stage and Satellite Solid Booster Building Yoshinobu Vehicle Assembly Building Assembly Building
No.1 Spacecraft Test Osaki No.1 Support Yoshinobu Block House and Assembly Building LSB gate Garage Osaki Supporting Liquid Hydrogen Storage Reserroir Launch Building Propellant New Yoshinobu Block House (underground) Osaki Restaurant Storage Facility Yoshinobu Second Camera Room Osaki office 2 Osaki office 1 Yoshinobu (H-IIA) Launch Complex LE-7 Engine Launch 0 100 200 300 400 500 Building Tool Room Test Stand Pad Service Tower (Launch Pad)
New Launch Pad 80m Meteorological Observation Tower Liquid Oxygen Storage
High-pressure Sea shore gate Gas Storage
to Kukinaga, Kaminaka
Pacific Ocean
to Takesaki Administration Building
Figure 1.2.4 Location of major facilities in Osaki Launch Range
1-10 Figure 1.2.5 No. 2 Spacecraft test and assembly building (STA2)
Figure 1.2.6 Spacecraft and fairing assembly building (SFA)
1-11 B
Figure 1.2.7 Vehicle assembly building (VAB) A
1-12 Figure 1.2.8 Overview of New Yoshinobu Launch Complex C
1-13 STA 2 Lift-off
No.2 Spacecraft test and assembly Launch pad building (LP) (STA 2)
¥ Spacecraft integration ¥ Terminal Countdown B ¥ Spacecraft functional test
Spacecraft and fairing Vehicle assembly assembly building building ML (SFA) (VAB)
¥ Propellant loading
SFA VAB
Transporting A from SFA to VAB
Figure 1.2.9 H-IIA launch operations process
1-14 4430
φ3700 12000 * 5800
* A * φ1194* Spacecraft 110 480 separation plane PLA / PSS interface PLF
90 φ2190 (REF) separation 1021 PSS / second plane stage interface φ4070 Model 4S 2601 φ3378 A 2063
φ4600 12000 * φ5100 4460.5
* A φ1194* Spacecraft 110 separation plane
* PLA / PSS interface 480 PLF φ2190 (REF) 90 separation 1965.5 PSS / second plane stage interface φ4070 Unit : mm Model 5S
Figure 1.2.10 Payload fairings for single launch with 1194M adapter A
1-15 4430
10000 φ3700 * 3800 * * φ1194* Spacecraft A 110 480 separation plane
UPLF PLA / PSS interface
φ1980 861 separation 1021 plane φ3640
φ3700 1000 φ3208* * A * LPLF * φ1194* Spacecraft
separation 1943.5 110 plane 480 separation plane 3444.5
4444.5 PLA / PSS interface φ2190 (REF) Aft end 1021 of LPLF Unit : mm φ4070
Model 4/4D-LS 4430
10000 φ3700 * 3800 * * φ1194* A Spacecraft 110 480 separation plane UPLF PLA / PSS interface separation φ1980 924 plane 1021
LPLF
separation φ3700 * plane 4436 * * φ1194* 5500 6000 Spacecraft 110 480 separation plane PLA / PSS interface φ2190 (REF) Aft end 1021 of LPLF
1-16 2601 φ3378 A 2063 * 9580.5 φ4600 * * A φ1194*
2041 Spacecraft 480 110 separation plane PLA / PSS interface φ1980
UPLF 1740 separation φ3640 1965.5 plane
φ3700 1000
φ3208* *
LPLF * * φ1194 separation * Spacecraft 1943.5 110 plane 480 separation plane 3444.5
4444.5 PLA / PSS interface φ2190 (REF) 1021
φ4070
Unit : mm
Model 5/4D
Figure 1.2.11(2/2) Payload fairings for dual launch with 1194M adapter A
1-17 C
φ3378
φ4600 15000
5100 φ 7885.6 2063 2601
φ3183.7 (REF)
Spacecraft 780 separation plane
1177.4 PLA PSS interface 843 ／ PLF φ3200 （REF） PSS／second separation plane 90 stage interface
φ4070 Unit:mm Model 5S-H
Figure 1.2.12 Payload fairing for single launch
1-18 A aluminum alloy semi-monocoque V band clamp, spin table aluminum alloy semi-monocoque 958*1 V band clamp, spin table φ B 1 φ959* A 1000 605
φ2190 φ2190 Model 937M-Spin adapter Model 937M-Spin-A adapter aluminum alloy semi-monocoque aluminum alloy semi-monocoque V band clamp V band clamp 1 *1 φ945* aluminum alloy semi-monocoque φ958 BC V band clamp φ1666 A 900 900 480
φ2190 φ2190 φ2190 Model 937M adapter Model 1666MA adapter Model 937MH adapter A aluminum alloy semi-monocoque aluminum alloy semi-monocoque V band clamp separation nuts (4 pl) *1 φ1666 φ1664 (PCD) aluminum alloy semi-monocoque V band clamp 1 φ1194* 1000 1000 480
2190 φ2190 φ2190 φ Model 1194M adapter Model 1666M adapter Model 1666S adapter
aluminum alloy semi-monocoque separation nuts (4 pl) φ3472 (PCD) graphite epoxy separation nut (4 pl)
φ2360 (PCD) 1089 300
φ2190 φ2190 Model 2360S adapter Model 3470S adapter
Figure 1.2.13 Payload adapters (example) C
1-19 Interface Coordination Mission Operations Department A User 1
Range Safety Control Spacecraft Interface Manager 1 Program Manager 1
Mission Planning User 2
Spacecraft Interface Manager 2 Program Manager 2 Launch Operations
Vehicle Manufacturer Control
Figure 1.2.14 Concept of users / NASDA relationship C
1-20 NASDA launch operations team
User 1 NASDA Mission Director 1 NASDA Program Manager 1
User 2 NASDA Mission Director 2 NASDA Program Manager 2
Contract coordination
Technical coordination
Figure 1.2.15 Concept of users / NASDA relationship C after establishment of NASDA launch operations team
1-21 1.3 H-IIA Launch System Related Documents NASDA provides two more documents related to the H-IIA launch system besides this user’s manual. These documents provide users with detailed information on launch facilities and range safety requirements as a reference for the preliminary planning phase. They are: a) H-IIA Payload-Related Facilities and Ground Support Equipment (GSE) Manual b) Launch Vehicle Payload Safety Requirements (NASDA-STD-14B) A
1.3.1 H-IIA Payload-Related Facilities and GSE Manual (in preparation) The H-IIA Payload-Related Facilities and GSE Manual provides information about TNSC and launch preparation operations of spacecraft. Appendix 2 of this user’s manual describes launch facilities and related launch preparation operations briefly. Detailed information is provided in the H- IIA Payload-Related Facilities and GSE Manual which includes a number of pictures and drawings of buildings and equipments.
1.3.2 Launch Vehicle Payload Safety Requirements This document provides the requirements for safety control, spacecraft safety design, and launch site operations at TNSC. A description related to safety reviews is provided in § 6.5.
1.4 Definition of terms The terms used in this manual are generally defined in the body of the text. The following terms are defined below: A
(1) User An organization or an individual who entrusts or wishes to entrust a launch of the spacecraft to NASDA.
(2) Spacecraft organization The user or the spacecraft builder to whom the user has entrusted the fabrication, testing and spacecraft launch operations. In case of a NASDA spacecraft, the spacecraft organization means the NASDA spacecraft organization and the term of NASDA corresponds to the launch vehicle organization.
1-22 (3) NASDA launch operations team A provisional team organized by NASDA to launch the launch vehicle.
(4) Launch vehicle organization A group within the NASDA launch operations team in charge of launch operations for the launch vehicle.
(5) Launch operations This is a generic term referring to the assembly, preparation, testing and joint operations of the spacecraft and launch vehicle organizations which are to be implemented at TNSC.
(6) Program manager A NASDA person responsible for contracts and interface coordination of the launch; a member of the “Mission Operations Department (MOD).” A
(7) Spacecraft interface manager A user person responsible for the launch activity coordination. In case of a NASDA spacecraft launch, the term refers to the manager of the spacecraft organization or his representative.
(8) NASDA mission director A NASDA staff member responsible for the interface and coordination of technical matters related to launch operations (from spacecraft arrival at TNSC through completion of the post lift-off operation) after establishment of the NASDA launch operations team.
(9) Launch vehicle system A system including the launch vehicle, related facilities, and Aerospace Ground Equipment (AGE).
(10) Spacecraft system A system including the spacecraft, its related facilities, and GSE.
(11) Y ± and X ± “Y - (numerical figure)” represents the number of days before the launch date in terms of actual working days. “Y + (numerical figure)” represents the number of days after the launch date in terms of actual working days. “Y - 0” represents the lift-off date. “X - (numerical figure)” represents the time before lift-off (holding (or margin) time not included).
1-23
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I
C H A P T E R 2 . H O N
A
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T F
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MI S S IO N P E R F O R MA N C E 2 A N
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C E CHAPTER 2. MISSION PERFORMANCE 2.1 General This chapter describes the mission performance of the H-IIA launch vehicle. The H-IIA launch vehicle provides a wide variety of mission performances. However, this manual describes only the following missions. a) Geostationary transfer orbit (GTO) mission b) Sun-synchronous orbit (SSO) mission c) Low Earth orbit (LEO) mission d) Earth escape mission Table 2.1.1 illustrates representative performance capabilities of the H-IIA family. The performance data are based on standard mission modifications of payload fairing (refer to § 4.4.3). Actual launch capability will vary with final A configurations.
2.1.1 Mission profile To provide users with H-IIA mission sequences, typical mission profile and sequences are briefly explained using a GTO mission for example in the following paragraphs. Figures 2.1.1 and 2.1.2 show flight profiles for typical GTO missions in case of H2A2024 and H2A212 respectively. Table 2.1.2 shows typical GTO mission sequence of events for each H-IIA vehicle. In case of other missions such as SSO, LEO, etc., sequence time of events and flight trajectories are different from a GTO mission, but most of these sequences are held similarly to a GTO mission. These data are representative and actual sequence and profile will be prepared to meet spacecraft requirements. As shown in Table 2.1.2, although sequence of events for each vehicle configuration are the same, the sequence time is different respectively. So, in the following paragraphs, sequence time for H2A2024 and H2A212 is used representatively according to Figures 2.1.1 and 2.1.2.
2.1.1.1 Booster and first stage phase The main engine LE-7A is ignited at about 4.7 seconds before lift-off (X-0). In case of H2A212, two more LE-7A engines of the LRB are ignited at the same time. After detecting the rise of the combustion pressure, two SRB-As are ignited (at about 0.5 second before X-0), subsequently the H-IIA vehicle rises away from the ML and the guidance control program on the guidance control computer of the second stage (GCC2) senses the lift-off (that is X-0). About 100 seconds after lift- A off (hereafter called X+100), two SRB-As burn out and are separated from the core stage. In case of H2A2024, the first pair of SSBs is ignited at about X+10, burns out at about X+70 and is separated from the core stage at about X+111. On the other A hand, the second pair of SSBs is ignited at about X+76, burns out at about X+136 and is also separated from the core stage at about X+142.
2-1 In case of H2A212, the LRB will experience a propellant depletion at about X+197 and its main engines are cut off at intervals of a several seconds. And then A the LRB is separated from the core stage. The payload fairing (PLF) is jettisoned after a free molecular heat flux becomes less than 1135 W/m2, resulting the A jettison timing of the PLF being different in each mission. At about X+389, the main engine shutdown command (MECOM) is sent from the guidance control computer of the first stage (GCC1). After the main engine has tailed off, the first and second stage separation is executed. The attitude control is conducted by the engine gimbaling of the core stage, the SRB-As and the LRB, and also auxiliary thrusters. The gimbaling of the core main engine (LE-7A) contributes to the pitch / yaw control throughout the booster and the first stage phase. The SRB-A engine nozzle is gimbaled for the pitch / yaw / roll control. The LRB engines gimbaling participates in the pitch / yaw control during all phase of the LRB, but in roll control, these engines contribute only after the SRB-A tail off. Auxiliary thrusters roll the H-IIA vehicle throughout the booster and first stage phase.
2.1.1.2 Second stage phase After the first and the second stage separation, the LE-5B is ignited by the first ignition command from the GCC2 (SEIG1). The LE-5B engine burns about 300 (or 220) seconds. As soon as the second stage (including the payload) is injected to the parking orbit, the engine is cut off by engine cutoff command from the GCC2 (SECO1). During engine burning, the pitch/yaw control is conducted by the gimbaling of the LE-5B and the roll control is conducted by the reaction control system (RCS). After the engine cutoff, the vehicle starts the coasting flight. During the coast phase, preparations for the second stage engine restart take place. These are propellant settling, pressure control of LOX / LH2 tanks, engine chilling down and so on. In this phase, the attitude control of the pitch/yaw/roll is performed only by the RCS. About 710 (or 630) seconds later, the LE-5B is restarted (SEIG2) and when the A second stage reaches the planned transfer orbit, the engine cutoff command (SECOM2) is sent from the GCC2. The second burning duration is approximately 210 (or 280) seconds in a normal GTO mission. A After spacecraft / second stage separation, using separation springs or GHe retro motor the collision avoidance maneuver (CAM, that is the collision and contamination avoidance maneuver) is conducted using the RCS and the residual GH2 in the LH2 tank.
2-2 Second stage Second stage engine restart Injection into GTO Second stage engine cutoff (SEIG2) engine ignition (SECO1) (SEIG1) Second stage First stage X+1421 s Spacecraft engine cutoff separation main engine cutoff Coasting (MECO) (SECO2) X+1664 s X+405 s X+713 s X+389 s X+1629 s
First stage separation X+399 s
Fairing jettison
X+220 s
Second SSB pair burnout / separation
X+136 s / X+142 s
First SSB pair separation
X+111 s
SRB-A burnout / separation
X+100 s / X+110 s
First SSB pair burnout / Second SSB pair ignition
X+70 s / X+76 s
First SSB pair ignition
X+10 s
Lift-off
X+0 s Figure 2.1.1 Typical GTO mission profile for H2A2024 A
2-3 Second stage Second stage engine restart Injection into engine cutoff (SEIG2) GTO (SECO1) Second stage engine ignition Second stage Spacecraft X+1269 s (SEIG1) Coasting engine cutoff separation (SECO2) X+636 s X+1584 s X+1549 s First stage X+405 s main engine cutoff (MECO) X+389 s First stage separation X+399 s
Fairing jettison
X+249 s
LRB burnout / separation
X+197 s / X+207 s
SRB-A burnout / separation
X+100 s / X+110 s
Lift-off
X+0 s
Figure 2.1.2 Typical GTO mission profile for H2A212 A
2-4 Table 2.1.1 Summary of H-IIA Payload capability C
Payload capability ; kg (lb) C Mission (Orbit)
SSO h = 800 km 3,600 (in Summer) A i = 98.6 deg (7,937) 4,400 (except Summer) — — — (9,700) (5S fairing)
Payload capability ; kg (lb) Mission (Orbit)
SSO h = 800 km 3,500 (in Summer) i = 98.6 deg (7,716) 4,200 (except Summer) ——— (9,259) (5S fairing)
2-5 Table 2.1.2 Typical sequence of events of H-IIA vehicle family for GTO mission A
Events (seconds) H2A202 *1 H2A2022 *1 H2A2024 *1 H2A212 *2 Remarks Guidance flight mode on -13.0 -13.0 -13.0 -13.0 LE-7A (pair) ignition -6.0 -6.0 -6.0 -6.0 SRB-A (pair) ignition -0.6 -0.6 -0.6 -0.6 A Liftoff 0.0 0.0 0.0 0.0 First SSB pair ignition N / A 50.0 10.0 N / A SRB-A burn out 100.0 100.0 100.0 100.0 A First SSB pair burn out N / A 109.8 69.8 N / A SRB-A separation 110.0 110.0 110.0 110.0 Second SSB pair ignition N / A N / A 76.0 N / A First SSB pair separation N / A 121.0 111.0 N / A Second SSB pair burn out N / A N / A 135.8 N / A Second SSB pair separation N / A N / A 142.0 N / A LRB main engine cutoff (LMECO) N / A N / A N / A 197.1 LRB separation N / A N / A N / A 207.1 Fairing jettison 263.8 275.0 220.2 249.0 A Main engine cutoff (MECOM) 389.3 389.3 389.3 389.3 First / Second stage separation 399.3 399.3 399.3 399.3 Second stage ignition 1 (SEIG1) 405.3 405.3 405.3 405.3 Second stage cutoff 1 (SECO1) 732.0 722.1 713.3 636.0 Second stage ignition 2 (SEIG2) 1466.2 1445.1 1420.8 1268.8 Second stage cutoff 2 (SECO2) 1651.7 1639.0 1628.6 1549.1 Spacecraft separation 1686.7 1674.0 1663.6 1584.1
2-6 2.2 Performance Ground Rules H-IIA performance ground rules for various missions are described in this section. 2.2.1 Payload mass definition A Performance capabilities referred to throughout this document are presented in terms of payload mass. Payload mass is defined as follows: A Payload mass is the total mass of the spacecraft injected to the target orbit, excluding the payload adapter (PLA) whose mass is assumed 100 kg including the separation system. This PLA is a standard type for the H-IIA vehicle and named 1194M. If a different type of the PLA is used or other hardware is required, mass difference between 1194M PLA and other PLA or hardware should be A considered in performance capabilities. Information concerning the PLA mass appears in the § 4.5. 2.2.2 Launch vehicle configurations Typical H-IIA performance presented in this document is based on using the 4S fairing unless noted otherwise. If user requires greater volume than the 4S C fairing for a spacecraft or dual launch, the H-IIA can also offer other types of a fairing such as the 5S, the 4/4D-LS, the 4/4D-LC, the 5/4D and the 5S-H fairing. But in case of using these fairings, performance will degrade mainly according to the fairing mass. Information related to the payload fairing mass is given in the § 4.4. 2.2.3 Launch vehicle performance confidence levels The H-IIA launch system is designed with 99.7 % performance confidence level to meet the requirements of each user with flexibility. Performance confidence levels can be set based on each mission’s requirements. 2.3 Geostationary Transfer Orbit (GTO) Mission 2.3.1 Payload capability for single launch Payload mass for GTO mission using 4S fairing , with the LE-7A lower nozzle C skirt is about 4,100kg (not including standard payload adapter) based on parameters in § 2.3.3. And payload mass for GTO mission using 4S fairing , without the LE-7A lower nozzle skirt is about 3,700kg (not including standard payload adapter) based on parameters in § 2.3.3. Payload mass using model 5S fairing is about 350 kg less, with the LE-7A lower nozzle skirt. And payload mass using model 5S fairing is about 300 kg less, without the LE- 7A lower nozzle skirt. Figure2.3.9 ~ 2.3.12 show payload capability of H2A202, H2A2022 and H2A2024 / 4S fairing, H2A212 / 5S for GTO mission for configuration with the LE- 7A lower nozzle skirt . Figure2.3.13~2.3.15 show payload capability of H2A202, H2A2022 and H2A2024 / 4S fairing for GTO mission for configuration without the LE-7A lower nozzle skirt.
2-7 2.3.2 Payload capability for dual (GTO and GTO) launch The combined mass of two spacecraft, upper payload adapter and the lower portion fairing equals the value of single spacecraft launch mass.
2.3.3 Typical orbital parameters Typical orbital parameters for GTO mission are as follows : Apogee altitude ha = 36,226 km *1 Perigee altitude hp = 250 km Inclination i = 28.5 ° Argument of perigee ω = 179.0 ° Assuming the Earth’s equator radius is 6,378.142 km and these values are osculating orbit parameters at spacecraft separation.
2.3.4 Injection accuracies Typical injection accuracies for GTO mission are as follows based on parameters in § 2.3.3. Apogee altitude ∆ ha = ± 180 km A Perigee altitude ∆ hp = ± 4km Inclination ∆ i= ± 0.02 ° Argument of perigee ∆ ω = ± 0.40 ° Longitudinal of ascend node ∆ Ω = ± 0.40 ° (These values are 3-sigma level.)
2.3.5 Typical sequence of events Table 2.1.2 shows a typical sequence of events for GTO mission in the case of A H2A202, H2A2022 and H2A2024 with the 4S fairing.
2.3.6 Typical trajectory Figures 2.3.2, 2.3.4, 2.3.6, 2.3.8 show a typical flight trajectory for GTO A mission in the case of H2A202, H2A2022 and H2A2024 with the 4S fairing, H2A212 with the 5S fairing respectively.
2-8 2.3.7 Typical flight parameters Figures 2.3.1, 2.3.3, 2.3.5, 2.3.7 show a typical flight parameters for GTO A mission in the case of H2A202, H2A2022 and H2A2024 with the 4S fairing, H2A212 with the 5S fairing respectively.
2-9 80 16 40.0 800 A 75 15 37.5 750
70 14 35.0 700
65 13 32.5 650 Dynamic pressure 60 12 30.0 600 Altitude
55 11 27.5 550 Acceleration Pa) ) 2 * 3 m/sec) 10 25.0 500 50 * 3 10 * m) 10 * * 45 9 22.5 3 450 10 40 8 20.0 * 400 Relative velocity
35 7 17.5 350 Altitude (
30 6 Acceleration (m/sec 15.0 300 Dynamic pressure 25 Relative velocity ( 5 12.5 250
20 4 10.0 200
15 3 7.5 150
10 2 5.0 100
5 1 2.5 50
0 0 0.0 0 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 Time from lift-off (sec)
Figure 2.3.1 Typical flight parameters for GTO mission (H2A202 with 4S fairing) A
1 CO
SE 5000 4000 6000
SE 3000 IG 2 1 7000 IG
SE 2000
8000 S E O C
C O E 1000 2 M S Range (km) 9000 /C S E P
10000
400 300 200 40.0 100 0
30.0 Altitude (km) Tanegashima Space Center 400 20.0 600 800 Ogasawara Down Range Station 1000 10.0 Time from lift-off (sec)
1200
1400 Latitude (deg) 0
Christmas 1600 Down Range -10.0 Station 1800
-20.0 120.0 130.0 140.0 150.0 160.0 170.0 180.0 190.0 200.0 210.0 220.0 230.0 Longitude (deg)
Figure 2.3.2 Typical flight trajectory for GTO mission (H2A202 with 4S fairing) A
2-10 16 40.0 800 80 A 75 15 37.5 750
70 14 35.0 700
65 13 32.5 650 Dynamic pressure 60 12 30.0 600 Altitude
55 11 27.5 550 Acceleration Pa) ) * 2 3 m/sec) 10 25.0 500 50 * 3 10 * m) 10 * * 45 9 22.5 3 450 10 * 40 8 20.0 400 Relative velocity
35 7 17.5 350 Altitude (
30 6 Acceleration (m/sec 15.0 300 Dynamic pressure ( 25 Relative velocity ( 5 12.5 250
20 4 10.0 200
15 3 7.5 150
10 2 5.0 100
5 1 2.5 50
0 0 0.0 0 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 Time from lift-off (sec)
Figure 2.3.3 Typical flight parameters for GTO mission (H2A2022 with 4S fairing) A
1 CO
SE 5000 A 4000 6000
SE 3000 IG 2 1 7000 IG
SE 2000
8000 S E O C
C O E 1000 2 M S Range (km) 9000 /C S E P
10000
400 300 200 40.0 100 0
30.0 Altitude (km) Tanegashima
Space Center 400 20.0 600 800 Ogasawara
Down Range Station 1000 10.0 Time from lift-off (sec)
1200
Latitude (deg) 0 1400
Christmas 1600 Down Range -10.0 Station
1800 -20.0 120.0 130.0 140.0 150.0 160.0 170.0 180.0 190.0 200.0 210.0 220.0 230.0 Longitude (deg)
Figure 2.3.4 Typical flight trajectory for GTO mission (H2A2022 with 4S fairing) A
2-11 80 16 40.0 800 A 75 15 37.5 750
70 14 35.0 700
65 13 32.5 650 Dynamic pressure 60 12 30.0 600 Altitude
55 11 27.5 550 Acceleration Pa) ) * 2 3 m/sec) 10 25.0 500 50 * 3 10 * m) 10 * * 45 9 22.5 3 450 10 * 40 8 20.0 400 Relative velocity
35 7 17.5 350 Altitude (
30 6 Acceleration (m/sec 15.0 300 Dynamic pressure ( 25 Relative velocity ( 5 12.5 250
20 4 10.0 200
15 3 7.5 150
10 2 5.0 100
5 1 2.5 50
0 0 0.0 0 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 Time from lift-off (sec)
Figure 2.3.5 Typical flight parameters for GTO mission (H2A2024 with 4S fairing) A
1 CO
SE 5000 4000 6000
SE 3000 IG 2 1 7000 IG
SE 2000
8000 S E O C
C O E 1000 2 M S Range (km) 9000 /C S E P
10000
400 300 200 40.0 100 0
30.0 Altitude (km) Tanegashima Space Center 400 20.0 600 800 Ogasawara
Down Range Station 1000 10.0 Time from lift-off (sec)
1200
Latitude (deg) 0 1400 Christmas Down Range 1600 -10.0 Station
1800 -20.0 120.0 130.0 140.0 150.0 160.0 170.0 180.0 190.0 200.0 210.0 220.0 230.0 Longitude (deg)
Figure 2.3.6 Typical flight trajectory for GTO mission (H2A2024 with 4S fairing) A
2-12 80 16 40.0 800 A 75 15 37.5 750
70 14 35.0 700 Altitude 65 13 32.5 650 Dynamic pressure 60 12 30.0 600
11 27.5 550 55 Acceleration Pa) ) * 2 3 m/sec) 10 25.0 500 50 * 3 10 * m) 10 * * 45 9 22.5 3 450 10 * 40 8 20.0 400 Relative velocity
35 7 17.5 350 Altitude (
30 6 Acceleration (m/sec 15.0 300 Dynamic pressure ( 25 Relative velocity ( 5 12.5 250
20 4 10.0 200
15 3 7.5 150
10 2 5.0 100
5 1 2.5 50
0 0 0.0 0 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 Time from lift-off (sec)
Figure 2.3.7 Typical flight parameters for GTO mission (H2A212 with 5S fairing) A
1 CO
SE 5000 4000 6000
SE 3000 IG 2 1 7000 IG
SE 2000
8000 S E O C
C O E 1000 2 M S Range (km) 9000 /C S E P
10000
400 300 200 40.0 100 0
30.0 Altitude (km) Tanegashima
Space Center 400 20.0 600 Ogasawara 800 Down Range Station Time from lift-off (sec) 10.0 1000
1200
Latitude (deg) 0 1400 Christmas Down Range -10.0 Station 1600
1800 -20.0 120.0 130.0 140.0 150.0 160.0 170.0 180.0 190.0 200.0 210.0 220.0 230.0 Longitude (deg)
Figure 2.3.8 Flight trajectory for GTO mission (H2A212 with 5S fairing) A
2-13 Payload mass (kg) Payload mass (kg)
6000 6000
Inclination : 5500 28.5 deg 5000 24 deg 5000 20 deg 4500 16 deg 4500 4000 4000
3500 3500
3000 3000
2500 2500
2000 2000 Inclination : 1500 1500 28.5 deg 24 deg 1000 1000 20 deg 16 deg 500 500
0 0 20000 30000 40000 50000 60000 70000 80000 20000 30000 40000 50000 60000 70000 80000 Apogee Altitude (km) Apogee Altitude (km)
Figure 2.3.9 Payload capability for Figure 2.3.10 Payload capability for GTO GTO mission (H2A202 with 4S fairing) mission (H2A2022 with 4S fairing)
Payload mass (kg) Payload mass (kg)
8000 9000
7500 8500 Inclination : Inclination : 28.5 deg 7000 8000 28.5 deg 24 deg 20 deg 6500 24 deg 7500 20 deg 16 deg 6000 16 deg 7000
5500 6500
5000 6000
4500 5500 4000 5000 3500 4500 3000 4000 2500 3500 2000 3000 20000 30000 40000 50000 60000 70000 80000 20000 30000 40000 50000 60000 70000 80000 Apogee Altitude (km) Apogee Altitude (km)
Figure 2.3.11 Payload capability for GTO Figure 2.3.12 Payload capability for GTO mission (H2A2024 with 4S fairing) mission (H2A212 with 5S fairing)
2-14 Payload mass(kg) Payload mass(kg) � 6000 6000 � Inclination：� � Inclination：� � 28.5 deg 28.5 deg 24 deg 5000 24 deg 5000 20 deg 20 deg 16 deg 4500 16 deg 4500 4000 4000 3500 3500 C 3000 3000 2500 2500 2000 2000 1500 1500
1000 1000
500 500
0 0 �20000 �30000 �40000 �50000 �60000 �70000 �80000 � �20000 �30000 �40000 �50000 �60000 �70000 �80000 � Apogee Altitude�(km) Apogee Altitde (km) Figure 2.3.13 Payload capability for Figure 2.3.14 Payload capability for GTO GTO mission (H2A202 with 4S fairing) mission (H2A2022 with 4S fairing)
�
Payload mass (kg) 6000
5500
5000
4500
4000
3500
3000 � 2500 Inclination：� 2000 28.5 deg 1500 24 deg 20 deg 1000 16 deg
500
0 � 20000 � 30000 � 40000 � 50000 � 60000 � 70000 � 80000 � Apogee Altitude (km)
Figure 2.3.15 Payload capability for GTO mission (H2A2024 with 4S fairing)
2-15 2.4 Sun-Synchronous Orbit (SSO) Mission
2.4.1 Payload capability Payload mass for SSO mission using 5S fairing, with the LE-7A lower nozzle C skirt is about 4,400 kg (except summer) based on parameters in §2.4.2. And payload mass for SSO mission using 5S fairing, without the LE-7A lower nozzle skirt is about 4,200kg (except summer) based on parameters in §2.4.2. Payload mass using model 4S fairing is about 300 kg greater, with the LE-7A lower nozzle skirt. Payload mass using model 4S fairing is about 250 kg greater, without the LE-7A lower nozzle skirt. Figure 2.4.1 shows payload capability of H2A202 / 5S fairing for SSO mission for both with and without the LE-7A lower nozzle skirt.
Typical orbital parameters 2.4.2 Typical orbital parameters for SSO mission are as follows : A Circular orbit altitude h = 800.0 km Inclination i = 98.6 ° B
Injection accuracies 2.4.3 Typical injection accuracies for SSO mission are as follows based on A parameters in § 2.4.2. Semi-major axis ∆ h= ± 10.0 km Inclination ∆ i= ± 0.18 ° Eccentricity ∆ E = 0 ~ 0.001 B (These values are 3-sigma level.) A
Typical trajectory 2.4.4 Figure 2.4.2 shows a typical flight trajectory for SSO missionin the case of A H2A202 with 5S fairing. B A
2-16 Payload mass（kg）�
8000�
7500� With the LE-7A lower nozzle skirt 7000� Without the LE-7A lower nozzle skirt 6500�
6000�
5500�
5000�
4500�
4000�
3500�
3000�
2500� 2000 � C 200��������������������� 300 400 500 600 700 800 900 1000 1100 1200 Apogee Altitude（km）�
Figure 2.4.1 Payload capability for SSO mission (H2A202 with 5S fairing except summer)
MECO SEIG 0
Tanegashima Space Center 1 0 0 0 40 2 0 SECO 0 0 30 400 3
0 20 600 0 0 S/C sep 10 800 4
0
0
0 0 1000 R
a
n
5
g
0 e
-10 0 0
1200 (
k
m
) -20 6 Latitude (deg)
1400 0
0 0 -30 Time from lift-off (sec)
1600 7 0
-40 0 0
1800
-50 8
0
0 0
-60
90 100 110 120 130 140 150 160 170 180 9
0 0
Longitude (deg) 0
0
200
400
1 600
0 800
0 1000
0 Altitude (km) 0
Figure 2.4.2 Typical flight trajectory for SSO mission (H2A202 with 5S fairing)
2-17 2.5 Low Earth Orbit (LEO) Mission
2.5.1 Payload capability Figures 2.5.1 and 2.5.2 show payload capability for LEO mission using model C 5S fairing for two inclinations in the case of configuration " with the LE-7A lower nozzle skirt". Figures2.5.3 and 2.5.4 show payload capability for LEO mission using model 5S fairing for two inclinations in the case of configuration "without the LE-7A lower nozzle skirt".
Figure 2.5.1 : Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 30.4 deg) With the LE-7A lower nozzle skirt Figure 2.5.2 : Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 51.6 deg) With the LE-7A lower nozzle skirt Figure 2.5.1 : Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 30.4 deg) Without the LE-7A lower nozzle skirt Figure 2.5.2 : Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 51.6 deg) Without the LE-7A lower nozzle skirt
Payload mass using model 4S fairing is greater than using model 5S fairing but the increase of payload mass is depends on the injection orbit.
2-18 Payload mass (kg) B 12000 A 11000 Circular orbit : 1 burn mode 10000 Elliptical orbit : perigee altitude 9000 500 km
8000 700 km
7000 1000 km > 6000
5000
4000
3000
2000
1000
0 100 1000 10000 100000
Apogee altitude (km) Figure 2.5.1 Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 30.4 deg) Payload mass (kg) B 12000 A 11000 Circular orbit : 10000 1 burn mode
9000 Elliptical orbit : perigee altitude 500 km 8000 700 km 7000 1000 km 6000
5000
4000
3000
2000
1000
0 100 1000 10000 100000
Apogee altitude (km) Figure 2.5.2 Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 51.6 deg)
2-19 Payload mass (kg)
12000
11000 Circular orbit : 1 burn mode 10000 Elliptical orbit : perigee altitude
9000 500 km 700 km 8000 1000 km
7000 C 6000
5000
4000
3000
2000
1000
0 100 1000 10000 100000 Apogee altitude (km) Figure 2.5.3 Payload capability for LEO mission (H2A202 with 5S fairing)
Payload mass (kg) (inclination 30.4 deg) 12000
11000 Circular orbit : 1 burn mode 10000 Elliptical orbit : perigee altitude 500 km 9000 700 km 8000 1000 km
7000
6000
5000
4000
3000
2000
1000
0 100 1000 10000 100000 Apogee altitude (km) Figure 2.5.3 Payload capability for LEO mission (H2A202 with 5S fairing) (inclination 51.6 deg)
2-20 2.6 Earth Escape Mission
2.6.1 Payload capability Figure 2.6.1 shows payload capability for Earth Escape mission using model C 4S and 5S fairing (See § 4.4). In this figure, the basic configuration is the two stage configuration.
Payload mass (kg)
6000
5000
Payload fairing type 4S fairing 4000 5S fairing
3000
2000
1000
0 -30 -20 -10 0 10 20 30 40 50
2 C3 (km/sec)
Figure 2.6.1 Payload capability for earth escape mission (H2A202)
2-21 2.7 Spacecraft Orientation and Separation
2.7.1 General description When the second stage is injected to the planned orbit, the second stage provides the required orientation and spin (if required, see the § 2.7.3) for the A spacecraft before its separation. After completion of the spacecraft separation, the second stage conducts the CAM to avoid collision and contamination to the A spacecraft.
2.7.2 Separation sequence (1) Before separation The second stage attitude is controlled by the guidance and control system according to the mission plan (desired direction, rate and sequence) with three axis direction using the reaction control (gas jet) system (RCS).
(2) Separation The second stage guidance control computer (GCC2) sends the separation command to the second stage sequence distribution box (SDB2), and SDB2 sends an ignition current to the separation system (pyrotechnics of clamp band or separation nuts). The separation is monitored by two separation switches located at the top of the adapter. The separation switches sense the spacecraft separation, and telemetry sends the separation event data to the ground station.
(3) After separation After spacecraft separation, the second stage conducts the CAM using the RCS and the residual GH2 in the LH2 tank (in addition, sometimes, residual GHe in the ambient helium bottle is used.)
2.7.3 Spin-up performance The second stage RCS can rotate the launch vehicle body and a spacecraft with speed up to 5 rpm clockwise or counterclockwise before spacecraft separation. The user should coordinate details with NASDA if spin-up is needed. If the user needs a spin-rate exceeding 5 rpm, 937M-spin or 937M-Spin-A A adapter with a spin table attached are available.
2.7.4 Pointing accuracy Pointing accuracy (error from the nominal value) just before spacecraft separation is shown below. Following values are for roll, pitch and yaw axes.
2-22 Without pointing maneuvering : Less than ± 10 ° With pointing maneuvering : Less than ± 3 ° (When spin-up is performed, above data is available before spin-up) Attitude rate accuracy (before separation) : Less than 0.3 °/sec for non-spin spacecraft separation Attitude rate accuracy (before separation) : Less than 0.25 °/sec for spin spacecraft separation
2.7.5 Relative separation velocity Relative velocity at spacecraft separation is nominal about 0.5 to 1.0 m/sec. A If the user needs a different velocity, the launch vehicle can provide it as long as the new velocity does not inhibit launch vehicle avoidance maneuvers. For upper spacecraft separation, the relative velocity will be set automatically by mission requirements. For all missions, NASDA will require a spacecraft attitude control and mission control plan in order to analyze and modify launch vehicle movement after spacecraft separation.
2.7.6 Separation tip-off rate Separation tip-off rate depends on the separation mechanism. With the spring ejection mechanism, springs installed in adapter push the spacecraft at separation. With the retro thruster mechanism, the second stage of launch vehicle moves backwards using retro thrusters at separation; there are no springs in the adapter. Separation spring mechanism Tip-off rate : Less than 2.0 °/sec Retro thruster mechanism Tip-off rate : Less than 1.0 °/sec
2.7.7 Dual launch sequence Figure 2.7.1 shows a typical separation sequence for dual launch on LEO (upper)-GTO (lower) mission. Figure 2.7.2 shows a typical separation sequence for dual launch on GTO (upper)-GTO (lower) mission.
2-23 Upper Second stage spacecraft engine cutoff separation Second stage (SECO1) engine ignition (SEIG1)
Coasting
Lower fairing jettison CAM
Lower Second stage spacecraft engine cutoff separation (SECO2)
Coasting
Second stage engine restart (SEIG2)
CAM
Figure 2.7.1 Typical separation sequence for dual launch on LEO-GTO mission
2-24 Upper Second stage spacecraft engine cutoff separation (SECO2) Second stage engine restart (SEIG2) Coasting
CAM
Lower fairing jettison
Lower spacecraft separation
CAM
Figure 2.7.2 Typical separation sequence for dual launch on GTO-GTO mission
2-25 C H A P T E R 3 .
E N V IR O N ME N T S
C E N
H V
A I
R
P O N
T M
E E
R N
T
3 S . CHAPTER 3. ENVIRONMENTS
3.1 General This chapter describes the environments to which the spacecraft will be exposed in prelaunch and flight phases. The spacecraft has to be designed and tested before launch according to these conditions. So the spacecraft test requirements are also described.
3.2 Mechanical Environments
3.2.1 General During lift-off and in-flight phases, the spacecraft is exposed to static and dynamic loads caused by the launch vehicle. These environments cover the phases of ground transportation in Tanegashima island (including transfer to the launch pad), encapsulation of the spacecraft and mating spacecraft to the second stage. The load factors described in § 3.2.2 to 3.2.5 should be considered as limit loads applied to the spacecraft.
3.2.2 Combined load factor A Figure 3.2.1 shows a typical longitudinal static acceleration flight time history. B During lift-off and in-flight phases, dynamic acceleration excited by aerodynamic factors (such as winds, buffeting at transonic phase) and / or forces of the propulsion systems (thrust buildup or tail-off transients of the first stage main engine or SRB-A, etc.) is imposed on the spacecraft. So a combination of static acceleration (the quasi-static acceleration) and low-frequency dynamic A acceleration should be considered as the design limit load factors for the spacecraft primary structure. For the H-IIA rocket, the combined load factor at lift-off and at main engine (of A the first stage) cutoff (MECO) transient cover the maximum loads of the spacecraft primary structure during lift-off and in-flight phases. Table 3.2.1 shows the load factors of 3-sigma high values. Lateral and longitudinal loads may act simultaneously during any phase. For secondary structures of the spacecraft which have low natural frequencies, load factor on the structures may exceed the above load factors. The acceleration distribution within the spacecraft should be determined using the CLA results. Therefore the user should discuss loads conditions on the spacecraft structures with NASDA.
3-1 3.2.3 Sinusoidal vibration The spacecraft is exposed to vibration environments that may be divided into two general frequency ranges as follows: (1) low-frequency sinusoidal vibration (2) high-frequency random vibration In this section, (1) is described. And (2) is described in § 3.2.4. The levels shown below are enveloped levels of the low frequency vibrations which A are exerted during lift-off and in-flight phases, in particular at lift-off, maximum Q α’ (which means the maximum product of dynamic pressure and total angle of attack), MECO, the first and the second stage separation, etc. These levels are B prescribed at the spacecraft interface (spacecraft separation plane).
Limit level (3-sigma high) Longitudinal 1.0 G0-P for 5 to 30 Hz A 0.8 G0-P for 30 to 100 Hz Lateral 0.7 G0-P for 5 to 18 Hz 2 0.6 G0-P for 18 to 100 Hz G = 9.80665 m/s B (Excitation is applied at the base of the adapter with a 4 octave/min sweep rate in up and down direction so that vibration levels at the spacecraft interface are equal to the above levels.)
These conditions do not include the influence of steady acceleration, therefore additional evaluations for this influence are necessary. If there is possibility during testing that the structure is subjected to overloads due to differences between the flight configuration and the vibration test configuration, a notching procedure will be allowed to avoid overloads. Notching conditions are defined in detail according to the coupled loads analysis (CLA). As to the structure for which sinusoidal vibration environmental test cannot simulate the dynamic load of the flight condition sufficiently, confirmation of the strength shall be carried out by tests and analyses separately.
3.2.4 Random vibration Spacecraft structure experiences the random vibration (high-frequency), which is primarily caused by the acoustic noise described in § 3.2.5. If random vibration conditions at the base of the spacecraft are required, contact NASDA.
3.2.5 Acoustics The spacecraft is exposed to an acoustic environment during the first stage
3-2 phase until the vehicle ascends to the altitude where an atmospheric influence can be disregarded. Random vibrations are generated by the noise of the first stage main engine and the SRB-A, and the pressure vibration caused by buffeting and boundary layer noise during the phase of the transonic flight and the high dynamic pressure. Figure 3.2.2 and Figure 3.2.3 show the acoustic environment level inside the payload fairing for H2A202 series and H2A212 respectively. This is the envelope A level during launch and flight and defined as 2-sigma high. This level is uniform around the spacecraft. The spacecraft should be able to endure to this level for 60 seconds. The reference point 0 dB of sound pressure level (SPL) is equivalent to 20 μPa. A
3.2.6 Shock Table 3.2.2 shows a summary of pyrotechnic shock events during flight on all kinds of H-IIA vehicle. A type and a location of each separation system are different according to each shock event. In these separation systems, the spacecraft separation device located at the spacecraft separation plane produces the highest shock. Figure 3.2.4 and Figure 3.2.5 show a typical spacecraft separation shock spectrum at the spacecraft separation plane with 1194M adapter and 2360S A adapter respectively. These shock spectrums are exerted uniformly in all directions.
3-3 50
45
40 ) 2 35
30
25
20
15 Acceleration (m/sec
10
5
0 0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800
Time from lift-off (sec)
Figure 3.2.1 Typical longitudinal static acceleration for GTO (H2A202)
3-4 Table 3.2.1 Combined loads (Limit level)
Event Acceleration Remarks
B Combined loads * Longitudinal 3.2 G 1.7 G (steady) + (1.5 G (dynamic)) for compression Lateral ± 1.8 G Lift-off Combined loads * Longitudinal 0.1 G for tension Lateral ± 1.8 G
Immediately Longitudinal 4.0 G before Lateral ± 0.5 G MECO At H2A202, H2A2022, H2A2024 MECO A MECO Longitudinal + 1.0 G Transit Lateral ± 1.0 G
B G = 9.80665 m/s2 * : Maximum load at the top of the payload adapter Lateral : ± may act in either direction As for the longitudinal loads, all of them are difined with the tension loads as positive. B
3-5 Center Frequency (Hz) SPL (dB)
31.5 125
63 126.5
125 131
250 133
500 128.5
1000 125
2000 120
4000 115
8000 113
OASPL 137.5
The reference point 0 dB = 20 μPa This level is defined as 2-sigma high
150
140 OA
130
SPL (dB) 120
110
100 31.5 63 125 250 500 1000 2000 4000 8000 Center Frequency (Hz)
Figure 3.2.2 Sound pressure level inside fairing with acoustic blanket A (H2A202, H2A2022, H2A2024)
3-6 Center Frequency (Hz) SPL (dB)
31.5 128
63 129.5
125 134
250 136
500 131.5
1000 128
2000 123
4000 118
8000 116
OASPL 140.5
The reference point 0 dB = 20μPa This level is defined as 2-sigma high
150
OA 140
130
SPL (dB) 120
110
100 31.5 63 125 250 500 1000 2000 4000 8000 Center Frequency (Hz)
Figure 3.2.3 Sound pressure level inside fairing with acoustic blanket (H2A212) A
3-7 Table 3.2.2 Summary of pyrotechnic shock events
A Pyrotechnic shock events H2A202 H2A2022 H2A2024 H2A212 H2A222 Remarks (REF)
SRB-A separation
First SSB pair separation N/A N/A N/A
Second SSB pair separation N/A N/A N/A N/A
LRB separation N/A N/A N/A
upper fairing in dual Fairing jettison launch
First / Second stage separation
upper spacecraft in dual Spacecraft separation launch
Fairing jettison (lower) in dual launch
Spacecraft separation (lower) in dual launch
(lower) means adaptable in case of dual launch. N/A means ‘not applicable’.
3-8 10000 1500Hz 3000Hz
4100G
1000 Acceleration (G)
100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2 B
Figure 3.2.4 Typical spacecraft separation shock spectrum with 1194M adapter A
3-9 10000
2000G
1000 Acceleration (G)
Q=10
100 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2 B
Figure 3.2.5 Typical spacecraft separation shock spectrum with 2360S adapter A
3-10 3.3 Thermal Environment
3.3.1 General Concerning thermal environments, four phases have to be considered. These are as follows : a) the phase of spacecraft preparation in the Ground Support Equipment (GSE) buildings and transport between these buildings (Refer to the GSE Manual for details) b) the phase of spacecraft encapsulation into the payload fairing (PLF) and mating to the launch vehicle in the Vehicle Assembly Building (VAB). c) the phase of transfer to the launch pad and the final prelaunch d) the in-flight phase The phases from a) to c) are concerned with prelaunch environment and d) A concerns launch and flight environment.
3.3.2 Prelaunch environment The spacecraft thermal environment is controlled during prelaunch activities such as spacecraft integration, functional test, propellant loading, mating to the payload adapter (PLA), encapsulation in the PLF, mating to the launch vehicle, final checkout, transport to the launch pad, etc. Environments during these activities are shown in Table 3.3.1. Environments in the No. 2 spacecraft test and assembly building (STA2) (or No. 1 spacecraft test and assembly building (STA1)), in which spacecraft integration and functional tests are main activities, are controlled at 22 ± 3 ºC (71.6 ± 5.4 ºF) and 50 ± 10 % relative humidity. A During the ground transport from the STA2 (or STA1) to the spacecraft and fairing assembly building (SFA) (or the third stage and spacecraft assembly building (TSA)), the temperature inside the PLF is not controlled because the spacecraft is loaded in the container (which is shielded completely) and the transport is carried out very quickly (about 30 minutes, not including preparation time in the STA2 or STA1). Environments in the SFA are controlled according to the kind of each activity. In case of a normal operation, the temperature is controlled at 21 ± 3 ºC (69.8 ± 5.4 ºF). For each operation of battery charging and propellant loading, the A temperature is controlled at 18 ± 3 ºC (64.4 ± 5.4 ºF) and at 21 ± 1 ºC (69.8 ± 1.8 ºF) respectively. The relative humidity is controlled at 45 ± 5 % during all activities. During the ground transport from the SFA to the VAB, the temperature inside the PLF is not usually controlled because the spacecraft is loaded in the PLF A (which is shielded completely) and the transport is conducted very quickly (about one hour, not including preparation time in the SFA). If required, A condition is controlled (option).
3-11 During operations of hoisting the encapsulated spacecraft and mating it onto the launch vehicle, air conditioning inside the PLF is not carried out for the reason that environments in the VAB are controlled at 24 ± 3 ºC (75.2 ± 5.4 ºF) (in A Summer), 20 ± 3 ºC (68.0 ± 5.4 ºF) (in Winter), and 50 ± 10 % relative humidity. After the air conditioning duct is connected to the PLF, environments inside the PLF are controlled at 10 – 25 ± 2 ºC (50 – 77 ± 3.6 ºF) and 40 – 50 % relative A humidity until lift-off (including a transfer phase from the VAB to the launch pad (LP)).
3.3.3 Launch and flight environment (1) Within fairing (from lift-off to fairing jettison) The spacecraft is protected by the payload fairing during ascent to a nominal altitude of about 130 km (430 kft). But aerodynamic heating heats the fairing B surface and the spacecraft receives a time-dependent radiant heating environment from the internal surface of this fairing prior to fairing jettison. Figure 3.3.1 shows typical internal surface temperature profiles of the 4S fairing or 4/4D upper fairing for a Geostationary Transfer Orbit (GTO) mission. C And Figure 3.3.2 shows typical internal surface temperature profiles of the 5S fairing or 5/4D upper fairing for a GTO mission. And Figure 3.3.3 shows typical C internal surface temparature profiles of the 5S-H fairing for a H-II Transfer Vehicle (HTV) mission. Further, Table 3.3.2 shows the maximum temperature and emittance of internal surfaces of the payload fairing during ascent flight. The maximum heat flux radiated by the fairing is less than 500 W/m2, and the maximum B temperatures remain below 140 ºC at the warmest location (the nose cap). C
(2) After fairing jettison The nominal timing for jettisoning the fairing on all flights is determined so that the maximum free molecular heat flux does not exceed 1135 W/m2. A typical free molecular heating profile for a GTO mission is shown in Figure 3.3.4. This C heat flux is evaluated as a free molecular flow acting on a plane surface perpendicular to the velocity vector and is based on a standard atmospheric model. The spacecraft thermal environment after jettisoning the fairing includes free molecular heating, solar heating, Earth heating, radiation to the second stage and deep space and thermal flux conducted from the forward end of the second stage through the spacecraft adapter. In addition, during the collision / contamination avoidance of the second stage after the spacecraft separation, the thruster plumes from the RCS and LH2 / GH2 and LOX / GOX vented through LH2 and LOX tanks might influence the spacecraft. Solar and Earth thermal heating can be controlled as required by the spacecraft by selecting launch time, vehicle attitude (including rolls) and proper mission design. Thermal heat conducted from the second stage, and influence of RCS plumes and expelled LH2 / GH2 and LOX / GOX are usually small. These values are estimated in the mission analysis.
3-12 Table 3.3.1 Gas conditioning capabilities
Inside building Inside payload fairing
Phase Room temp Flowrate Inlet Relative Inlet temp Flowrate Inlet Relative pressure pressure capability capability humidity Cleanliness capability *1 capability humidity Cleanliness capability capability ° ° 3 ° ° 3 C ( F) Sm /min kPa % C ( F) Sm /min kPa % B 22±3 class Ñ Ñ 50±10 N/A N/A N/A N/A N/A STA2 (71.6 ± 5.4) 100000
Transport N/A N/A N/A N/A N/A N/A N/A N/A N/A N/A from STA2 to SFA
21±3 class Normal operation Ñ Ñ 45±5 (69.8 ± 5.4) 100000 A
18±3 class Battery charging Ñ Ñ 45±5 N/A N/A N/A N/A N/A (64.4 ± 5.4) 100000 SFA Propellant 21±1 class Ñ Ñ 45±5 loading (69.8 ± 1.8) 100000
After 21±3 class Ñ Ñ 45±5 According to room circumstances encapsulation (69.8 ± 5.4) 100000
Transport *2 10Ð25(±2) Upper : 50 class N/A N/A N/A N/A N/A 7 40Ð50 from SFA to VAB (50 Ð 77 (± 3.6)) Lower : 50 5000 B
24 ± 3 (in Summer) A (75.2 ± 5.4) class 10Ð25(±2) Upper : 50 class VAB Ñ Ñ 50±10 7 40Ð50 20 ± 3 (in Winter) 100000 (50 Ð 77 (± 3.6)) Lower : 50 5000 B (68.0 ± 5.4) A Transport 10Ð25(±2) Upper : 50 class N/A N/A N/A N/A N/A 7 40Ð50 from VAB to LP (50 Ð 77 (± 3.6)) Lower : 50 5000 B
3-13 Table 3.3.2 Maximum temperatures and emittances of fairing internal surface
Max. temp. (℃) Emittance Section (T.B.D.) (T.B.D.)
Nose cap 140 0.15 B
Nose cone 100 0.15 B
Cylinder 120 0.15 B
Acoustic blanket 34 0.85
3-14 Cylinder section Nose cone section A Nose cap section
Stagnation point
200 B
150 Stagnation point
Nose cap 100 Nose cone
50 Cylinder Fairng internal temperature (¡C) 0 0 50 100 150 200 250 300 350 Time from lift-off (sec) Figure 3.3.1 Internal surface temperature profiles of 4S or 4/4D upper fairing for GTO mission
Second nose cone section Cylinder section A First nose cone section B
Stagnation point A 200
150 Stagnation point
Cylinder 100 No.2 nose cone
50 No.1 nose cone Fairng internal temperature (¡C) 0 0 50 100 150 200 250 300 350 Time from lift-off (sec) Figure 3.3.2 Internal surface temperature profiles of 5S or 5/4D upper fairing for GTO mission
3-15 Second nose cone section Cylinder section C First nose cone section
Stagnation point
200
Stagnation point 150
100
No.2 Nose cone No.1 Nose cone 50 Cylinder Fairng internal temperature (¡C) 0 0 50 100 150 200 250 300 350 Time from lift-off (sec)
Figure 3.3.3 Internal surface temperature profiles of 5S-H fairing for HTV mission
3-16 10
C )
2
1
0.1 Aerothermal heat flux (kw/m
0.01 0 200 400 600 800 1000 1200 1400 1600 Time from fairing jettison (sec)
Figure 3.3.4 Aerothermal heat flux of the free molecular flow (GTO mission) C
3-17 3.4 Fairing Internal Pressure Environment Air inside the payload fairing is vented during the ascent phase through one- way vent ports. Typical predicted internal pressure profiles and pressure decay rate profiles for the 4S fairing are shown in the following figures respectively. a) Figure 3.4.1 : Typical internal pressure profile for GTO mission in case of the 4S fairing b) Figure 3.4.2 : Typical pressure decay rate profile for GTO mission in case of the 4S fairing c) Figure 3.4.3 : Typical internal pressure profile for SSO mission in case of the 4S fairing d) Figure 3.4.4 : Typical pressure decay rate profile for SSO mission in case of the 4S fairing
As shown in these figures, the pressure decay rate typically varies while the launch vehicle is in the transonic flight phase. Maximum pressure decay rates for the 4S fairing are as follows : a) GTO mission : 2.50 kPa/sec C b) SSO mission : 2.57 kPa/sec
The vent area of the launch vehicle payload adapter and fairing is designed assuming that the spacecraft vents some amount of internal volume through the payload adapter. If a user requires such venting, the user should check with NASDA. In case of the 1194 adapter, it has two vent holes whose diameter is 45 mm respectively (refer to A3.3).
3-18 120
100
80
60
40
Fairing internal pressure (kPa) 20
0 0 102030405060708090100 Time from lift-off (sec)
Figure 3.4.1 Typical internal pressure profile for GTO mission (Model 4S fairing)
3-19 0
-0.5
-1
-1.5
-2
-2.5
-3
Pressure decay rate (kPa/sec) -3.5
-4
-4.5
-5 0 102030405060708090100 Time from lift-off (sec)
Figure 3.4.2 Typical pressure decay rate profile for GTO mission (Model 4S fairing)
3-20 C 120
100
80
60
40
Fairing internal pressure (kPa) 20
0 0 102030405060708090100 Time from lift-off (sec)
Figure 3.4.3 Typical internal pressure profile for SSO mission (Model 4S fairing)
3-21 0 C
-0.5
-1
-1.5
-2
-2.5
-3
Pressure decay rate (kPa/sec) -3.5
-4
-4.5
-5 0 102030405060708090100 Time from lift-off (sec)
Figure 3.4.4 Typical pressure decay rate profile for SSO mission (Model 4S fairing) A
3-22 3.5 Contamination and Cleanliness
3.5.1 Prelaunch contamination and cleanliness The launch vehicle hardware which may affect the spacecraft environment shall be designed, manufactured and handled according to strict contamination control guidelines. This hardware is defined as a contamination critical item, and includes the payload support structure, the payload adapter and the interior surface of the payload fairing (PLF). In addition, operation activities at the launch site from arrival at the STA2 (or STA1) and air supplied to the spacecraft are also controlled strictly according to the contamination control guidelines.
(1) Contamination control before spacecraft encapsulation Surfaces of the contamination critical hardware items are cleaned and inspected to maintain cleanliness of less than Class 100,000 conditions. Supplied gas (air) is also controlled at less than Class 100,000 conditions through a filter.
(2) Contamination control after encapsulation After the spacecraft is encapsulated in the PLF, the payload compartment inside the PLF is completely closed up except access doors or large doors. And then, air controlled at less than Class 5,000 conditions is supplied through the air conditioning duct and maintained in higher pressure condition than surrounding pressure of the PLF.
3.5.2 Flight contamination control The sequence of the spacecraft separation and the collision / contamination avoidance maneuver is carried out as described in § 2.7.2. In this phase, the RCS is used for attitude control and propellant settling (if required, used as retro- motors to move the vehicle away from the spacecraft after separation). As part of the maneuver, GH2 is expelled from the LH2 tank through the propulsive vent port to increase second stage / spacecraft separation distance. Further, for safety disposal of the second stage, LH2 and LOX are expelled from each tank through each engine chill down port which is nonpropulsive, GOX is expelled from the LOX tank through the vent port (nonpropulsive), GH2 is expelled from the LH2 tank through the nonpropulsive or propulsive vent port, and cryogenic GHe is also discharged from GHe bottles through the LOX tank. The installation location and cant angle of RCS thrusters, vent ports and chill down ports are roughly shown in Table 3.5.1. The RCS is of a module type and two RCS modules are installed on the component equipment panel located at the
3-23 aft end of the second stage LOX tank and are located in vehicle axis symmetry. One RCS module consists of four 50 N hydrazine (N2H4) thrusters. (If required, six more 50 N thrusters and one 4 N thruster are added per module.) The thruster’s exhaust gas is mainly composed of ammonia, nitrogen and hydrogen. Use of RCS thrusters is restricted to minimize contamination products. The expelled products from each vent and chill down port are hydrogen, oxygen and a small amount of helium which are almost noncontaminating to the spacecraft. If necessary, a contamination analysis is carried out. Outgases from A inside the fairing, which are measured per ASTM E595-77*1, are consistent with NASA regulation.
3-24 Table 3.5.1 Source of gaseous contaminant to spacecraft and its installation location
Gas producting item / quantity Products Purpose Location and cant angle to X axis *1 Remarks
< RCS module > Under surface of the component equipment panel A 50 N pitch / roll thruster : 4 (4) Attitude control 90 deg Ammonia, 50 N yaw thruster : 2 (2) 90 deg ( ) means Nitrogen, 50 N settling / retention thruster : 2 (2) Propellant settling 164.4 deg *2 option thruster Hydrogen 4 N settling / retention thruster : 0 (2) 164.4 deg *2
50 N retro-thruster : 0 (4) Retrogression 45 deg *2
< LH2 / GH2 vent >
Engine chill down port : 1 Hydrogen Engine chill down and Aft end of LH2 tank : about 90 deg
residual LH2 expulsion
LH2 tank vent port : 1 Hydrogen Residual GH2 expulsion Aft end of LH2 tank : about 90 deg A
< LOX / GOX vent >
Engine chill down port : 1 Oxygen Engine chill down and Aft end of LH2 tank : 90 deg residual LOX expulsion
LOX tank vent port : 1 Oxygen Residual GOX expulsion Aft end of LH2 tank : 90 deg
< CHe vent >
LOX tank vent port : 1 Helium Residual CHe discharge Use LOX tank vent port through LOX tank
3-25 3.6 Radiation and Electromagnetic To ensure that electromagnetic compatibility (EMC) is achieved for each launch, the electromagnetic environment is thoroughly evaluated. The spacecraft and the launch vehicle system must prevent mutual disturbance for devices and wiring of each system, and be designed to endure any anticipated disturbance. The spacecraft system will be required to provide all data necessary to support EMC analysis employed for this purpose.
3.6.1 Launch vehicle generated radio environment Launch vehicle intentional radiations are limited to the UHF-band telemetry transmitters at 8 W (nominal), the VHF-band telemetry transmitter at 3 W AC (nominal), the radar transponder at 400 W peak (nominal) and the SHF-band telemetry transmitter at 2 W (nominal). Launch vehicle RF systems and their frequencies are as follows: VHF telemetry system 295 to 297 MHz (standard) C UHF telemetry system 2.289 to 2.291 GHz (standard) Radar transponder system 5.23 to 5.786 GHz (standard) SHF telemetry system 14.855 to 14.865 GHz (option) C
3.6.2 LV generated electromagnetic environment When the launch vehicle is not transmitting any radio signals, the H-IIA radiated emission level is below that shown in Figure 4.6.8. Detailed information C related to RF requirements is presented in § 4.6.8.2. This level is defined at the spacecraft separation plane (at the lower spacecraft A separation plane in a dual launch).
3.7 Spacecraft Compatibility Test Requirements NASDA requires that the spacecraft be capable of withstanding maximum expected flight loads multiplied by minimum factors of safety to preclude loss of critical function. An environmental test report is required to summarize the performed tests and to document the adequacy of the spacecraft structure for flight loads. The spacecraft tests required for demonstration of compatibility are listed in Table 3.7.1. This table describes tests, margins, and durations appropriate for recommendation in three phases of development. The structural test model (STM) is considered a test-dedicated qualification article. The protoflight model (PFM) is the first flight article produced without a qualification or STM program. The flight model (FM) is defined as a flight article produced after the qualification or protoflight article. NASDA also suggests that the spacecraft organization demonstrate the spacecraft compatibility to thermal and EMI / EMC environments.
3-26 Flight hardware fit checks are performed to verify mating interfaces and envelopes. Table 3.7.2 identifies recommended spacecraft qualification and acceptance tests to validate adequate compliance with H-IIA environments.
3-27 Table 3.7.1 Spacecraft structural tests, margin, and duration
STM PFM FM Test (Qualification) (Protoflight) (Flight)
Static Level 1.25 x Limit 1.25 x Limit 1.0 x Limit Analyses (DLF or CLA) (CLA) (CLA)
Acoustic Level Limit + 3 dB Limit + 3 dB Limit Level Duration 2 minutes 1 minutes 1 minutes
Sine Vib Level 1.25 x Limit 1.25 x Limit 1.0 x Limit Sweep Rate 2 Oct / minutes 4 Oct / minutes 4 Oct / minutes
Shock 2 Firings 2 Firings 1 Firing
Table 3.7.2 Spacecraft qualification and acceptance tests requirement
Sine EMI / Modal Static Fit Acoustic Shock Vib EMC Survey Loads Check
Qualification
Acceptance
3-28 C H A P T E R 4 .
IN T E R F A C E R E Q U IR E ME N T S
I
N
T
E
C
R
H F
A
A C
E
P
R E
T
Q
E U
I
R R E
M
4
E
N
.
T S CHAPTER 4. INTERFACE REQUIREMENTS 4.1 General This chapter describes interface requirements between the spacecraft and the H-IIA launch vehicle. In this chapter, frequency requirements, balance requirements, mechanical requirements, electrical and radio requirements are included. All interface information given is a baseline, so slight modifications or tailoring might be permitted according to negotiation with NASDA.
4.2 Frequency Requirements
4.2.1 General To avoid dynamic coupling modes in the low-frequency between the launch vehicle and the spacecraft during the ascent phase, the spacecraft should be designed with a structural stiffness which satisfies the fundamental frequency requirements. Figure 4.2.1 shows the H-IIA launch vehicle axes used in this document for A reference.
4.2.2 Fundamental frequencies Under the assumption that the spacecraft is connected rigidly to the separation plane, its primary structure fundamental frequency should be as follows: a) Lateral direction ≥ 10 Hz b) Longitudinal direction ≥ 30 Hz
If the spacecraft can not satisfy the above conditions, the spacecraft organization should discuss loads, environmental conditions and usable A volume, etc., with NASDA using a result of coupled loads analysis at the interface meetings, and confirm that there are no problems.
4-1 XB
Roll (+)
Vehicle body axes [XB YB ZB]
[IV]
[III]
Yaw (+) YB [I] Pitch (+) [IV] ZB [III] [II] Fairing separation plane (down range) is from II to IV axis. UM
PST
[I] ML UM UM, ML and PST 15m are fixed structures. [II] At new LP, PST does not exist.
Figure 4.2.1 H-IIA launch vehicle axes A
4-2 4.3 Balance Requirements
4.3.1 General This section describes balance requirements for the spacecraft design. These are related to the payload adapter strength and the center of gravity (CG) of the spacecraft. The CG of the spacecraft concerns a disturbance at the spacecraft separation and spin-up unbalance for the spun-up spacecraft. In this manual, the height of the spacecraft CG means a distance from the spacecraft separation plane to the spacecraft CG.
4.3.2 Height limit of the center of gravity The height limit of the spacecraft CG is defined according to the strength requirement of each payload adapter. Therefore the height limit of CG depends on the used adapter type. 4.3.3 Balance requirements
4.3.3.1 Static balance The CG offset of the spacecraft in radial direction shall be as follows: a) Separation without spin-up Less than 25 mm from the launch vehicle center axis in radial direction b) Separation with spin-up (less than 5 rpm) Less than 25 mm from the launch vehicle center axis in radial direction B If the separation mechanism with springs is not used, the spacecraft organization should coordinate with NASDA about static balance requirements.
4.3.3.2 Dynamic balance For the spacecraft which requires to be spun-up, the angle between the principal roll inertia axis of the spacecraft and the launch vehicle roll axis (or the spacecraft spin axis) shall be less than 3°. B
4.4 Payload Fairing This section describes the H-IIA launch vehicle payload fairing, especially the configuration, usable volume, and mission modification of five fairing types. The spacecraft is located at the top of the launch vehicle second stage with a payload adapter, and encapsulated within the payload fairing for launch environment protection.
4-3 For dual launch, the upper spacecraft is located on the top of the lower fairing using an adapter for the upper spacecraft.
4.4.1 Fairing types There are 6 standard payload fairing models. C a) 4 m diameter single launch model (Model 4S) b) 5 m diameter single launch model (Model 5S) c) 4 m diameter dual launch model (Model 4/4D-LS) — upper : long type / lower : short type d) 4 m diameter dual launch model (Model 4/4D-LC) — upper : long type (same as Model 4/4D-LS) / lower : clamshell type e) 5 m / 4 m diameter dual launch model (Model 5/4D) f) 5 m diameter single launch model (Model 5S-H) C
* “4” in model names indicates 4 m diameter, “5” indicates 5 m diameter. “S” in model names indicates single launch, “D” indicates dual launch.
Table 4.4.1 shows characteristics of payload fairings. Usable volumes of the following payload fairings take into account dynamic displacement of spacecraft which satisfies frequency requirements described in § 4.2.
4.4.1.1 Model 4S This payload fairing is available for a single launch of 3.7 m diameter spacecraft. Figure 4.4.1 shows the model 4S. Figure 4.4.2 shows the usable volume of the model 4S fairing.
4.4.1.2 Model 5S This payload fairing is available for a single launch of 4.6 m diameter spacecraft. Figure 4.4.3 shows the model 5S. Figure 4.4.4 shows the usable volume of the model 5S fairing.
4.4.1.3 Model 4/4D-LS This payload fairing is available for a dual launch of 3.7 m diameter spacecraft. Figure 4.4.5 shows the model 4/4D-LS. Figure 4.4.6 shows the usable volume of the 4/4D long upper fairing. Figure 4.4.7 shows the usable volume of the 4/4D short lower fairing.
4-4 4.4.1.4 Model 4/4D-LC This payload fairing is available for a dual launch of 3.7 m diameter spacecraft. Figure 4.4.8 shows the model 4/4D-LC. Figure 4.4.9 shows the usable volume of the 4/4D clamshell lower fairing.
4.4.1.5 Model 5/4D This payload fairing is available for a dual launch of 4.6 m diameter and 3.7 m diameter spacecraft. Figure 4.4.10 shows the model 5/4D. Figure 4.4.11 shows the usable volume of the model 5/4D upper fairing. Figure 4.4.12 shows the usable volume of the model 5/4D lower fairing.
4.4.1.6 Model 5S-H C This payload fairing is available for a single launch of 4.6 m diameter spacecraft. This payload fairing is developed for the launch of HTV. Figure 4.4.13 shows the model 5S-H. Figure 4.4.14 shows the usable volume of the model 5S-H fairing.
4.4.2 Stay out zone around the payload adapter Figure 4.4.15 (1/3) shows the general configuration of the 1194M adapter, Figures 4.4.15 (2/3) and (3/3) show the stay-out zone around the 1194M adapter. C
4.4.3 Large door After encapsulation of the spacecraft in the fairing, the user can enter the fairing through a large door prepared for spacecraft operations, but this door is optional and the user should discuss with NASDA about the number and locations of this door. Standard size of large door : 600 mm x 600 mm Figure 4.4.16 shows the large door. C Aside from above large door, two large doors for launch operations of the H-IIA vehicle are located at the bottom of the cylindrical part for each fairing model. After encapsulation of the spacecraft in the fairing, the user can enter the fairing through these doors, but the user should discuss with NASDA about using these doors.
4-5 4.4.4 Mission modification
4.4.4.1 Access door After encapsulation of the spacecraft in the fairing, the user can access the spacecraft through access doors.
Standard size of access door : 450 mm in diameter (ø 450) C Optional size of access door : 600 mm in diameter (ø 600) Standard number of the access doors of ø 450 Four for model 4S, 5S, 5S-H, 4/4D and 5/4D upper fairings Two for model 4/4D and 5/4D lower fairings
Figure 4.4.17 shows the access door of ø 450 Figure 4.4.18 shows the access door of ø 600 Allowable Access door areas of ø 450 in case of several model fairing are shown in Figure 4.4.19 to Figure 4.4.26 Allowable Access door areas of ø 600 in case of several model fairing are shown in Figure 4.4.27 to Figure 4.4.34
The distance between centers of access doors shall be more than 1,000 mm. Access doors shall normally be located at the cylindrical section of the payload fairing.
4.4.4.2 Umbilical connectors Restrictions of locating the umbilical connectors are as follows: a) Center angle from I and III axes of launch vehicle axes shall be less than 14°. A b) Standard number of umbilical connectors : 2 for model 4S, 5S, 5S-H and upper fairing of model 4/4D and 5/4D C Details of umbilical connectors are described in § 4.6.6. If the spacecraft provider requires it, the umbilical connectors can be mounted on the interface plane of the payload adapter.
4.4.4.3 Transparent window After encapsulation of the spacecraft in the fairing, the user can link the spacecraft and the ground stations with the radio signals through a transparent window. Material of transparent window : Glass fiber reinforced plastic (GF skin honeycomb sandwich) Size of transparent window : 450 mm diameter Standard number of transparent windows : 1 A
4-6 Figure 4.4.35 shows the transparent window. C It may be installed in the same area as the access door. Instead of using a transparent window, users can link the spacecraft and the ground stations with the radio signals directly with an internal / external antenna A connected through a coaxial cable. It can transmit and receive radio signals before and after lift-off as a substitute for transparent windows. Standard number of internal / external antennas : zero A
4.4.4.4 Internal antenna After encapsulation of the spacecraft in the fairing, the user can link the spacecraft and the ground stations using the antenna inside the fairing. It can receive radio signals from the spacecraft and can transmit radio signals from the ground stations to the spacecraft. Standard number of internal antennas : one Provided model 5S-H : two C Figure 4.4.36 shows typical installation of an internal antenna. The internal antenna is connected to Ground Support Equipment (GSE) through coaxial lines.
4.4.4.5 Separate air conditioning If fairings will be used for a dual launch, an umbilical ventilation inlet, bulkhead and relief valves can be installed in the lower fairing to provide separate air conditioning between the upper and lower fairings. Size of ventilation inlet : 400 mm diameter Standard number of ventilation inlets : one for model 4S, 5S, 5S-H and upper fairing of model 4/4D and 5/4D C one for lower fairing of model 4/4D and 5/4D Size of relief valve : 120 mm (in case of 45 º type of the payload support structure) 160 mm (in case of 25 º type of the payload support structure) Number of relief valves : two
4.4.4.6 Acoustic blankets Acoustic blankets can be attached to the inner surface of the fairing to reduce acoustic sound pressure level. Acoustic blankets are made of glass fiber and the cover material. Thickness and size of a blanket are determined according to A user’s requirements. The standard blanket is 10 mm thick. Figure 4.4.37 shows a typical configuration of an acoustic blanket. C A
4-7 Table 4.4.1 Characteristics of payload fairings
Items External Usable volume Portion Launch Application Height Diameter of fairing Height Diameter Model (m) (m) (m) (m) NASDA 4S Single 12.0 4.07 Ð 10.23 3.7 ETS-VI, COMETS
5S Single 12.0 5.1 Ð 9.12 4.6 NASDA ADEOS
upper 8.23 3.7 NASA TRMM 4/4D-LS Dual 14.5 4.07 lower 3.80 3.7 NASDA ETS-VII
upper 8.23 3.7 None 4/4D-LC Dual 16.0 4.07 lower 5.36 3.7 None A
upper 6.70 4.6 ISAS SFU 5/4D Dual 14.1 5.1 / 4.07 lower 4.68 3.7 NASDA GMS-5
5S-H Single 15.0 5.1Ð 12.9 4.6 None C
4-8 φ4070
12000
7414
STA 10408 A 90
Unit : mm
Figure 4.4.1 Model 4S
4-9 R750 R565 (REF) 269
18° 18° 4430
(REF) φ3700 113 12000 * 5800 7414
* φ1194* Spacecraft
480 Separation Plane
* A °
42 110 φ2190 (REF) 1021
Unit : mm φ4070
* These values will vary with adapter model.
Figure 4.4.2 Usable volume of model 4S with 1194M adapter
4-10 φ5100
12000
9000
6937
1987.5
STA 10408 A 90
Unit : mm
Figure 4.4.3 Model 5S
4-11 R939.2 R509.2 (REF) 430 A
25° 28° 2601 (REF)
31 φ3378
15° 16.5°(REF) 2063 (REF) 31
φ4600 12000 *
φ5100 4460.5 4949.5
Spacecraft φ1194* Separation plane * * 25° 110 φ2190 (REF) 480 15° 1987.5 1965.5
φ4070 Unit : mm
* These values will vary with adapter model.
Figure 4.4.4 Usable volume of model 5S with 1194M adapter
4-12 φ4070
Upper fairing 10000
14444.5 (REF)
Lower fairing adapter
Lower fairing 4444.5
STA 10408 A 90
Unit : mm
Figure 4.4.5 Model 4/4D-LS
4-13 R750 R565 (REF) 269
18° 18° 4430 10000 (REF) φ3700 113 * 3800 * * A 5414 φ1194* Spacecraft 110 480 Separation Plane
42° 2190 (REF)
1021 φ
Unit : mm φ4070 * These values will vary with adapter model.
Figure 4.4.6 Usable volume of model 4/4D long upper fairing with 1194M adapter
4-14 (REF) φ1980 160
42° 1021 φ3640 861
1.7° 3700 (REF)
φ 1000 6.4° φ3208* (REF) * A
* φ1194* Spacecraft 1943.5
480 Separation 5465.5 (REF) Plane 3444.5
* A 42° φ2190 (REF) 110 1021
Unit : mm φ4070
* These values will vary with adapter model.
Figure 4.4.7 Usable volume of model 4/4D short lower fairing with 1194M adapter
4-15 φ4070
Upper fairing 10000
16000 (REF) A
Lower fairing adapter
Lower fairing 6000 A
STA 10408 A 90
Unit : mm
Figure 4.4.8 Model 4/4D-LC
4-16 A φ2190 (REF) (REF)
φ1980 160 1021 924 Separation Plane
φ3700 * 4436 6000 * * 5500 φ1194 * Spacecraft 110 480 Separation Plane
42° φ2190 (REF) 1021
Unit : mm φ4070
* These values will vary with adapter model.
Figure 4.4.9 Usable volume of model 4/4D clamshell lower fairing with 1194M adapter
4-17 φ5100
Upper fairing 9580.5 Lower fairing adapter
14025 (REF)
Lower fairing 4444.5
STA A 90 10408
φ4070 Unit : mm
Figure 4.4.10 Model 5/4D
4-18 R939.2 R509.2 (REF) 430
25° 28° 2601 (REF)
31 φ3378 A
° 15 16.5° (REF) 2063 2063 (REF) 31
9580.5 φ4600 * * φ1194 * A
* 2041 Spacecraft 2530 110 480 Separation Plane
25° φ2190 1965.5 1987.5 Unit : mm
φ4070
* These values will vary with adapter model.
Figure 4.4.11 Usable volume of model 5/4D upper fairing with 1194M adapter
4-19 φ1980
25° 25.5° (REF) 1740 1965.5 φ3640
1.7° φ3700 (REF) 1000 6.4° φ3208* (REF) * A 6410 * * φ1194* Spacecraft 1943.5 110 480 Separation Plane 3444.5
42° φ2190 (REF) 1021 Unit : mm φ4070 * These values will vary with adapter model.
Figure 4.4.12 Usable volume of model 5/4D lower fairing with 1194M adapter
4-20 C φ�5100
15000
12000
9937
1987.5
STA 10408 90
Unit:mm
Figure 4.4.13 Model 5S-H
4-21 R939.2 C R509.3 430 （REF）�
25°� 28°� 2601
Φ�3378
31 15°� 16.5°� 2063
31
Φ�4600
15000
9937 7885.6 Φ�5100
Φ�3183.7 （REF）�
A 780 A 1177.4 （REF）� 1987.5 15°� Spacecraft Φ�3200 843 Separation plane 25°� 90
Φ�4070
Unit:mm
Figure 4.4.14 (1/2) Usable volume of the model 5S-H #1
4-22 Diameter of usable volume ( D) C STA 8102.3 : 4600 mm C STA 8140.8 : 4400 mm STA 8195.5 : 4200 mm STA 8290.9 : 4000 mm STA 8410.2 : 3800 mm ） STA 8610.5 : 3600 mm B 1563 TYP
TYP （ ）
（ STA 9054.3 : 3400 mm 751（TYP） 1689 B STA 9359.3 STA 9565.0 : 3360 mm
C Prohibited area (2 pl) (STA 9456 STA 9565 only) SECT A-A ） 33 318.2 3183.7 REF （ （REF）� STA 8420.5
STA 9565 1353.7
1048.7 STA 10408 Φ3360 Φ3400 Φ4070 SECT B-B ） 33 3183.7 REF 250 （REF）� （ STA 8420.5
STA 9565 STA 10408
Φ4070 SECT C-C
Unit : mm
Note : The constraints by PAF are not included. This usable volume is under development.
Figure 4.4.14 (2/2) Usable volume of the model 5S-H #2
4-23 I Micro swtich (2 pl)
II IV
Spring housing (8 pl)
Spacecraft III separation plane
φ1215 (REF) 480
φ2190 (REF)
Unit : mm
Figure 4.4.15 (1/3) General configuration of the 1194M adapter C
4-24 [I] φ3700 mm ±20°
φ1815 mm [II] [IV]
φ1215 mm ±20° [III]
φ1194 mm Spacecraft rear frame φ1000 mm Usable volume φ820 mm
φ134 mm 200 mm 100 mm 60 mm
Spacecraft 46° ° Separation 46 110 mm Plane Stay out zone 450 mm 336 mm 480 mm
15° (REF)
This area must be coordinated with NASDA.
Figure 4.4.15 (2/3) Stay-out zone around the 1194M adapter (I / III -axis ±20°) C
4-25 φ3700 mm
[I] φ1815 mm ±70°
φ1215 mm [II] [IV] Spacecraft 1194 mm ±70° φ rear frame [III] φ1000 mm Usable volume φ820 mm � φ134 mm 200 mm 60 mm
Spacecraft
separation 110 mm plane Stay out zone 46° 450 mm 336 mm 480 mm
15° 12° (REF)
This area must be coordinated with NASDA.
Figure 4.4.15 (3/3) Stay-out zone around the 1194M adapter (II / IV -axis ±70°) C
4-26 A A 600 mm (23.6 in. )
Large Door A
Aluminium panel 733.0 mm (28.9 in.) A
Aluminium honeycomb
600 mm (23.6 in.)
(80.1 in. ) SECT A-A R2035 mm
Note : This figure shows the configuration of the large door in case of the 4S fairing.
Figure 4.4.16 Large door C
4-27 AA
Aluminum panel Access door
576 mm (22.7 in.)
Aluminum honeycomb
φ450 mm (17.7 in.)
SECT A-A R2035 mm (80.1 in.)
Note : This figure shows the configuration of the access door in case of the 4S fairing.
Figure 4.4.17 ø 450 access door AC
4-28 C
AA
Aluminum panel Access door
732 mm (28.8 in.)
Aluminum honeycomb
φ600 mm (23.6 in.)
(80.1 in.) SECT A-A R2035 mm
Figure 4.4.18 ø 600 access door
4-29 Access door prohibited Access door allowed STA -892 Center of A access door
800 596
STA 2994
500
600 IV IV 600 1954
1954 400 400 400 1954 I 1954 400 III 600 600 II Unit : mm
12786
6002197 400 4002197 600 6002197 400 4002197 600 Upper edge 170 14 of cylinder A STA 500 2994 1400(REF)
Air conditioning 1900 inlet 7414
Large door* 1 Large door* 1 2100
(with air vent hole) (with air vent hole) 1300 (REF)
STA 1180.5 10408 1000 1000 1000 1000 Lower edge of cylinder IVI II III IV
Figure 4.4.19 Allowable areas of ø 450 access door on model 4S fairing C
4-30 Access door prohibited STA -842 A Access door STA 800 allowed 1408 Center of 500 access door 2250 1015
IV
IV
600
1803 600
400 400 STA 1408 500 I 400 III 400 STA 600 600 2063 2340 3471 II 500
IV IV 600 2802 600
400 400 400 I 400 III 600 600 II C Unit : mm 1400 (REF) 600 3006 400 400 3006 600600 3006 400 400 3006 600
619 374 STA 3471 Air conditioning inlet
1900 A 4949.5
Air vent hole Air vent hole Air vent hole Air vent hole 500 500 1600
STA 600 (REF) 8420.5 2503 2503 2503 2503 IV I II III IV
Figure 4.4.20 Allowable areas of ø 450 access door on model 5S fairing C
4-31 A Access door STA prohibited -4892 * 3 Access door allowed STA -3336.5 Center of * 2 access door
STA -1006 800 * 4086 3 596
STA 549.5 * 2 500
600 IV IV 600 1954
1954 400 400 400 1954 I 1954 400 III 600 600 II Unit : mm
12786
6002197 400 4002197 600 6002197 400 4002197 600 *2 STA 170 14 500 C 549.5 A *3
STA 1400(REF) -1006 Air conditioning
inlet 1900 5414
1 1 *2 Large door* Large door* 2100
(with air vent hole) (with air vent hole) 1300
STA 1100 1180.5 (REF) 5963.5 C 1150 1000 1000 1150 1150 1000 1000 1150 *3 STA IVI II III IV A 4408
4-32 12786 570 2056.5570 570 2056.5 570570 2466.5 570 2056.5 570 A 160 STA 5963.5
Separation plane 1000 ) TYP
Air conditioning 970 inlet ( ) ) 4444.5 TYP REF
Large door*1 Large door*1 ) 2200 ( TYP
(with air vent hole) 2144.5(REF) (
STA 1150 1180.5 ( 10408 1150 1150 1150 IV I II III IV
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.22 Allowable areas of ø 450 access door on model 4/4D short lower fairing C
4-33 12786 A 6701956.5 570 5701956.5 670 670 1896.5 570 5701956.5 670 60
STA 500 4408 Separation� plane Air conditioning
970 inlet 6000 4200 (REF) 3100
Large door* 1 Large door* 1 2200
(with air vent hole) (with air vent hole) 1400 (REF) STA 1180.5 10408 1150 1150 1150 1150
IVI II III IV Separation� Separation� Separation� plane plane plane
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.23 Allowable areas of ø 450 access door on model 4/4D clamshell lower fairing C
4-34 Access door prohibited STA -2867 A Access door STA allowed 800 -617 Center of 500 access door 2250
1015
IV
IV
600
1803 600
400 400 STA -617 A 500 III 400 III 400 STA 600 600 2063 2340 1446 II 500
IV IV 600 2802 600
400 400 400 I 400 III 600 600 II
16022 A 600 400 400 600 600 619374 400 400 600 STA 1446 1400 (REF)
Air vent hole Air vent hole Air vent hole 2530
500 500 Air conditioning inlet 1600 STA 2503 2503 2503 3976 IV I II III IV
600 (REF) Unit : mm
Figure 4.4.24 Allowable areas of ø 450 access door on model 5/4D upper fairing C
4-35 A
12786 570 2056.5570 570 2056.5 570570 2466.5 570 2056.5 570 160 STA 5963.5
Separation plane 1000 ) TYP
Air conditioning 970 inlet ( ) ) 4444.5 TYP REF
Large door*1 Large door*1 ) 2200 ( TYP
(with air vent hole) 2144.5(REF) (
STA 1150 1180.5 ( 10408 1150 1150 1150 IV I II III IV
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.25 Allowable areas of ø 450 access door on model 5/4D lower fairing C
4-36 STA Access door C -3842 Prohibited STA 600 -1592 Access door 500 Allowed Center of 600 Access door 600
1803
STA 1803
-1592 400 500 400
STA 1803 400 1803 500 471 600 600 400
600 600
2802 2802
400 400 2802 2802 400 600 600 400
400 400 400 400
600 3006 3006600 600 3006 3006 600 93 413 413 93 93 413 413 93 STA 500 471 3000
STA 3471 500 400 400 400 400
600 3006 3006 600 600 3006 3006 600 625 380 1322
STA 620 3471 1400 (REF)
1900 Air conditioning inlet 135 659 292 4949.5 135 135 137 1720 392 STA 8420.5 1101 1000 1237 403 603 1237
Unit：mm
Figure 4.4.26 Allowable areas of ø 450 access door on model 5S-H fairing
4-37 C
Access door prohibited Access door allowed STA -892 Center of access door
920 4086
STA 2994
620
720 IV IV 720
1653
1653 520 520 520 1653 I 1653 520 III 720 720 II
12786
7201957 520 5201957 720 7201957 520 5201957 720
STA 620 2994 1400(REF)
Air conditioning inlet 1900 7414
Large door* 1 Large door* 1 2220
(with air vent hole) (with air vent hole) 1420 (REF)
STA 1180.5 10408 1120 1120 1120 1120 IVI II III IV Unit : mm
Figure 4.4.27 Allowable areas of ø 600 access door on model 4S fairing
4-38 C
Access door STA prohibited -842 STA Access door 920 1408 allowed 620 Center of access door 2250
IV 720
IV
720
1480 0 48 1
520 STA 520 520 1408 I 1 520 III 620 480 80 14 STA 2063 720 720 3471 620 II
IV 720 720 IV
2512 2512
520 520 520 I 2512 520 2512 III 720 720 II
720 2766 520 520 2766 720720 2766 520 520 2766 720
379 134 STA 3471 1400(REF)
Air conditioning 1900 inlet 4949.5
Air vent hole Air vent hole Air vent hole Air vent hole 620 620 600 (REF) 1720 STA 8420.5 2623 2623 2623 2623 IV I II III IV Unit : mm
Figure 4.4.28 Allowable areas of ø 600 access door on model 5S fairing
4-39 Access door C STA -4892 prohibited * 3 Access door allowed -3336.5STA * Center of 2 access door STA -1006 920 * 3 4086
STA 549.5
* 2 620
720 IV IV 720 1653
1653 520 520 520 1653 I 1653 520 III 720 720 II
12786
7201957 520 5201957 720 7201957 520 5201957 720 *2 170 14
STA 620 549.5 *3
STA 1400(REF) -1006 Air conditioning inlet 1900 5414
*2 Large door* 1 Large door* 1 2220 (with air vent hole) (with air vent hole) 1420 STA 1180.5 5963.5 1270 1120 1120 1270 1270 1120 1120 1270 *3 STA IVI II III IV 4408
Unit : mm
Figure 4.4.29 Allowable areas of ø 600 access door on model 4/4D long upper fairing
4-40 12786 C 675 1846.5675 675 1846.5 675675 2521.5 675 1846.5 675 STA 5963.5
Separation plane 1000
Air conditioning
inlet 1075 ) 4444.5 REF 2305 Large door*1 Large door*1 2200 (REF)
STA (with air vent hole) 1255 1092.5 ( 10408 1255 1255 1255 IV I II III IV
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.30 Allowable areas of ø 600 access door on model 4/4D short lower fairing
4-41 12786 C
7751746.5 675 6751746.5 775 775 1746.5 675 6751746.5 775
STA 500 4408 Separation� plane Air conditioning
1075 inlet 6000 4200 (REF) 2995
1 1 Large door* Large door* 2305 1505
(REF) (with air vent hole) (with air vent hole) STA 1092.5 10408 1255 1255 1255 1255
IVI II III IV Separation� Separation� Separation� plane plane plane
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.31 Allowable areas of ø 600 access door on model 4/4D clamshell lower fairing
4-42 C
Access door STA prohibited STA -617 Access door -2867 allowed 2250 Center of access door 920 620
IV
IV
720
1480 720
1480 520 STA 520 -617
III 520 148 III 620 2063 STA 0 1480 520 1446 720 720 620 II
IV IV 720 2512 720 2512 520 520 520 2512 I 2512 520 III 720 720 II
720 2766520 520 2766720 720 2766520 520 2766 720 379 134 STA 1446 1400(REF)
Air vent hole Air vent hole 2530 600Air 620 vent hole Air conditioning
1720 inlet STA 2623 2623 2623 3976 IV I II III IV 620 (REF) Unit : mm
Figure 4.4.32 Allowable areas of ø 600 access door on model 5/4D upper fairing
4-43 12786 C 675 1846.5675 675 1846.5 675675 2521.5 675 1846.5 675 STA 5963.5
Separation plane 1000
Air conditioning
inlet 1075 ) 4444.5 REF 2305 Large door*1 Large door*1 2200 (REF)
STA (with air vent hole) 1255 1092.5 ( 10408 1255 1255 1255 IV I II III IV
Access door Unit : mm prohibited Access door allowed Center of access door
Figure 4.4.33 Allowable areas of ø 600 access door on model 5/4D lower fairing
4-44 Access door STA C -3842- Prohibited STA 920 Access door -1592- 620 Allowed Center of 720
720 Access door
1480
STA 1480 -1592 - 520 620 520
STA 1480 520 1480 620 471 720 720 520
720 720
2512 2512
520 520 2512 2512 520 720 720 520
520 520 520 520 173 173 173 173 STA 620 471
STA 3471 620
520 520 520 520
720 2766 2766 720 720 2766 2766 720 385 140 1322 STA 620 3471 1900 135 659 292 135 135 1720 152 STA 8420.5 1101 1000 997 163 363 997
Unit：mm
Figure 4.4.34 Allowable areas of ø 600 access door on model 5S-H fairing
4-45 A A
Transparent window GFRP panel
608mm (23.9 in.)
NOMEX
φ450mm (17.7 in.)
SECT A-A R2035mm (80.1 in.)
Note : This figure shows the configuration of a transparent window installed in the cylinder section of the 4S fairing. The size of the window should be equivalent when it is installed in the cone section.
Figure 4.4.35 Transparent window C
4-46 IV A I A
III II
30.44 (REF)
Support
Antenna
B
B VIEW AÐA VIEW BÐB
Unit : mm
Figure 4.4.36 Typical Installation of internal antenna C
4-47 Cover sheet (PTFE) B
Filter (Stainless steel mesh) (200×200) 2500 500
Heat seal
Acoustic blanket
Insulator (Glass Wool)
10
Velcro (Polyester)
Unit : mm
Note : Blanket size will vary depending on equipment inside the fairing.
Figure 4.4.37 Typical Configuration of acoustic blanket C
4-48 4.5 Payload Adapter
4.5.1 Configuration of payload adapter The interface parameters of the payload adapter are defined below. a) Figure and dimensions of separation plane b) Detailed dimensions of mating point of adapter and spacecraft c) Installation dimensions of separation switches
4.5.1.1 Separation mechanism The H-IIA launch vehicle provides clamp bands and separation nuts as the separation mechanism on the adapter.
(1) Clamp band mechanism The V flange of the rear frame of the spacecraft is connected to that on the top front of the adapter, and bound by clamp bands with concave V shaped blocks. At the spacecraft separation, bolts which connect the clamp bands are cut by pyrotechnic bolt cutters. The clamp band tension is designed to ensure no gaps exist between the C spacecraft and the adapter frames on the ground and in flight environment.
(2) Separation nut mechanism The spacecraft is connected to the top of the adapter at 4 or 8 points by separation nuts. To separate the spacecraft, the separation nuts are released.
4.5.1.2 Ejection mechanism Separation springs can be installed in the adapter to separate the spacecraft from the launch vehicle. If the spacecraft requires the launch vehicle to back away using retro thrusters, separation springs are not installed. When the separation springs are installed, static loads shall act on each mating structure as a limiting load, in addition to the acceleration load described in § 3.2.1.
4.5.1.3 Spacecraft separation monitoring switches The standard launch vehicle has two microswiches on the top of the adapter to monitor spacecraft separation. It is possible to locate the actuator pad of the spacecraft microswitch on the top of the adapter.
4-49 4.5.1.4 Nomenclature of payload adapter (1) “1194” means the interface diameter at the top of the adapter is 1194 mm. (2) “M” means the adapter has a clamp band separation mechanism. (3) “S” means the adapter has a separation nut separation mechanism.
4.5.2 Adapter types Table 4.5.1 shows characteristics of the payload adapters for the H-IIA launch C vehicle. Detailed information related to the payload adapter is presented in APPENDIX C 3.
4.5.3 Mission modifications
4.5.3.1 Separation springs Separation springs can be installed inside an adapter to force the spacecraft away from the launch vehicle. These springs are not applied to adapters 2360S and 3470S.
a) Number of separation springs : 4 to 8
4.5.3.2 Umbilical connectors Umbilical connectors can be installed on an adapter to meet the user’s requirements.
a) Number of umbilical connectors : 2 b) Location of umbilical connectors : Decided by discussion with user c) Specification of umbilical connectors : Described in § 4.6.6.4
4-50 Table 4.5.1 Characteristics of payload adapters
Items Spacecraft Interface Height Mass Connecting Connecting Applications diameter (mm) (kg) device points Model (mm) A 937M 945 900 90 clamp Ð None band
937MH 958 900 80 clamp Ð None band B A 937M-Spin 959 605 130 clamp Ð NASDA : GMS-5 band
937M-Spin-A 958 1000 150 clamp Ð None band B
NASA : TRMM clamp 1194M 1215 480 100 Ð NASDA : ETS-VII band MOT : MTSAT B ESA : ARTEMIS A 1666M 1666 1000 100 clamp Ð None (T.B.D.) band
1664 1666S 1000 100 separation 4 None (PCD) (T.B.D.) nuts B
2360 separation NASDA : 2360S 300 40 8 (PCD) nuts ETS-VI, COMETS B
3472 separation 3470S 1089 350 4 NASDA : ADEOS (PCD) nuts BA
clamp C 1666MA 1666 480 100 Ð None band
clamp 239M 239 100 5 Ð None band
Launch vehicle interface diameter : 2190 PCD Spin : With spin table
4-51 4.6 Electrical and RF Interface
4.6.1 General This section describes electrical interfaces such as electrical grounding between the launch vehicle and the spacecraft; service interfaces from the launch vehicle; specifications of interface connector (umbilical connector); and RF interface issues such as transparency of payload fairing, constraints of spacecraft RF transmission, etc.
4.6.2 Electrical grounding The launch vehicle and spacecraft must maintain the same electrical potential in flight, for which grounding is necessary. The grounding reference point should be located on the separation plane of the spacecraft and launch vehicle, where the grounding should be provided by mechanical contact of both sides at connection. MIL-B-5087B class S (less than 1Ω) resistance requirement is applied to the H- IIA launch vehicle grounding. Therefore, surface finish of the spacecraft separation structure should satisfy the above requirement. The spacecraft provider should contact NASDA if MIL-B-5087B class R (less than 2.5 mΩ) is required.
4.6.3 Umbilical interface
4.6.3.1 Umbilical lines for single launch mission (1) The umbilical interface block diagram between the spacecraft and the ground facilities for a single launch is shown in Figure 4.6.1. (2) Two coaxial cable lines can be provided for the spacecraft RF link. (3) The specifications for interface connectors with the spacecraft (the spacecraft umbilical connector) are defined in the Interface Control Specifications (ICS) for each spacecraft. The maximum number of connector pins is 120. (4) The interface to the ground facilities is conducted using the umbilical connector located on the payload fairing.
4.6.3.2 Umbilical lines for dual launch mission (1) The umbilical interface block diagram between the spacecraft and the
4-52 ground facilities for a dual launch is shown in Figure 4.6.2. (2) The specifications of interface connectors for the spacecraft in the upper fairing and lower fairing are the same as for a single launch A mission, § 4.6.3.1 (3). A
4.6.4 Command and power interface Table 4.6.1 shows the electrical interfaces which the H-IIA launch vehicle provides to the spacecraft.
4.6.4.1 Pyrotechnic command The H-IIA launch vehicle can provide pyrotechnic commands to the spacecraft for the following two functions :
(1) Spacecraft separation pyrotechnic command When the spacecraft organization provides a separation mechanism, the launch vehicle system can provide pyrotechnic commands for spacecraft separation. Details of design conditions, etc., are to be determined at the A interface meetings. This command is provided as a standard service when the spacecraft organization provides a separation mechanism.
(2) Other spacecraft related pyrotechnic commands If the spacecraft system needs other commands, the launch vehicle system can provide one additional command as an option. However, for dual launch, the launch vehicle hardware spare channels are limited, so special arrangements must be made at the interface meetings. The launch vehicle system can provide ten command signals in total including other pyrotechnic commands, electrical commands and dry loop commands for a dual launch.
Main electrical characteristics of the pyrotechnic command are the following and the wiring diagram for the pyrotechnic command is shown in Figure 4.6.3.
4-53 a) Battery Engine battery (26 AH) C +6 b) Voltage 28 -4 VDC A c) Ignition timing To be determined at the interface meetings d) Pulse width of igniting signal To be determined at the interface meetings e) Minimum igniting current To be determined at the interface meetings f) Recommended igniting current To be determined at the interface meetings g) Number of power cartridges To be determined at the interface meetings h) Non-igniting current To be determined at the interface meetings i) Bridge wire resistance To be determined at the interface meetings j) Insulation resistance To be determined at the interface meetings k) Insulation resistance after ignition To be determined at the interface meetings
4.6.4.2 Electrical command (discrete signal) (1) The spacecraft organization may request electrical commands (discrete signals) from the launch vehicle organization, if necessary. Total number of command signals is as described in § 4.6.4.1. If in a dual launch, both C spacecraft require these signals, special arrangements must be made at the interface meetings. These commands are an optional service to the spacecraft.
(2) The launch vehicle wiring diagram for the electrical commands is shown in B Figure 4.6.4.
(3) Main electrical characteristics of the electrical command
+6 a) Voltage 28 -4 VDC A b) Load resistance To be determined at the interface meetings c) Supply current To be determined at the interface meetings B d) Insulation resistance To be determined at the interface meetings e) Supply timing To be determined at the interface meetings f) Supply time (duration) To be determined at the interface meetings g) Spacecraft circuit condition Insulated from ground and structure
Design details will be established at the interface meetings.
4.6.4.3 Dry loop command (1) The spacecraft organization may request dry loop commands from the launch vehicle organization, if necessary. Total number of command signals is as described in § 4.6.4.1. If in a dual launch, both spacecraft require these signals, special adjustment is required at the interface meetings. These commands are an optional service to the spacecraft.
4-54 (2) The launch vehicle wiring diagram The wiring diagram of the electrical commands is shown in Figure 4.6.5.
(3) Main electrical characteristics of the electrical command A
a) Supply timing To be determined at the interface meetings b) Supply time (duration) To be determined at the interface meetings c) Supply current To be determined at the interface meetings d) Circuit resistance in launch vehicle To be determined at the interface meetings e) Insulation resistance To be determined at the interface meetings
Design details will be arranged at the interface meetings.
4.6.4.4 Power supply A If the spacecraft organization requires power supply, the launch vehicle can provide it to the extent which depends on the mission. Detailed specifications of CA power will be arranged at the interface meeting and so on.
4.6.5 In-flight telemetry
4.6.5.1 Separation status transmission The second stage telemetry system transmits spacecraft separation status signals for monitoring from the ground. Spacecraft separation can be ascertained from the signals initiated by microswitches installed on the top of the adapter at the separation plane. This is a standard service. In case of the separation after spin up by the launch vehicle (in longitudinal), it may be impracticable to monitor the spacecraft separation in real time.
4.6.5.2 Dynamic environments data transmission The dynamic environments data measured at the adapter are transmitted by the second stage telemetry system. The in-flight dynamic environments data for each spacecraft measured as standard service are listed below. The frequency is less than 100 (T.B.D.) Hz.
a) Spacecraft separation status 2 ch. B b) Temperature of adapter structure 1 ch.
4-55 c) Acceleration at adapter structure 3 ch. B
4.6.6 Interface connectors between spacecraft and launch vehicle This section specifies the interface connectors between the spacecraft and the H-IIA launch vehicle.
4.6.6.1 Interface connectors procurement responsibility A Interface connector receptacles (the spacecraft umbilical connectors) which is installed on the spacecraft, and plugs shall be provided by the H-IIA launch vehicle organization. The spacecraft organization can procure both plugs and receptacles and perform the fit checks, and the plugs may be furnished to the H-IIA launch vehicle. A
4.6.6.2 Interface connectors for single launch A Interface connectors for transferring electrical signals for a single launch shall C be installed on the payload fairing. In this case, electrical interfaces are maintained until fairing jettison. For this reason, lanyard style push-pull connectors are used for the H-IIA vehicle plugs. If interface connectors are installed on the adapter, electrical interfaces will be maintained until spacecraft separation. Figure 4.6.6 shows interface connectors for a single launch.
4.6.6.3 Interface connectors for dual launch A Interface connectors for transferring electrical signals for a dual launch shall C normally be the same as in a single launch case for the upper fairing, and on the adapter for the lower fairing. If interface connectors are installed on the adapter, electrical interfaces will be maintained until spacecraft separation. Figure 4.6.7 shows interface connectors for a dual launch.
4.6.6.4 Standard interface connector specifications The standard interface connector specifications shall conform to the requirements of NASDA-ESPC-915 or ESA/SCC SPEC. No. 3401/008, contacts B SPEC. No. 3401/009.
4-56 Connectors with the specifications shown in Table 4.6.2 shall generally be used. These connectors are manufactured by CIE DEUTSCH. Table 4.6.2 shows the standard interface connector specifications.
4.6.6.5 Other interface connector characteristics (1) Plug disconnecting characteristics at fairing jettison a) Disconnecting angle and force of the plugs: The disconnecting angle of the H-IIA vehicle plugs with lanyard shall be within ±10° at separation from spacecraft receptacles. The disconnecting force of the plugs shall be as shown below. b) Location : The interface connectors shall normally be placed within ±5° from the axis I or III. c) Angle : The connectors shall be arranged within ±2° A between the connector face and the axis II - IV of the H-IIA launch vehicle. d) Pulling direction : Nothing shall be located within ±15° of the pulling direction. e) Key position : Key position of interface connectors shall be set so that connector keys are oriented in the forward direction of the H-IIA launch vehicle.
B Shell size Min. disconnecting force (N) Max. disconnecting force (N)
3 5.3 88.3 7 6.6 88.3 12 8.8 150.0 19 13.2 167.7 27 17.7 176.5 37 26.5 194.2 61 30.9 196.6
(2) When interface connectors are located on the adapter The key position of interface connectors shall be set so that connector keys are oriented in the external radial direction of the H-IIA launch vehicle. Interface connectors are located as each adapter type defines; 2 connectors should be located 180° opposite each other. The receptacle surface should face the separation plane.
4-57 4.6.7 RF constraints
4.6.7.1 Fairing transparency for spacecraft RF communications The fairing RF transparency will be specified after the location of the spacecraft antenna is fixed.
4.6.7.2 Operating constraints The spacecraft shall not radiate narrow-band electrical fields exceeding the “acceptable H-IIA radiation susceptibility level” as the worst case of the sum A spurious level shown in Figure 4.6.8. The radiation emission level is defined at the spacecraft separation plane (at the lower spacecraft separation plane in a C dual launch).
4.6.8 Electrical and RF requirements for launch phase
4.6.8.1 Electrical requirements The spacecraft organization shall satisfy the following constraints in the final preparation phase leading up to lift-off.
(1) The spacecraft organization shall design the spacecraft so that the umbilical cable carries only low current signals at lift-off. Recommended voltage and current are 28 VDC, and less than 10 mA. B (2) The spacecraft power shall be switched from external to internal, and the ground power supply must be switched off at about 5 minutes before lift-off. Details are coordinated at the interface meeting.
4.6.8.2 RF requirements Launch vehicle on-board equipment has frequencies as follows: C VHF telemeter transmitters : 295 to 297 MHz (standard), 294 to 296 MHz (option) UHF telemeter transmitters : 2289 to 2291 MHz (standard), 2200 to 2290 MHz (option: changeable according to RF frequency of spacecraft) SHF telemeter transmitters : 14.855 to 14.865 GHz (option) Radar transponder : 5.23 to 5.786 GHz (transmission & reception) Command destruct receiver (CDR) : 400 to 500 MHz Global positioning system receiver (GPSR): 1.425 to 1.675 GHz (option) For “Hot launch” spacecraft, the spacecraft organization shall satisfy the following constraints. a) Spurious radiation interference levels from the launch vehicle and C TNSC will not exceed those given in Figure 4.6.8. Spacecraft’s acceptable spurious radiation levels shall satisfy this constraints.
4-58 b) The spacecraft telemetry frequency band above must not overlap the launch vehicle bands. c) The spacecraft shall not radiate a narrow-band electrical field at the C spacecraft separation plane exceeding the limit set (as the worst case of the sum spurious level) in Figure 4.6.8. 4.6.9 RF Link Interface
4.6.9.1 General RF link for S-band telemetry/command between the spacecraft on the H-IIA A launch vehicle and the user GSE can be provided. RF link between the spacecraft on-board and Masuda Tracking and Communication Station (MTCS) can be also provided (option). The interface conditions are defined as follows: (1) Link path a) From the spacecraft on-board to STA2 b) From the spacecraft on-board to ML GSE room c) From the spacecraft on-board to MTCS (2) Operation phase a) When ML is in VAB after the spacecraft VOS b) During the transfer of ML from VAB to PAD c) Before the launching after ML transfer to PAD
4.6.9.2 RF link with ML/STA2 RF link for S-band telemetry/command between the spacecraft and STA2 can be A provided after the spacecraft VOS. The following routes can be prepared to make RF coupling with the spacecraft. (a) Fairing internal antenna and RF coaxial umbilical cable (b) Air link with ML umbilical mast horn antenna via fairing RF transparent window The route above can be switched by remote-control from STA2 or ML GSE room. Available numbers and locations of fairing RF transparent windows need the discussion. A The spacecraft is connected to ML GSE room with RF through a coaxial cable on the umbilical mast and bi-directional amplifier. The users can communicate with the spacecraft by hooking up their RF GSE to the IDF in ML GSE room. When users need to communicate with the spacecraft from STA2 checkout room, fiber optics can be used between STA2 and ML (See 4.6.9.2(1),(2)).
(1) RF link when ML is in VAB or on Launch Pad (LP) Before ML is transferred from VAB or after being fixed on the LP, RF link between ML and the checkout room in STA2 is provided through fiber optics and RF/Optical Signal Converter (modulator/demodulator). The similar RF/Optical Signal Converter is installed in the checkout room in STA2. The user GSE can be connected with RF/ Optical Signal Converter in STA2. RF/Optical Signal Converter can be remote- controlled from STA2. The RF link schematic in VAB is shown in Figure 4.6.9. The RF link schematic on LP is shown in Figure 4.6.11. For dual launch, RF link can be provided for each user. The communication with the spacecraft can be performed by installing RF GSE in ML GSE room instead of using a route to STA2, however,
4-59 personnel are not allowed to stay at ML GSE room after X-5Hr (the time all personnel should leave).
(2) RF link during ML transfer During transfer of ML from VAB to LP, fiber optic network is not available. RF link during transfer is established by using air propagation between ML and VAB. ML during transfer is connected with VAB via RF Air link. RF / Optical Signal Converter is installed in VAB. STA2 and VAB are connected via fiber optics (see Figure 4.6.10). The route from STA2 should be switched on VAB side after ML transfer begins. When the communication with the spacecraft is performed by installing RF GSE in ML GSE room without using a route to STA2, personnel can stay at ML GSE room.
4.6.9.3 RF Link with MTCS A RF link in S-band between the spacecraft on-board and Masuda Tracking and Communication Station (MTCS) can be provided. The installation of user's equipment in MTCS or the use of NASDA tracking network would be optional.
(1) When ML is in VAB The spacecraft in the VAB cannot be connected directly from the MTCS via Air link. RF link between MTCS and the spacecraft shall be performed via STA2. As described in 4.6.9.2, RF link between the spacecraft and the checkout room in STA2 is provided. From the checkout room in STA2 to MTCS, the link can be provided through a STA2 inhouse coaxial cable and STA2 outside antenna for MTCS.
(2) During ML transfer The spacecraft can be linked directly from MTCS via Air link. In this case, RF transparent windows need to be installed. The numbers and locations of RF transparent windows need the discussion. A (3) After ML is transferred on LP The spacecraft can be linked directly from MTCS via Air link (as same as (2) above). Furthermore, via fiber optic network, the spacecraft on the LP can be connected with MTCS through STA2.
One faring RF transparent window is provided for each user as standard. When Air link is needed for the link between umbilical mast horn antenna and the spacecraft, the transparent window shall be located on the line connecting TT/C antenna with horn antenna. When Air link between MTCS and the spacecraft is used, the RF window shall be located on the line connecting TT/C antenna with MTCS. As umbilical mast horn antenna is seen in the completely different direction from MTCS from the view of the spacecraft, two fairing RF transparent windows (or more) may be installed. In this case, the installation of the second window or further would be optional. See § 4.4.4. for the constrains regarding the installation of RF transparent windows.
The relative clocking between the spacecraft and the fairing may affect the locations of the transparent windows. The locations of other access doors or the large doors may also be affected.
4-60 A
UM Payload fairing
S / C Junction box
Second stage Umbilical lines
ML Spacecraft umbilical STA 2 facility (connection changeable) IDF Spacecraft *1 Spacecraft A GSE Optical lines IDF IDF GSE
Figure 4.6.1 Umbilical interfaces for single launch
4-61 A
Upper UM fairing
A
Junction S / C box
Junction
box
Lower fairing S / C Umbilical lines Umbilical lines
Second stage ML Spacecraft IDF A umbilical Spacecraft STA 2 facility GSE 1 (connection IDF changeable) Spacecraft Optical lines *1 GSE IDF
IDF Spacecraft Spacecraft Optical lines *1 GSE 2 GSE IDF IDF
4-62 Table 4.6.1 Electrical interfaces
Related Item Interface access Description paragraph(s)
1) Single launch (a) Wire quantity : 120 wires 4.6.3.1
Umbilical Electrical interface 2) Dual launch interface DBAS type connector Upper spacecraft (a) Wire quantity : 120 wires 4.6.3.2 A Lower spacecraft (a) Wire quantity : 120 wires
1) Spacecraft separation command --- standard 4.6.4.1 (When user provides payload adapter) Command Same as umbilical 2) Pyrotechnic command --- option 4.6.4.1 Interface interface 3) Electrical command --- option 4.6.4.2 4) Dry loop command --- option 4.6.4.3
Telemetry 1) Separation status --- standard 4.6.5.1 ------interface 2) In-flight environmental data --- standard 4.6.5.2
4-63 H-IIA Spacecraft
Battery Battery Bus #1 Bus #2
ARM
Protective B resistor A
Pyrotechnic
Protective resistor A
Pyrotechnic
GCC2
Ground Bus
SDB2
(Reference)
Figure 4.6.3 Pyrotechnic command wiring diagram
4-64 H-IIA Spacecraft
A
Battery Bus
B
RL
GCC2
Ground Bus
SDB2
A
(Reference)
Figure 4.6.4 Electrical command wiring diagram
4-65 H-IIA Spacecraft
A
B
GCC2
Ground Bus
SDB2 (Reference)
Figure 4.6.5 Dry loop command wiring diagram
4-66 Via Payload fairing
Payload fairing
Spacecraft
UM A
Adapter
Payload support structure
Second stage
Fairing A inner wiring Adapter Via Payload adapter wiring PSS wiring
Note & : Interface connector
Payload fairing
Spacecraft
UM A
Adapter
Payload support structure
Second stage
Figure 4.6.6 Interface connectors for single launch
4-67 Upper Spacecraft UM A
Fairing inner wiring Adapter Adapter wiring
Lower fairing Lower fairing adapter adapter wiring
Lower fairing PSS wiring A
Lower Spacecraft UM A
Adapter
Payload support structure
Second stage
Example of upper spacecraft interface connectors via payload adapter B Note & : Interface connector A Figure 4.6.7(1/2) Interface connectors for dual launch (via payload adapter)
4-68 Upper Spacecraft
UM A
Fairing inner wiring Adapter Adapter wiring Lower fairing Lower fairing adapter adapter wiring
PSS wiring Lower fairing Lower Spacecraft UM
Adapter
Payload support structure
Second stage
Example of upper spacecraft interface connectors via payload fairing B Note & : Interface connector
Figure 4.6.7(2/2) Interface connectors for dual launch (via payload fairing)
4-69 Table 4.6.2 Standard interface connector specification
DBAS -O N-
Lanyard cable length (to be applied only to plugs)
Ñ A614 = 134.3 ± 2.1mm A Ñ B614 = 164.3 ± 2.1mm Ñ B864 = 188.0 ± 2.5mm Ñ C614 = 194.3 ± 2.1mm Ñ E614 = 217.0 ± 2.1mm Ñ L614 = 244.4 ± 2.1mm
P : Pin S : Socket
Shell size 3,7,12,19,27,37,61
Grounding - :No G : Yes
70 ..... Square Flange Receptacle 74 ..... Single Hole Mounting Receptacle B 78 ..... Plug with Lanyard 79 ..... Rack and Panel Plug
4-70 (V/m ) (dBμV/m) C 5.23 to 5.786 GHz 170 (171.1 dB V/m)
100 160 50 150 400 to 500 MHz 2.200 to 2.290 GHz (133.8 dB V/m) (122 dB V/m) 10 140 137 5 130 H-IIA radiation susceptibility level 2.5 14.855 to 14.865 GHz C 294 to 297 MHz 1 120 (111 dB V/m)* (114 dB V/m) 0.5 B 110 (*126 dB V/m)
0.1 100 0.05 H-IIA radiation emission level 90
0.01 80 -3 5×10 70 Electric field level 5.23 to 5.786 GHz -3 10 60 (72 dB V/m) 5 10-4 × 50
10-4 40 -5 5×10 30 -5 10 20 400 to 500 MHz -6 5×10 (18 dB V/m) 1.425 to 1.675 GHz 10 2.4 dB V/m * 10-6 0 0.01 (10 KHz) 0.1 1 10 100 1000 (1 GHz) 10000 100000
Frequency (MHz) Remarks : Option
Figure 4.6.8 Acceptable spurious radiation levels
4-71 VAB A
ML Umbilical mast horn antenna Fairing Radio UM internal transparent antenna window *1 S/C *1 Antenna S/C switch / MTCS Relay unit
S/C AMP
S/C SFA STA 2
S/C #1 S/C S/C Checkout RF Link Unit Unit ML Optical cable RF Link S/C #2 S/C NASDA RF high frequency IDF Optical cable NASDAEquipment RF / /Optical Optical Interface Interface Unit Checkout RF Link high frequency equipment Unit Unit Spacecraft GSE
Figure 4.6.9 VAB RF link schematic (for dual launch)
4-72 ML Umbilical mast upper antenna A S-band only*2
ML Umbilical mast horn antenna UM
Fairing Radio AMP internal transparent antenna window *1 SC Antenna 1 VAB * switch / SC Relay MTCS unit AMP S/C AMP IDF
Equipment / Optical S/C SFA STA 2 Interface
S/C #1 S/C Optical cable S/C Checkout RF Link Unit Unit ML
NASDA RF high frequency IDF RF Link S/C #2 S/C Equipment / Optical Interface Unit Checkout RF Link Unit Unit Spacecraft GSE
During transfer
Figure 4.6.10 RF link schematic during ML transfer (for dual launch)
4-73 ML Umbilical mast horn antenna A
Fairing Radio UM internal transparent antenna window *1 SC Antenna *1 switch / SC Relay MTCS unit
S/C
AMP
S/C SFA STA 2
S/C #1 S/C S/C Checkout RF Link Unit Unit ML
Optical cable NASDA RF high frequency RF Link S/C #2 S/C IDF Optical cable Equipment / Optical Interface Unit Checkout RF Link Unit Unit Spacecraft GSE
Figure 4.6.11 Launch pad RF link schematic (for dual launch)
4-74 4.7 Other Ground Equipment Interfaces
4.7.1 Power Several types of electrical power are available at the launch complex for spacecraft use. Commercial AC power is used for basic facility operations. Usable electrical power in the STA2 and the SFA, and after mating to the launch vehicle (on the ML) is described in APPENDIX 2 (A2.3, A2.4 and A2.6).
4.7.2 Liquids and gases All chemicals to be used will be in compliance with the requirements restricting ozone-depleting chemicals. Gaseous helium (GHe) and gaseous nitrogen (GN2) are available at STA2, SFA and VAB for spacecraft use. The gas quality is MIL-P-27407A Type 1 Grade A or equivalent for GHe, and MIL-P-27401C Type 1 Grade B or equivalent for GN2 respectively.
4.7.3 Propellant / gas sampling and analyzing A Liquids and gases provided for spacecraft use will be sampled and analyzed. Gases, such as hypergolic fuels and oxidizers, water, solvents, and hypergolic decontamination fluids can be analyzed, if necessary.
4.7.4 Filling equipment room For filling the spacecraft with propellant and pressurized gas, a filling equipment room is available either at SFA and TSA. The filling equipment room is designed for the safety of the operators when hazardous work is performed and the electrical equipment in this room is proof against hydrazine explosion to prevent secondary accidents due to leakage of the propellant from the spacecraft propulsion system. Air for a protective suit (scape-suit) is usable.
4-75 C H A P T E R 5 .
L A U N C H O P E R A T IO N S
L
A
C U N
H C
A H
P O P
T E
E
R
R A T
5 I
O
. N S CHAPTER 5. LAUNCH OPERATIONS 5.1 General
5.1.1 Scope This chapter provides users with information on typical launch operations at the launch site. The users are required to meet the requirements specified in this chapter and specified separately in the “Launch vehicle payload safety requirements (NASDA- STD-14B)” for spacecraft and in the “Ground Support Equipment (GSE) manual”, A with respect to the safety management, the safety design and the launch site operations at Tanegashima Space Center (TNSC).
5.2 Overview of the launch-related organizations To coordinate the launch services, a member of the “Office of Space Transportation Systems” shall be appointed to a Program Manager for each C spacecraft. The Program Manager shall be responsible for contracts and interface coordination (including technical items) of the launch. Figure 5.2.1 shows the NASDA headquarters’ launch operations organization and individual responsibilities. The user shall appoint a spacecraft interface manager to act as a single contact C point with NASDA for the launch service coordination. The spacecraft interface manager shall be responsible for coordination required after the launch contract is signed. NASDA shall organize a launch operations team for launch operations to be performed at TNSC. The Program Manager appointed as a single contact point and the spacecraft C interface manager shall be responsible for interface coordination on technical matters before implementation of the launch operations at TNSC before the NASDA launch operation team is established. The NASDA launch operations team, with the spacecraft interface manager, shall be responsible for interface coordination of technical matters in the launch operations after the NASDA launch operations team is established, via the NASDA Mission Director (NMD) acting as a contact point. Although the NMD shall be responsible for the technical interface coordination of the launch operations carried out at TNSC, the Program Manager finally commands this coordination. Figure 5.2.2 shows the relationship between the user and NASDA after A establishment of the NASDA launch operations team.
5-1 Spacecraft Interface Manager
Mission Operations Department Safety and Reliability Division A Safety Requirements Documents Launch Service Coordination Safety Review Interface Coordination Launch Vehicle Integration Contract Department Space Transportation System Spacecraft / Launch Operation Interface Engineering Department Spacecraft / Launch Vehicle Interface Contract for Launch Mission Analysis Flight Safety Program Manager Tanegashima Space Center Launch Operations Launch Facilities Operation Range Facilities Operation Facilities and AGE Maintenance
NASDA Launch Operations Team
Figure 5.2.1 Launch operations organization of NASDA
5-2 NASDA launch operations team
User 1 NASDA Mission Director 1 NASDA Program Manager 1
User 2 NASDA Mission Director 2 NASDA Program Manager 2
Contract coordination
Technical coordination
Figure 5.2.2 Relationship between the user and NASDA A after establishment of NASDA launch operations team
5-3 5.3 Launch Operations Requirements The launch operations requirements which specify the safety requirements, operation interface requirements and associated requirements should apply to the design and fabrication of spacecraft, and the launch operations. It should be noted, however, that the information related to the hardware and software of the launch vehicle is not included, for which the Program Manager should be C responsible. After the launch contract is signed, the documents (refer to Chap. 6.) containing the detailed description of the interface shall be prepared and shall be updated in due course with the results at the interface meetings.
5.3.1 Safety requirements The spacecraft organization shall meet the requirements specified in the “Launch vehicle payload safety requirements (NASDA-STD-14B)” (safety A requirements) separately specified by NASDA with regard to the safety management, the safety design and the launch site operation at TNSC of the spacecraft. (Refer to § 6.5.)
5.3.2 Launch operations interface requirements The launch operations interface requirements are specified in the spacecraft / H-IIA Interface Control Specifications (ICS) related to launch operations to be established separately upon agreement. It should be noted that the interface control specifications for tracking control and associated matters shall be established separately, if they are necessary. (Refer to § 6.3.)
5.4 Responsibility and Organization The NASDA launch operations team launches the H-IIA vehicle from the Yoshinobu launch complex of Osaki Range in TNSC. The spacecraft preparation shall be under the responsibility of the user. The buildings and the related facilities and GSE to be actually utilized shall be determined when the launch contract is signed.
5.4.1 Launch operations organization During the launch operations at TNSC, the user is required to appoint an Operations Manager for the spacecraft. The spacecraft organization’s Operations Manager (Spacecraft Interface Manager acting as a contact point) shall coordinate
5-4 the actual operations with the NASDA Mission Director. All coordination prior to the launch operations and coordination of matters (other than technical matters) related to the launch operations after the NASDA launch operations team is established shall be conducted by the spacecraft organization’s Operations Manager and the Program Manager. Figure 5.4.1 shows the organization chart for launch operations except for Y-2 A ~ Y-0. The countdown (preparation for lift-off) shall be supervised by the NASDA launch conductor (LCDR). The spacecraft organization must appoint the following officers and assign them to the above operations.
(1) Operations Manager for spacecraft The operations manager for spacecraft conducts all countdown operations of spacecraft, and informs the NASDA General Director of completion of prelaunch operations for the spacecraft.
(2) Spacecraft Conductor (SC) The Spacecraft Conductor shall get correct information on the progress of the spacecraft operations and issue appropriate instructions for respective operations. The Spacecraft Conductor shall notify LCDR of the spacecraft preparation progress.
The operations prior to the countdown are performed according to the spacecraft organization’s network of command. The organization chart should be submitted to NASDA before starting operations at TNSC. (When these operations are carried out by NASDA, NASDA will set up such an organization.) Figure 5.4.2 shows the organization chart for Y-2 ~ Y-0. A
5.4.2 Responsibility The operations manager for spacecraft shall be responsible for all the spacecraft operations to be performed at TNSC.
5-5 User Program Manager
Spacecraft Interface Manager
General Director
Operations Manager NASDA Mission Director Launch Director Support Director for Spacecraft Spacecraft Operations Team Planning Staff Launch Vehicle Director Flight Safety Director
Planning Group Launch Vehicle Group Flight Safety Group Logistics Support Team Mission Analysis Group Calculation Team
Range Director Range Safety Director General Affairs Director
Range Group Range Safety Group General Affairs Group Range Facilities Group Pad Safety Group Accounting Group Down Range Group Facilities Group Public Relations Group
Figure 5.4.1 Organization chart for launch operations (except for Y-2 ~ Y-0) A
5-6 General Director
NMD
Launch Director Order flow Planning & Coordination Group Check flow Logistics Support Group
Operations Manager Launch Vehicle Director Range Director Flight Safety Director Ground Safety Director for Spacecraft Range Group Launch Vehicle Group Flight Safety Group Range Safety Group Range Facilities Group Spacecraft Mission Analysis Group Pad Safety Team Down Range Group Calculation Team Operations Team
RCO FSO
SC LCDR NSO
A LSOM NQA NSAFE
SC Spacecraft Conductor LSOM Launch Site Operations Manager RCO Range Control Officer NSO NASDA Safety Officer C LCDR Launch Conductor NQA NASDA Quality Assurance Monitor FSO Flight Safety Officer NSAFE NASDA Pad Safety Officer NMD NASDA Mission Director A
Figure 5.4.2 Organization chart for Y-2 ~ Y-0 A
5-7 5.5 Restrictions
5.5.1 Restrictions on the ground In order to ensure safety, some spacecraft operations, such as accessing the spacecraft, RF radiation and switching, power interruption, and so on are prohibited in some cases. The launch date is subject to change due to weather, launch vehicle malfunction, or other reasons. Restrictions on the launch date and launch window are shown in § 5.5.2. Detailed restrictions shall be coordinated and confirmed at the interface meetings.
5.5.2 Restrictions on launching The launch date and launch window must be determined by taking many factors into account. This paragraph describes the major restrictions. Details are determined at the interface meetings.
5.5.2.1 Launch window (1) Launch period The H-IIA launch vehicle is launched during two periods, June to September and November to February. Further details are to be decided after negotiation with NASDA.
(2) Launch date The launch date, including the alternative date, shall be set in one of the two launch periods through coordination among NASDA, the users and other concerned organizations.
(3) Launch window The launch window shall be established by NASDA within the period determined by spacecraft organization analysis (including tracking control), considering restrictions such as shadow and sun angle and other relevant factors.
The launch window shall be set after the final mission analysis is completed. To maximize launch opportunity, a period of the launch window should be 45 minutes or more.
5-8 5.5.2.2 Launch postponement When the launch vehicle cannot be launched within the launch window on the scheduled date, launch shall be postponed 24 hours or more. Figure 5.5.1 shows a typical case of the operations to be performed on the launch day and the number of days of postponement. Although one day postponement is generally the case for H-IIA launch operation, it may vary depending on the necessary operations and the causes of the postponement. Depending on the causes of postponement, access to the spacecraft may take much time.
5-9 Y-1 Y-0 Y+1 Date 1820 22 246810121416182022 246810121416182022
T1 : Vehicle arming T2 : ML transfer preparation T3 : ML transfer T4 : ML / LP connection T5 : Vehicle final preparation
T6 : Facilities final preparation X-0 X-0 T7 : Terminal countdown A T8 : Facilities settling < Before vehicle arming > ~ < Before ML transfer > ~ Recycle preparation Disarming (If necessary) A ~ < Before LOX / LH2 loading > ~ VAB preparation ML / LP disconnection ML return Disarming (If necessary) A < After LOX / LH2 loading > ~
LOX/LH2 detanking Tank pressure-swing and heating AB VAB preparation ML / LP disconnection ML return Vehicle settling A
Figure 5.5.1 Typical launch postponement schedule
5-10 5.6 TNSC Facilities and GSE Related to Launch Operations of Spacecraft The following describes the major facilities and GSE which can be utilized by the spacecraft organization at TNSC. The facilities and GSE actually utilized by the user are specified in the Spacecraft / H-IIA ICS. The special facilities and GSE are operated by NASDA persons under the responsibility of the spacecraft organization.
a) Pyrotechnics storage facility and solid propellant storage facility b) Hazardous material storage (LPSA, LOSA) (propellant storage and loading) c) STA 1 and STA 2 (spacecraft functional test) d) Nondestructive Test Facility (NDTF) (solid motor X-ray inspection) e) Spacecraft and Fairing Assembly Building (SFA) (propellant loading, battery charging, encapsulating into the payload fairing, etc.) A f) Third stage and Spacecraft Assembly Building (TSA) (propellant loading) g) Solid Booster Test Building (SBB) h) Vehicle Assembly Building (VAB) (mating to launch vehicle, battery charging, final checkout, arming, etc.) i) Movable Launcher (ML) B j) Takesaki Range Control Center (RCC) k) Other specifically requested facilities and GSE
Figure 5.6.1 shows the location of spacecraft-related buildings in TNSC’s Osaki Launch Range.
5-11 A Non-Destructive Test Facility
N No.2 Spacecraft Test to Hirota and Assembly Building
Spacecraft and Fairing Spin Test Building Assembly Building Osaki Power Station Spin Liquid Propellant Measument storage Building Liquid Oxydizer Osaki 5 roads storage crossing gate Third Stage and Satellite Solid Booster Building Yoshinobu Vehicle Assembly Building Assembly Building
No.1 Spacecraft Test Osaki No.1 Support Yoshinobu Block House and Assembly Building LSB gate Garage Osaki Supporting Liquid Hydrogen Storage Reserroir Launch Building Propellant New Yoshinobu Block House (underground) Osaki Restaurant Storage Facility Yoshinobu Second Camera Room Osaki office 2 Osaki office 1 Yoshinobu (H-IIA) Launch Complex LE-7 Engine Launch 0 100 200 300 400 500 Building Tool Room Test Stand Pad Service Tower (Launch Pad)
New Launch Pad 80m Meteorological Observation Tower Liquid Oxygen Storage
High-pressure Sea shore gate Gas Storage
to Kukinaga, Kaminaka
Pacific Ocean
to Takesaki Administration Building
Figure 5.6.1 Location of spacecraft-related buildings in TNSC’s Osaki Launch Range
5-12 5.7 Launch Operations This section describes the typical spacecraft launch operations performed at TNSC.
5.7.1 Spacecraft-related operations and programs The spacecraft launch operations at TNSC can be broadly classified into three phases:
Phase 1 Spacecraft preparation and functional test Phase 2 Spacecraft hazardous operations Phase 3 Joint operations by spacecraft and launch vehicle organizations
Phase 1 operations are independent without any direct interface with the launch vehicle system. They are performed in STA2 and / or STA1.
Phase 2 operations are independent without any direct interface with the launch vehicle system. They consist of hazardous operations and are performed in the SFA, the TSA, the SBB, the NDTF, and others.
Phase 3 operations are performed jointly by the launch vehicle and spacecraft organization, and are performed in the SFA and Yoshinobu Vehicle Assembly Building (VAB).
Generally in the initial planning of spacecraft launch operations, the spacecraft organization is expected to perform all operations from delivery of the GSE (checkout equipment) to TNSC to its removal from TNSC within 45 days (working days) (40 days before the launch, and five days after the launch). The spacecraft / fairing assembly shall generally be mated to the launch vehicle (spacecraft VOS) four days before launch (Y-4). Figure 5.7.1 shows a typical spacecraft launch operations schedule. Figure 5.7.2 depicts a typical operations flow diagram for spacecraft preparation at TNSC.
5.7.2 Phase 1 (spacecraft preparation and functional test) After arrival at Tanegashima Airport, Nishino-omote or Shimama Seaport, the spacecraft and GSE are transported to TNSC on the public roads. (See Appendix 2.) The spacecraft delivered by the spacecraft organization to the STA 1 or STA 2 (specified in advance), are unpacked and installed by the spacecraft organization.
5-13 NASDA will support these tasks. Hazardous materials, such as solid motors, pyrotechnics, propellant, and explosives, are delivered by the spacecraft organization to the specified place where they are to be stored by NASDA. Consumable materials required for the operations shall be prepared by the spacecraft organization. Propellant and high-pressure gas may be provided by NASDA with extra charge. The spacecraft is transported to the test room (clean room), and GSE is set up in the checkout room adjacent to the test room. The spacecraft assembly and functional tests are performed in the STA 1 or STA 2. Hazardous operations, such as installing pyrotechnics and loading propellant, must not be performed in these facilities. However, inert high- pressure gas systems can be charged if sufficient safety is ensured (this operation requires permission of the NASDA launch site safety division). After assembly and functional tests, the spacecraft is loaded by the spacecraft organization into the container for transportation prepared by the spacecraft organization or by NASDA and transported to the SFA or TSA by the spacecraft A organization or by NASDA (option). Figure 5.7.3 depicts a typical Phase 1 operations flow diagram.
5.7.3 Phase 2 (hazardous operations for spacecraft) Hazardous operations such as installing solid motors and pyrotechnics, loading propellants, and charging high-pressure gas systems shall be performed in the SFA or TSA. Hazardous operations shall be performed by the minimum required number (but more than two) of persons who have received safety instruction and training. Other persons give operation instructions or monitor the status from the monitor room. Only explosion-proof GSE shall be set up in the room where the hazardous operations are to be performed. Operations in Phases 1 and 2 may be performed in parallel. Figure 5.7.4 depicts a typical Phase 2 operations flow diagram.
5.7.3.1 Preparing and assembling pyrotechnics and solid motor Before installation on the spacecraft, the pyrotechnics and the solid motor shall be inspected in the NDTF or SBB (pyrotechnics only). After inspection and assembly, the pyrotechnics are transported into the SFA or TSA. All these operations, including transportation between buildings, shall be under the responsibility of the spacecraft organization. NASDA shall provide support for handling X-ray test equipment of the NDTF, movement of materials, and similar operations requiring use of the NASDA equipment and materials. After the pyrotechnics have been inspected or installed, the pyrotechnics and
5-14 the solid motor must be stored in their respective storeroom.
5.7.3.2 Spacecraft operations (1) Transfer The spacecraft in the container is transferred by the spacecraft organization or A by NASDA (option) from the STA 1 or STA 2 to the SFA or TSA, and is unloaded from the dolly in the air lock entrance room. After the cleanliness in the air lock room has been confirmed, the spacecraft is taken out of the container and is moved into the assembly room (clean room) by the spacecraft organization.
(2) Loading the propellant and charging the high-pressure gas system The propellant and the high-pressure gas for pressurization up to the flight level are loaded in the SFA or TSA (or in the VAB for special cases) by the spacecraft organization. Also the operations such as depressurization, purging and flushing should be under the responsibility of the spacecraft organization. Batteries can be charged in the SFA and / or TSA, if hazardous operations are permitted.
(3) Installing the pyrotechnics The pyrotechnics and the solid motor shall be installed by the spacecraft organization in the SFA and / or TSA, if special permission is obtained from NASDA. However, pyrotechnics wire connection and arming shall be conducted by the spacecraft organization in the VAB as part of the countdown operations.
5.7.3.3 Final spacecraft assembly (1) Weight measurement The spacecraft can be weighed in the STA, SFA, or TSA under the responsibility of the spacecraft organization. NASDA has the right to request the spacecraft organization to weigh the spacecraft, concerning its effect on flight performance. Whether the weight is to be measured before or after loading the propellant and charging the high-pressure gas systems will be determined by coordination with the spacecraft organization. The spacecraft organization can utilize the weighing equipment of NASDA.
(2) Final inspection Electrical and mechanical inspection and solid propellant motor arming inspection must be completed before spacecraft encapsulation into the payload fairing. After entering Phase 3 operations, direct access to the launch vehicle body (including the spacecraft) is allowed only when it is approved in the interface meeting; operations by telecommunications signals via the umbilical line or RF signals are allowed.
5-15 5.7.4 Phase 3 (joint operations by spacecraft and launch vehicle organizations) Phase 3 consists of joint operations by spacecraft and launch vehicle organizations. Operations from mating the spacecraft and the payload support structure to mating the encapsulated spacecraft on the launch vehicle are conducted by the launch vehicle organization; the spacecraft organization provides support and monitoring. Operations after mating are performed by both spacecraft and launch vehicle organizations under various restrictions. Figure 5.7.5 depicts a typical Phase 3 operations flow diagram for a single launch. Figure 5.7.6 depicts a typical Phase 3 operations flow diagram for a dual launch. In case of typical operations, the payload adapter is mated to the spacecraft prior to Phase 3, unless there are any restrictions for the spacecraft.
5.7.4.1 Encapsulation into the payload fairing The first operations in Phase 3 (joint operations by spacecraft and launch vehicle organizations) are to mate the spacecraft and payload support structure and to encapsulate them into the payload fairing in the SFA. These operations are performed by the launch vehicle organization with the support of the spacecraft organization. The spacecraft is usually handled together with payload adapter by means of the handling jig prepared by the spacecraft organization. The additional weight of the adapter must be considered in the spacecraft and spacecraft handling jig design.
5.7.4.2 Encapsulation for single launch After the checkout of the spacecraft in the SFA, the spacecraft with the payload adapter is encapsulated into the payload fairing. The major operations here are as follows:
(1) The spacecraft with the payload adapter (PLA) is mated to the payload support structure (PSS) and then the ordnance for the spacecraft separation is mounted. (The PLA can be usually mated to the spacecraft in any phase prior to Phase 3. If the spacecraft does not accept this sequence, however, the PLA is mated to the PSS prior to mating with the spacecraft and then the spacecraft is mated on the top of the PLA in this phase.)
5-16 (2) The payload fairing encapsulates the spacecraft from the top and then the C fairing bottom flange and forward connection flange of the PSS are bolted to each other. (In case of the model 5S fairing, each half-shell is mated in parallel with the spacecraft.) Figure 5.7.7 depicts a typical encapsulation sequence for single launch.
5.7.4.3 Encapsulation for dual launch After the checkout of the spacecraft in the SFA, two spacecraft are encapsulated into the payload fairing separately. The major operations here are as follows:
(1) The lower spacecraft with the PLA is mated on a corresponding PSS and then the ordnance for the spacecraft separation is mounted. (The PLA can be usually mated to the spacecraft in any phase prior to Phase 3. But if the spacecraft does not accept this sequence, the same operations are conducted as a single launch.)
(2) The upper spacecraft with the PLA is mounted with the ordnance for the spacecraft separation in parallel with the lower spacecraft.
(3) The lower payload fairing for a dual launch (hereafter referred to as the lower fairing) encapsulates from top of the lower spacecraft and then the bottom flange of the lower fairing is bolted to the connection flange of the lower PSS.
(4) The upper spacecraft with the PLA is mounted on the upper PSS which is connected to the forward end of the lower fairing.
(5) The upper fairing encapsulates from top of the upper spacecraft and then the upper fairing is bolted to the upper PSS. Figure 5.7.8 depicts a typical encapsulation sequence for a dual launch.
5.7.4.4 Transportation of encapsulated spacecraft The encapsulated spacecraft is mounted on the dolly in the SFA air lock room (1) and is transported to the VAB by tractor. It is separated from the dolly on the lower floor of the VAB, and is hoisted. Figure 5.7.9 depicts a typical transportation sequence of the encapsulated spacecraft. During transportation, NASDA will monitor the conditions to meet the following requirements. Temperature : 5 ºC to 30 ºC Humidity : less than 60 % RH
5-17 Vibration : less than 0.6 G O-P (for each axis) G = 9.80665 (m/s2) B If required air conditioning will be prepared by NASDA (option). A
5.7.4.5 Mating with launch vehicle After the encapsulated spacecraft is lifted on the upper floor of the VAB, it is directly mated on the top end of the forward skirt of the second stage. After that, the air conditioning duct is connected to the fairing, and the payload fairing A assembly transport jig is removed and then the payload fairing opening spring, quick disconnector (QD), and fairing separation pyrotechnics are attached to their positions. Figure 5.7.10 depicts a typical sequence of the mating the encapsulated spacecraft to the launch vehicle.
5.7.4.6 Spacecraft inspection after installation The spacecraft functional test can be conducted according to the joint operations schedule determined in advance. (Detailed coordination is to be made with the NASDA mission director.) This also applies to the RF link test, leakage inspection, battery charging, and visual inspection. Spacecraft arming and disarming operations shall be inspected and witnessed by the NASDA launch site safety division. During these operations, turning on the electric system or RF radiation system is prohibited in the launch vehicle system . The electric or RF radiation systems of the spacecraft are prohibited to operate during the installation of pyrotechnics of the launch vehicle. While some propulsion system operations are being performed, switching may be prohibited. Details shall be determined during the joint operation schedule coordination with the NASDA launch site safety division.
5.7.4.7 Y-3 operation A The final preparation for the countdown configuration is performed during precountdown operations. Details shall be determined at the interface meetings and the coordination meeting before countdown.
5.7.4.8 Y-2 ~ Y-0 operation A The spacecraft organization shall perform the final functional test using the ground line and RF, and charge the battery within the specified time. Details shall be determined at the interface meetings and the coordination meeting
5-18 before countdown. Figure 5.7.11 shows a typical countdown schedule. The following gives some parts of restricted operations during the countdown.
(1) Hazardous operations a) Leakage check of the high pressure bottle of the launch vehicle (Y-3) A During this operation, access to the spacecraft is prohibited. (VAB entrance is controlled.)
b) Loading propellant for the second stage gas jet system of the launch vehicle (Y-2) During this operation, all of spacecraft operations are prohibited. (VAB C entrance is controlled.) A
c) Pyrotechnics wire connection by the launch vehicle organization (Y-2) During this operation, spacecraft operations such as switching or turning on the electric and RF radiation system are prohibited.
d) Vehicle body arming by the launch vehicle organization (Y-1) During this operation, spacecraft operations such as switching or turning on the electric and RF radiation system are prohibited.
e) Pyrotechnics wire connection and removing the SAD safety pin by the spacecraft organization (Y-0) During these operations, launch vehicle operations such as switching or turning on the electric and RF radiation system are prohibited.
(2) Others a) ML transfer (Y-0) During and after this operation, access to the spacecraft is prohibited.
In case of a dual spacecraft launch, all restrictions from each spacecraft side will be imposed on each other. The problems will be coordinated at the interface meetings and the coordination meeting before countdown.
5.7.4.9 Terminal countdown Figure 5.7.12 shows a typical terminal countdown sequence of the launch vehicle system on the launch date (Y-0). During terminal countdown, the spacecraft organization shall perform the following operations:
5-19 (1) Spacecraft RF flight configuration The final spacecraft RF flight configuration shall be completed before X-10 (T.B.D.) minutes. No change is allowed until 20 seconds after the separation of the spacecraft, unless the launch vehicle organization agrees.
(2) Final spacecraft inspection and Switching the spacecraft power supply The final spacecraft inspection and switching the spacecraft power supply from an external to infernal source shall be completed before the completion of the spacecraft preparation.
(3) Solid propellant apogee motor arming (when required) The spacecraft solid propellant apogee motor shall be armed before X-10 minutes. The arming timing requires agreement with the launch vehicle organization.
(4) Completion of spacecraft preparation The spacecraft organization shall complete spacecraft preparation incorporated in the launch vehicle sequence before X-270 seconds. B
Completion of the spacecraft preparation is a prerequisite for the automatic countdown sequence.
(5) Automatic countdown sequence The launch vehicle organization starts the automatic countdown sequence after X-270 seconds. B
(6) Countdown recycle If the recycle command is issued during the countdown, all countdown operations are reset to X-25 minutes. B
5.7.4.10 Recycle operations (Launch postponement) If launching is postponed on the launch date, recycle operations are performed. The major recycle operations are as follows: a) Discharging the launch vehicle propellant and purging b) Transporting the ML to the VAB A c) Disarming spacecraft systems (if required) d) Disarming launch vehicle systems (if required) e) Operations required for repetition of Y-0
5-20 Launch date
Working day Y-50 45 40 35 30 25 20 19 18 17 16 15 14 13 12 11 10 9 8 7 6 5 4 3 2 1 0 +1 +5
Workplace Operations
STA 1 Spacecraft preparation or Functional test
Phase 1 STA 2 Battery charging & discharging
SBB Pyrotechnics preparation
NDTF Solid motor X-ray inspection
Propellant loading & high-pressure gas charging Phase 2 Pyrotechnics & solid motor SFA installing Spacecraft final assembling TSA Payload adapter connection Encapsulation into the fairing Preparation to transport to VAB
Transportation to VAB Encapsulated spacecraft Phase 3 VAB mating to the launch vehicle Spacecraft test after mating STA2 Countdown A
Post-lift-off operation
Figure 5.7.1 Typical spacecraft launch operations schedule
5-21 Tanegashima STA 2 SFA VAB (ML) Airport Spacecraft preparation Pyrotechnics installation Functional tests Functional check SPM installation RF compatiblity tests RF link tests to tracking station Propellant loading Spacecraft rehearsal Battery charging Mating with adapter Spacecraft arming Spacecraft weighing Battery charging Nishinoomote Pressurization Final countdown Port Battery charging Encapsulation in fairing
STA2 STA 1 TSA Shimama Remote control Port Same as STA 2 only for Pyrotechnics installation small S/C SPM installation Propellant loading Spacecraft weighing Pressurization Battery charging
Solid propellant storage NDTF SPM storage X-ray inspection
SBB Pyrotechnics storage Pyrotechnics preparation Pyrotechnics storage SPM preparation
Figure 5.7.2 Typical operations flow diagram for spacecraft preparation at TNSC
5-22 Spacecraft Unloading Unpacking Preparation Functional tests Preparation Tanegashima for transport Airport Ground support equipment
Unpacking Validation @ STA 2 or @ STA1 Unloading Transportation Nishinoomote By road A Port Propellant loading & pressurizing system
Unpacking Validation @ SFA or @ TSA Unloading Shimama Solid propellant motor Port Unpacking Storage @ Solid propellant storage
Pyrotechnics
Unpacking Storage @ Pyrotechnics storage
Figure 5.7.3 Typical phase 1 operations flow diagram
5-23 Spacecraft Battery charging
STA 2 Installing Weighing & or Installing pyrotechnics Propellant loading Pressurizing SPM Final check STA 1
Preparation for GSE Preparation Validation
Pyrotechnics to Phase 3 Storage Preparation SBB
Solid propellant motor (SPM)
X-ray inspection Storage NDTF
Figure 5.7.4 Typical phase 2 operations flow diagram
5-24 A Y - 7 Y - 6 Y - 5 Y - 4 Y - 3 ~ 1 Y - 0
Mating Encapsulation Preparation Transportation Spacecraft test Countdown with payload in fairing to transport to VAB after mating support structure to VAB ML transfer Hoisting of Preparation encapsulated of encapsulation spacecraft
Mating with launch vehicle
All system (spacecraft / vehicle) rehearsal (dry run)
@ SFA @ VAB VAB LP
Figure 5.7.5 Typical phase 3 operations flow diagram for single launch (for 4S fairing)
5-25 Y - 10 Y - 9 Y - 8 Y - 7 Y - 6 Y - 5
Preparation of Upper spacecraft Encapsulation in Preparation to Upper connection to connects to lower Upper fairing transport to VAB spacecraft lower fairing fairing lid
Mating Encapsulation in Preparation of with payload lower fairing encapsulation Lower support structure spacecraft Preparation of encapsulation
@ SFA
Y - 4 Y - 3 ~ 1 Y - 0 A
Transportation Spacecraft test Countdown to VAB after mating ML transfer Hoisting of encapsulated spacecraft
Mating with launch vehicle All system (spacecraft / vehicle) rehearsal (dry run)
@ VAB VAB LP
Figure 5.7.6 Typical phase 3 operations flow diagram for dual launch
5-26 Working platform
Sling
Payload Spacecraft support structure Payload adapter (PSS) (PLA) Base ring Base (jig) (jig)
Lifter
Mating preparation Mating spacecraft Spacecraft checkout to PSS before encapsulation
Payload fairing
Spacecraft encapsulation Encapsulated spacecraft Sealing and preparation for transportation
Figure 5.7.7 Typical encapsulation sequence for single launch (for 4S fairing)
5-27 Upper fairing Upper spacecraft
Payload adapter Dolly (PLA) A
PLA PSS
Lower fairing
Lower Payload spacecraft support PLA structure (PSS) Mating upper Upper spacecraft Sealing and preparation spacecraft encapsulation for transportation Base (jig) Base ring (jig) to PSS in upper fairing
Lifter
Mating lower spacecraft Lower spacecraft encapsulation to PSS in lower fairing
Figure 5.7.8 Installation sequence for dual launch
5-28 Working platform
Tractor Base (jig) Base ring (jig) Dolly Lifter
Gas purge unit
Preparation for Installation on the dolly Transportation transportation (SFA VAB)
VAB
A
Grounding wire
Preparation for hoisting Hoisting
Figure 5.7.9 Transportation sequence of the encapsulation spacecraft
5-29 Movable floor
Second stage Mating fairing to Access to launch vehicle the second stage
Handling ring
Air conditioning Air conditioning duct duct
Connecting air conditioning duct to the fairing Removing handling ring
Figure 5.7.10 Encapsulated spacecraft and launch vehicle mating
5-30 YÐ3 Y-2 Y-1 Y-0 Date 246810121416182022246810121416182022246810121416182022246810121416182022
Access limit S/C access allowed time AB
Battery charging and final configuration set up
Prop-system valve check
Pyrotechnics circuit check
Pyrotechnics connection
Second stage gas jet propellant loading andloading unit carrying out
Guidance and Control system, RF system check A Final closure and preparation for transfer
Mechanical / umbilical system final preparation
Launch vehicle arming
ML transfer
Vehicle final preparation X-0 Terminal countdown A
Figure 5.7.11 Typical countdown schedule
5-31 Real-time (min) —480—450—420—390—360 —300—270—240 —180 —120 —60 0 +30 +90
X time (min) —420—390—360 —300 —240—210—180 —110 —40 —40 0 Launch / Vehicle System operations B Terminal Countdown preparation
Personel arrangement and briefing
RF system check Hold X-270 s Automatic C/D (Official 60 min.) sequence start Terminal countdown
Terminal countdown settling / recycling
Propulsion system operations
Operation Items Air conditioning Switching Mode III → I Air conditioning stop
First / second stage LOX / LH2 Final preparation Chilling down and Loading propellant loading (98% Additional loading)
Ambient / Cryogenic bottle pressurizing (X-18 s) Guidance system operations Initial alignment Flight mode ON A *1
(*1 : Perform befor 60 min. Real time)
Figure 5.7.12 Typical launch vehicle system countdown schedule
5-32 C H A P T E R 6 .
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N T CHAPTER 6. INTERFACE MANAGEMENT
6.1 General This chapter describes the spacecraft / H-IIA launch vehicle interface management.
6.1.1 Launch service organization The NASDA organization for launch services is described below.
(1) Technical coordination section Office of Space Transportation Systems a) Mission Operations Department (MOD) A The MOD is responsible for general coordination related with launch services as a coordination contact point with the spacecraft organization. In the concrete, the MOD is responsible for the interface between the spacecraft / the H-IIA launch A vehicle and between the spacecraft / launch operations. Furthermore, the MOD is responsible for procurement of the H-IIA launch A vehicle and also spacecraft safety reviews. b) Space Transportation System Engineering Department This section takes charge of mission analyses such as a trajectory analysis, a CLA and so on, and mission modifications of the H-IIA launch vehicle based on each spacecraft’s requirements.
(2) Contract section Contract Department This section is responsible for a contract related with launch services.
(3) International coordination section External Relations Department This section is the contact point in case that an agreement or coordination with the government organization is necessary.
6.1.2 Interface management document
6.1.2.1 Documents to be submitted by the spacecraft organization The spacecraft organization shall submit to NASDA the following documents which are required for coordinating the interface between the spacecraft and H- IIA launch vehicle system.
6-1 Item When to submit
a. Documents and data required for mission analyses Refer to ¤ 6.2.1 (mission requirements, mass properties, and coupled loads analysis model and integrated A thermal analysis model of the spacecraft b. Schedule of the spacecraft Refer to ¤ 6.5.1 and 6.5.4 (development, test, launch operation, etc.) c. Safety program plan Refer to ¤ 6.6 d. Safety data package Refer to ¤ 6.6 e. Documents required for legal procedures When NASDA requires
6.1.2.2 Document to be submitted by NASDA After concluding the launch contract, NASDA and the spacecraft organization have an interface meeting, and the results of the agreement between the concerned organizations are incorporated in the “Spacecraft / H-IIA interface control specifications” (hereinafter referred to as “ICS”) which specify details of the interface items. This ICS is maintained and managed by NASDA until launch.
6.2 Interface Work with Spacecraft Organization Spacecraft / H-IIA launch vehicle interface work is handled by the NASDA MOD. A The results of the interface coordination are incorporated in the ICS, whenever applicable.
6.2.1 Interface schedule / Interface items The interface schedule is established to define interface tasks including analysis, fit check and review meetings. It specifies the time of the tasks implementation and data exchange. The details will be coordinated at interface meetings for each mission.
6.2.1.1 Standard mission Interface schedule and items of the standard mission in which the H-IIA launch vehicle will launch a spacecraft to a standard orbit (such as a geostationary transfer orbit and a polar orbit) are shown in Figure 6.2.1.
6-2 6.2.2 Mission analysis The following analyses are conducted by NASDA for the spacecraft organization. However, if results of analyses on other similar mission can be adopted, these results are substituted for results of the following analyses. a) Trajectory analysis b) Orbit dispersion analysis c) Sun angle analysis d) Spacecraft separation analysis e) Relative trajectory analysis of spacecraft and launch vehicle f) Spacecraft coupled loads analysis g) Radio compatibility study h) Integrated thermal analysis A
6.2.2.1 Trajectory analysis Using the mission requirements for the spacecraft organization insertion trajectory (mass properties, sun angle restrictions, requirements at the time of separation, etc.) and launch vehicle data, the trajectory is calculated to meet the mission requirements while satisfying various restrictions, and the major sequence of events and the trajectory parameters at the time of spacecraft separation are supplied to the spacecraft organization.
6.2.2.2 Orbit dispersion analysis This analysis uses the trajectory worked out in a) to estimate the orbit error at the time of orbit injection caused by error sources including the vehicle body and inertial sensors, and to provide the covariance matrix at the time of spacecraft separation.
6.2.2.3 Sun angle analysis This analysis uses the trajectory worked out in a) to provide information on the temporal change of the sun angle and eclipse.
6.2.2.4 Spacecraft separation analysis This analysis provides the attitude angle error and attitude rate at the time of C separation by analyzing separation motion of the second stage and spacecraft, based on the most current information on spacecraft mass properties obtained from the spacecraft organization.
6-3 6.2.2.5 Relative orbit analysis of vehicle and spacecraft This analysis provides the relative orbit and the relative position between the vehicle and the spacecraft after spacecraft separation. Further this analysis supplies information on pressure and thermal impacts and amount of contamination deposit to which the spacecraft is exposed by the vehicle during the collision avoidance maneuver after the spacecraft separation. A
6.2.2.6 Spacecraft coupled loads analysis This analysis uses the most current spacecraft dynamic model supplied from the spacecraft organization to obtain the load imposed on the spacecraft and the relative displacement from the payload fairing during launch in combination with the vehicle body model. The analysis result will be supplied to the spacecraft organization. Coordination will be made with the vehicle organization as regards the timing for spacecraft coupled load analysis.
6.2.2.7 Radio frequency compatibility study RF compatibility is checked based on the H-IIA launch vehicle / spacecraft spurious radiation acceptable levels designated in Figure 4.6.8 and actual data of the spacecraft EMC test.
6.2.2.8 Integrated thermal analysis A Integrated thermal analysis will be conducted using a thermal model provided by the user if required. This analysis covers the period after mating to the launch vehicle in the VAB up-to the injection into the orbit.
6.2.3 Interface test
6.2.3.1 Fit check of spacecraft and PLA The fit check of the spacecraft and the payload adapter is conducted by NASDA using the flight model payload adapter or an equivalent model. The payload adapter or test jigs are prepared and transported by NASDA. Time of fit check is given in Figure 6.2.1. Details will be established at the interface meeting. When the equivalent model is used for this check, the fit check using the flight model must be conducted at the launch site prior to launch operations.
6-4 6.2.3.2 Umbilical connector disconnection test A disconnection function between the spacecraft umbilical connector and the payload fairing umbilical connector is checked by the following two tests to verify that normal connection and disconnection are possible. These tests are conducted by NASDA. The test timing is illustrated in Figure 6.2.1. a) Preliminary disconnection test b) Launch site disconnection test
(1) Preliminary disconnection test The preliminary disconnection test is conducted using either of the following two methods:
a) A single connector of the payload fairing (flight model or its equivalent) is transported to the spacecraft, and is subjected to the disconnection test. b) A single connector of the spacecraft (flight model or its equivalent) is transported to the payload fairing fabrication factory, and is subjected to the disconnection test.
(2) Launch site disconnection test The umbilical connector disconnection test is conducted after the spacecraft and payload fairing are mated to the second stage of the vehicle at the launch site. This test is applied for only lanyard type connector . A
6.2.3.3 Separation shock test The separation shock test at the spacecraft separation event is conducted by users using the flight model payload adapter or an equivalent test model. This test is usually held together with the spacecraft / payload adapter fit check. Time schedule of the shock test is given in Figure 6.2.1. Details will be established at the interface meeting.
6.2.4 Mission modification
6.2.4.1 Payload fairing mission modification The following items can be mounted on the payload fairing as requested by the spacecraft organization. The standardized numbers of respective items as described hereinafter are summarized in Table 6.2.1.
6-5 Details are described in § 4.4.4.
a) Access doors b) Large doors c) Umbilical connectors d) Radio transparent window e) Internal antenna f) External antenna g) Internal and external antennas h) Acoustic blankets i) Independent air conditioner (air conditioned inlet, partition wall and relief valve)
Figure 6.2.1 illustrates the interface schedule for this mission modification.
6.2.4.2 Payload adapter mission modification The following items can be mounted on the payload adapter as requested by the spacecraft organization. a) Separation springs b) Umbilical connectors
6.2.5 Standard services and optional services Standard services and optional services to launch the spacecraft are provided by the NASDA.
6.2.5.1 Distinction between standard and optional service items Table 6.2.1 shows a distinction between standard service items and optional service items in the WBS (working breakdown structure) of the mission integration.
6.2.5.2 Notes a) The payload adapter selected by NASDA is a standard payload adapter. If the spacecraft organization will use a new payload adapter (PLA), the spacecraft organization is responsible for development and costs of the PLA.
b) The spacecraft organization can utilize the NASDA’s PLA for the spacecraft development. But the spacecraft organization is responsible for costs of preparation of the PLA.
6-6 c) Standard mission modifications of the mission analysis and the payload fairing are shown in Table 6.2.1. In case that there are additional mission modification items as option, NASDA requires additional costs from the spacecraft organization. d) Configurations and specifications of spacecraft related facilities and Ground Support Equipment (GSE) of NASDA are shown in the concerned chapter of this manual and H-IIA Payload-Related Facilities and GSE Manual (in preparation).
The spacecraft organization is responsible for costs if modifications of these facilities for the spacecraft are required.
6-7 year -4 -3 -2 -1 +1 month -46 -44 -42 -40 -38 -36 -34 -32 -30 -28 -26 -24 -22 -20 -18 -16 -14 -12 -10 -8 -6 -4 -2 0 2 4 68 10
Spacecraft system milestone (REF) PDR CDR PQR PSR
Interface work milestone ICS Draft ICS Revisionl ICS Final MRR / PSR
A Mission requirements B Injection orbit A B A B Flight Trajectory Analysis parameters
C Covariance matrix C Orbit Dispersion Analysis
D Sun angle histories D Sun Angle Analysis
E Spacecraft mass properties E Inflight Mass Property Analysis
G Atitude rate, G S/C separation dynamics analysis F Characteristics F
H S/C model I Dynamic load H I S/C H-IIA Coupled Loads Analysis H-IIA S/C *1 *1
Reference Thermal Coupled Analysis (if required) J S/C model K Thermal condition J K Detailed *1 *1 Final Radio Compatibility Study ( If necessary )
Payload adapter F/C Fit check S/T Shock test
Mission modification M/M Payload fairing U/P Umbilical connector pulling test
*1 : This item is planned on the spacecraft requirement.
Figure 6.2.1 Typical spacecraft / H-IIA launch vehicle interface schedule for standard mission
6-8 Table 6.2.1(1/6) Mission Integration WBS A
Responsibility ※1 WBS No. Items Description Remarks ※2 Cost Work
overall 01 -01 interface meeting SC and LV SC and LV standard
interface management document
-01 ICS LV (SC) LV (SC) standard B 02 -02 operational plan SC SC SC work
-03 range safety plan LV (SC) LV (SC) standard
flight analysis C
trajectory analysis
-01 SCmass properties before flight plan,final trajectory SC SC SC work B
(preliminary trajectory analysis [for new mission] ) LV LV standard flight plan,final trajectory
-02 orbit dispersion analysis flight plan LV LV standard
(preliminary trajectory -03 sun angle analysis LV LV standard [for new mission] and) flight plan
(preliminary trajectory -04 SC separation analysis LV LV standard [for new mission] and) flight plan 03 (preliminary trajectory -05 relative orbit analysis LV LV standard [for new mission] and) flight plan
coupled loads analysis -06 analysis prior to SC test (normally twice) LV LV standard
-07 radio compatibility study prior to SC test SC and LV SC and LV standard
integrated thermal analysis A
-08 SC mathematical thermal model SC SC SC work
analysis prior to SC test (normally once) LV LV option
-09 dual launch compatibility analysis LV LV standard
SC : Spacecraft LV : Launch Vehicle
6-9 Table 6.2.1(2/6) Mission Integration WBS A
Responsibility ※1 WBS No. Items Description Remarks ※2 Cost Work
vehicle manufacture
-01 standard body (H2A202) addition for other version LV LV standard 04 -02 fairing (4s) addition for other version LV LV standard
in case of 937M spin, -03 PLA LV LV standard 1194M, 2380S, 3470S B
PLA
-01 new PLA development except for WBS04-03 SC LV option
mission modification 05 number of separation springs standard:4 to 8 LV LV standard -02 umbilical connector on PLA to be prepared by LV or SC LV SC or LV standard
number of separation switches standard : 2 LV LV standard
fairing
mission modification
number of access doors standard:4(2forlower fairing) LV LV standard
number of large doors 4 for vehicle operation in principle LV LV standard A
number of umbilical connectors selective PLA / PLF type LV LV standard
RF link method 06 number of internal antennas standard -01 (standard : 1) LV LV
number of radio transparent windows standard (standard : 1) LV LV C number of internal / external option antennas (standard : none) SC LV
other methods (standard : none) SC LV option
acoustic insulation blankets 10 mm LV LV standard
air conditioner LV LV standard
SC : Spacecraft LV : Launch Vehicle
6-10 Table 6.2.1(3/6) Mission Integration WBS A
Responsibility ※1 WBS No. Items Description Remarks ※2 Cost Work
interface verification test
SM and PFM s/c test plan -01 test for launch vehicle LV LV standard
mechanical compatibility test / PLA for test to be prepared by static loads ¥ vibration test SC in principle PLA will be re-fablicated PLA preparation and transportation SC LV option or newly developed (standard : no separation spring and no connector) -02 AGE and jigs preparation SC-PLA mating portable stand SC SC SC work
mating operation SC LV (SC) option
test SC SC SC work
separation shock test
PLA preparation / transportation PLA will be re-fablicated SC LV option or newly developed
pyrotechnics preparation SC SC or LV SC work or option -03 07 AGE and jigs preparation SC-PLA mating portable stand SC SC SC work
mating operation SC LV (SC) option (SC work)
test SC SC SC work
fit check of flight component standard joint work of LV and SC
PLA preparation / transportation LV LV standard -04 AGE and jigs preparation SC-PLA mating portable stand SC LV option
fit check LV and SC LV and SC standard
disconnect functional test of umbilical connectors
umbilical connector preparation to be prepared by LV or SC SC or LV SC or LV SC work or standard -05 AGE and jigs preparation SC-PLA mating portable stand SC LV option
test LV and SC LV and SC standard
test result of STM and PFM -06 for LV LV LV standard
SC : Spacecraft LV : Launch Vehicle
6-11 Table 6.2.1(4/6) Mission Integration WBS A
Responsibility ※1 WBS No. Items Description Remarks ※2 Cost Work
safety reviews
08 -01 SAFETY SUBMISSION in every phase SC SC SC work
-02 reviews LV LV standard
support for legal procedure if necessary for foreign SC 09 law for high-pressure gas, radio, -01 SC LV option pyrotechnics, etc. in response to requirements not 10 modification of facilities specified in user's manual SC LV option standard working days : within 45 launch operations special consideration for over 45 days
interface document
-01 SC launch operation plan SC SC
joint operation plan LV (SC) LV (SC)
phase 0 preparation for SC operation
facilities validation -02 test except GSE (validation cable) LV LV standard
test including GSE SC SC SC work
phase 1 SC independent work
fitting pyrotechnics SC SC SC work
11 -03 visual inspection SC SC SC work
final assembly of SC SC SC SC work
inspection after final assembly SC SC SC work
phase 2 SC independent hazardous work
fitting pyrotechnics SC SC SC work to be prepared by SC side preparation of loading equipment in principle SC SC SC work (including NASDA's equipment)
inspection of loading equipment SC SC SC work -04 propellant loading and high-pressure gas charging SC SC SC work post-treatment of loading equipment SC SC SC work
storage of loading equipment SC LV or SC option or SC work C
treatment of waste SC LV or SC option or SC work
6-12 Table 6.2.1(5/6) Mission Integration WBS A
※1 Responsibility ※2 WBS No. Items Description Remarks Cost Work
phase 3 joint work of LV and SC standard (SC work) (costs are shared)
SC / PLA mating LV (SC) LV (SC) standard (SC work)
PLA / PSS mating LV (SC) LV (SC) standard (SC work)
SC encapsulation into the fairing LV (SC) LV (SC) standard (SC work)
SC / fairing mating LV (SC) LV (SC) standard (SC work) onto the second stage -05 SC system test
testing SC SC SC work
operation of elevator type floor LV LV standard
SC work and Y-0 reheasal SC and LV SC and LV LV work
Y-2 ~ Y-0 operation SC SC SC work A
analysis of liquid and gas
gas analysis SC LV option 11 propellant analysis
-06 chemical SC LV option
contamination SC LV option
IPA circulation and filtering SC LV option
post treatment of preparation / operation SC LV option scrubber and bubbler
consumable materials
propellant SC LV or SC option or SC work -07 high-purity gas and IPA, etc. for purging, flushing, etc. SC LV or SC option or SC work
others SC LV or SC option or SC work
SC facility operation
one person for standard working day -08 supporting operators LV LV standard (within 45)
driver for special motor vehicle if needed SC LV (SC) option (SC work)
SC : Spacecraft LV : Launch Vehicle
6-13 Table 6.2.1(6/6) Mission Integration WBS A
Responsibility ※1 WBS No. Items Description Remarks ※2 Cost Work
SC transportation in launch site
SC handling SC SC (LV) SC work (option) -09 canistar preparation dia 5m x 10mH SC LV option
transportation operation SC LV option
other services
security control in SC area SC LV option
clean tent class 1000 LV SC standard A
equipment calibration user's equipment LV LV standard
refrigerator for battery preparation LV LV standard
operation LV SC standard
FALSE floor for SC checkout equipment
11 telephone ¥ facsimile extension LV LV standard
outer line SC LV option
-10 copy machine preparation SC LV option
copying paper preparation SC LV option
SC storage for launch postponement SC SC and LV SC work and option
modem
Blast Shield
test connector
customer's sticker on the PLF SC LV option
special suits to be prepared by SC, SC work or protective suit and dust-proof suit SC or LV SC or LV if required for SC standard
special cleaning of clean room if specially required SC LV option
umbilical line change LV LV standard
12 orbit parameter of SC separation within 30 minutes after separation LV LV standard
report of post-flight analysis 13 orbit parameter after SC separation within one week after launch SC SC SC work (report of SC tracking data)
SC : Spacecraft
6-14 6.3 Spacecraft / H-IIA and Spacecraft / Launch Operations Interface Control
6.3.1 Interface control document The “Spacecraft / H-IIA interface control specifications (ICS)” related to launch capability, interface restrictions and launch operations are edited and maintained through the interface meetings to be held as required in the spacecraft design, fabrication and test phases, thereby clarifying the requirements and characteristics inherent to each mission. Table 6.3.1 shows an example of the table of contents for the ICS.
6.3.2 Coordination items and timing The master schedule is prepared according to results of coordination with each spacecraft organization by NASDA.
6-15 Table 6.3.1 Standard spacecraft / H-IIA interface control specifications
0 SCOPE 3.4 SPACECRAFT CONFIGURATION AT LAUNCH 3.5 DATA ACQUISITION IN LAUNCH PHASE 1 INTRODUCTION 3.6 SPACECRAFT MASS AND DYNAMIC PROPERTIES 1.1 H-IIA MISSION 3.6.1 Fundamental Frequency 1.2 SPACECRAFT MISSION OBJECTIVES 3.6.2 Primary Structure 1.3 SPACECRAFT DESCRIPTION 3.6.3 Secondary Structure and Flexible Elements 1.4 H-IIA DESCRIPTION 3.7 SPACECRAFT DIMENSIONING REQUIREMENTS 1.4.1 Outline of the H-IIA Launch Vehicle 3.7.1 Mass Constraints 1.4.2 H-IIA First Stage 3.7.2 CoG Constraints 1.4.3 H-IIA Second Stage 3.7.3 Inertia Constraints (T.B.D.) 1.4.4 H-IIA Guidance and Control System 3.7.4 Balancing Constraints 1.4.5 Payload Fairing 3.8 MECHANICAL INTERFACES 1.4.6 Payload Adapter 3.8.1 Axis Definition 1.5 OSAKI RANGE DESCRIPTION 3.8.2 Assembly Characteristics 1.5.1 Yoshinobu Launch Complex 3.8.3 Access and Mounting 1.5.2 Spacecraft and Fairing Assembly Facilities 3.9 ENVIRONMENTAL CONDITIONS UNDER THE 2 APPLICABLE DOCUMENTS FAIRING DURING LAUNCH PREPARATION 2.1 SPECIFICATIONS 3.9.1 Air Conditioning 2.2 INTERFACE DRAWINGS 3.9.2 Thermal Characteristics 2.3 RELEVANT DOCUMENTS 3.10 IN FLIGHT ENVIRONMENTS 2.4 DOCUMENT AMENDMENT PROCEDURE 3.10.1 Mechanical Environments 2.4.1 Introduction 3.10.2 Other Environments 2.4.2 Procedure 3.11 ELECTRICAL INTERFACES 3 REQUIREMENTS 3.11.1 Earth Potential Continuity 3.1 GENERAL 3.11.2 Umbilical Link 3.2 MISSION CHARACTERISTICS 3.11.3 Electrical Link to the H-IIA Second Stage 3.2.1 Launch Capability 3.12 RF LINK INTERFACE 3.2.2 Launch Date 3.12.1 Spacecraft TX / RX Characteristics 3.2.3 Launch Vehicle Configuration 3.12.2 H-IIA TX / RX Characteristics 3.2.4 Launch Window 3.12.3 Spacecraft Transmission Plan 3.3 INJECTION ORBIT PARAMETERS AND 3.12.4 H-IIA Transmission Plan CONDITIONS 3.12.5 Radio EMC Compatibility 3.3.1 Flight Plan 3.12.6 Operational Constraints 3.3.2 Orbit Parameters 3.12.7 Radio Link Requirements 3.3.3 Injection Accuracy 3.13 FLUID INTERFACE 3.3.4 Conditions at Separation 3.14 PYROTECHNIC INTERFACE
4 SPACECRAFT PREPARATION FACILITY 6.1.3 RF / Video / Data Links INTERFACES 6.1.4 Umbilical Links 4.1 OPERATION FLOW 6.1.5 Main Power 4.1.1 Spacecraft Operations 6.1.6 Fluids 4.1.2 Combined Operations 6.1.7 Telecommunications 4.2 BUILDING STA1 / STA2 (Spacecraft Test and 6.2 SPACECRAFT / BLOCK HOUSE INTERFACE Assembly) 7 RANGE FACILITIES 4.3 BUILDING SFA (Spacecraft and Fairing Assembly) 7.1 TRANSPORT AND HANDLING 4.4 BUILDING TSA (Third Stage and Spacecraft 7.2 FLUID AND PROPELLANTS Assembly) 7.3 TECHNICAL SUPPORT 4.5 BUILDING VAB (Vehicle Assembly) 7.4 SAFETY FACILITIES 4.6 ML (Mobile Launch table) 7.5 RANGE COMMUNICATIONS 4.7 RF / VIDEO / DATA LINKS IN EACH BUILDING 7.6 TELECOMMUNICATIONS 4.8 ELECTRICAL LINKS IN EACH BUILDING 7.7 BUILDING FACILITIES 4.9 TELECOMMUNICATION IN EACH BUILDING 7.8 STORAGE AREA 4.10 SPACECRAFT TRANSPORT CONTAINER 7.9 MISCELLANEOUS 4.10.1 Canister Description and Interface 4.10.2 Requirements during Transfer on Site 8 PLANS 4.11 OTHERS 8.1 DEVELOPMENT AND TEST PLAN 8.2 LAUNCH OPERATION PLAN 5 ENVIRONMENTAL CONDITIONS, INTERFACES 8.3 PLANNING CONTROL DURING TRANSFER 8.3.1 Planning Control 5.1 TRANSFER WITHOUT FAIRING 8.3.2 Launch Site Meeting 5.1.1 Environmental Conditions during Transportation 9 LAUNCH OPERATION REQUIREMENTS 5.2 TRANSFER WITH FAIRING 9.1 REQUIREMENT FOR WASH OF TEST 5.2.1 Environmental Conditions during EQUIPMENT Transportation 9.2 REQUIREMENT FOR ANALYSIS OF GAS & 5.3 TRANSFER FROM VAB TO LAUNCH PAD (LP) LIQUID 5.3.1 General 10 SAFETY REQUIREMENTS 5.3.2 Ground Environments 5.3.3 Interfaces 11 DOCUMENT ITEMS DELIVERY AND 6 SPACECRAFT / LAUNCH AREA INTERFACE REVIEWS 6.1 SPACECRAFT / LP INTERFACE 12 RESPONSIBILITY MATRIX 6.1.1 Environmental Conditions 6.1.2 Access Facilities / Requirements
6-16 6.4 Mission Analysis
6.4.1 General Mission analysis is conducted by the vehicle organization to confirm that the spacecraft is injected into the required orbit in the required condition. If a special orbit is specified, analyses for the mission plan and environment will be made according to the following steps, which are to be initiated almost at the same time as the spacecraft development. If the spacecraft mission requirements (mass, orbit, etc.) are similar to those of prescribed missions, the reference planning phase may be omitted.
(1) Reference planning phase (From 32 months to 20 months before launch) In the Reference Planning Phase, the major scenario of the mission plan is determined and identified problems are solved.
(2) Detailed planning phase (From 18 months to 8 months before launch) The Detailed Planning Phase determines and approves the mission plan; the vehicle is actually launched according to the mission plan completed in this phase. Therefore, the data required for launch operation, tracking control operation, and flight safety operation are generated in this phase.
(3) Final planning phase (From 6 months to immediately before launch) In the Final Planning Phase, the measured data for both the vehicle and the spacecraft are used to reconfirm that there is no problem in launching the vehicle according to the mission plan completed in the detailed planning phase.
6.4.2 Reference planning phase In this phase, the major scenario of the flight plan is determined, and the data below are generated to solve identified problems and to meet the spacecraft development requirements. Items b) through g) may be normally omitted except when the trajectory must be changed drastically from prescribed missions.
a) Mass property b) Required trajectory c) Preparation and estimation of reference trajectory
6-17 d) Orbit injection accuracy e) Loads imposed on the spacecraft during flight f) Position, velocity, and attitude at orbit injection g) Sequence of events h) History of the sun direction during flight i) Spacecraft separation conditions j) Thermal conditions imposed on the spacecraft during prelaunch and flight
6.4.3 Detailed planning phase This phase solves problems identified in the reference planning phase and completes the mission plan. The following data are generated based on the mission plan:
a) Preparation and estimation of the detailed trajectory b) Orbit injection accuracy c) Load imposed on the spacecraft during flight d) Position, velocity and attitude at orbit injection e) Sequence of events f) History of the sun direction during flight g) Spacecraft separation conditions h) Flight safety i) Thermal conditions imposed on the spacecraft during prelaunch and flight
6.4.4 Final planning phase In this phase, the following work is performed for the final confirmation of the mission plan:
a) Review of the mission plan (mission readiness review) b) Final confirmation of the mission plan based on the measured data for both the spacecraft and the vehicle c) Influence analysis using measured wind at launch site.
6.4.5 Spacecraft coupled loads analysis (CLA) Spacecraft / H-IIA CLA is conducted to support the spacecraft design and verification with the spacecraft structural dynamic environments (refer to § 3.2) induced by the H-IIA launch. Two cycles of the CLA, preliminary (reference
6-18 phase) CLA and final CLA, are conducted in normal procedure. Each CLA cycle is conducted at the spacecraft customer’s request. The preliminary CLA using preliminary spacecraft mathematical model can support spacecraft customer to assess compatibility of the H-IIA launcher. The final CLA using a final spacecraft mathematical model shall verify the specified values of acceleration, loads, and deflections in the ICS. In case that the payload design properties exceed the specified values in the H- IIA User’s Manual, the extensive coordination based on the CLA can tailor interface requirements between the spacecraft and the H-IIA launch vehicle, and may enable the H-IIA to launch the protruded payload.
6.4.6 Integrated thermal analysis A Integrated thermal analysis is conducted to verify the spacecraft thermal environments during the H-IIA launching phase. At the spacecraft customer’s request, the integrated thermal analysis using a spacecraft mathematical model is conducted to verify the specified values of spacecraft thermal environments in the ICS. The integrated thermal analysis covers thermal environments from encapsulation to separation of the spacecraft.
6.5 Safety Reviews
6.5.1 Payload safety requirements The design and launch operations of the spacecraft must satisfy requirements of “NASDA-STD-14B: Launch Vehicle Payload Safety Requirements”. A The spacecraft organization should enforce the safety program according to requirements of NASDA-STD-14B. A The requirements on a safety program plan are described in § 6.5.1., § 6.5.2.
6.5.2 Requirements on the safety program plan of the spacecraft organization The spacecraft organization must establish a safety program plan according to NASDA-STD-14B and submit to NASDA arranging its contents for the document A “Payload Safety Program Plan”. When the spacecraft organization establishes the safety program plan, it is recommended to coordinate with NASDA through an interface meeting.
6-19 6.5.3 Safety review data package The spacecraft organization must submit the safety data package to NASDA according to NASDA-STD-14B in every operation phase. NASDA holds internal A safety review of this safety data package as a basis for safety. NASDA will inform the spacecraft organization of review results. Details on these contents are described in the document “CFX-97010: Outline of Payload System Safety Review for external Users”.
6.5.4 Outline of NASDA safety review NASDA system safety program plan is established for each spacecraft. Before establishment of this plan, NASDA will coordinate with the spacecraft organization at the interface meeting. Details are explained in CFX-97010. According to NASDA system safety program plan, NASDA reviews the safety of the spacecraft based on the safety data package submitted by the spacecraft organization, informs the spacecraft organization of review results and requires any treatments if necessary.
6.5.5 Interface The safety review meeting is set up based on coordination between the spacecraft organization and NASDA.
6.6 Reviews and Other Meetings The following describes the reviews and other meetings required to launch the spacecraft and launch vehicle.
6.6.1 Reviews before launch operation
6.6.1.1 Mission readiness review (MRR) This meeting is held once unless otherwise required. The review is carried out in conformity to the mission analysis report, and the results are confirmed to satisfy the requirements of the ICS. The user has the right to attend this meeting as an observer. This meeting is held about 3 months prior to the launch day.
6.6.1.2 Spacecraft interface confirmation review (SIC) This meeting confirms that the spacecraft manufactures, tests and working
6-20 program have conformed to the requirements of the ICS, and launch operations can be proceeded at the launch site. This meeting is held just before spacecraft transportation to Tanegashima Space Center (TNSC).
6.6.2 Reviews for launch operation
6.6.2.1 Launch vehicle readiness review (LVRR) This meeting is held in TNSC to verify that the launch vehicle and AGE and mission analysis have satisfied technical proceedings to be ready for the countdown operation. The verified results are to be reported in FRR. The user has the right to attend this meeting as an observer. This meeting is held toward Y-5.
6.6.2.2 Spacecraft readiness review (SCRR) This meeting is held in TNSC to verify that the payload and GSE have satisfied technical proceeding to be ready for the countdown operation. The verified results are to be reported in FRR. This meeting is held just before encapsulation of the spacecraft into the payload fairing.
6.6.2.3 Flight readiness review (FRR) This review confirms that all the organizations involved in the launch are ready for the countdown operations to be started. Each organization reports the current progress of preparation and conclusion of the review meeting. When reviews are finished, countdown operations are started after approval of the launch director is obtained. The user must report the current preparation progress of the spacecraft and conclusion of the reviews. This meeting is held toward Y-4.
6.6.3 Safety review The safety review is explained in § 6.5.
6-21 6.6.4 Meetings
6.6.4.1 Interface meeting This meeting coordinates matters to be described in the ICS and other interface requirements, and is held whenever required.
6.6.4.2 Daily meeting This meeting confirms the progress of both the launch vehicle and the spacecraft organizations and facilities coordination. It is held every day (except on holidays). SIM and related members of the spacecraft organization have to attend this meeting.
6.6.4.3 Launch site readiness meeting The purpose of this meeting, held just before the arrival of the spacecraft and associated equipment at TNSC, is to verify that the facilities are configured according to the requirements contained in the ICS.
6.6.4.4 Precountdown coordination meeting
6-22
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A P P E N D IX 1 .
H IS T O R Y O F N A S D A L A U N C H V E H IC L E S APPENDIX 1. HISTORY OF NASDA LAUNCH VEHICLES A1.1 General Japan has developed four kinds of large launch vehicles, N-I, N-II, H-I and H-II since early seventies when Japan started active space development under NASDA. Figure A1.1.1 shows configuration of these launch vehicles. Table A1.1.1 summarizes launch results of NASDA launch vehicles. BC Abstracts of the H-I and H-II launch vehicles are described in the following sections.
A1.2 Abstract of The H-I Launch Vehicle NASDA has used the H-I launch vehicle as the main launch vehicle since 1986. It has a Delta-based first stage (the MB-3 engine), which is basically the same as the N-I and N-II rockets with the nine strap-on boosters (SOBs), and a domestically developed LOX / LH2 second stage propulsion system (the LE-5 engine). Up to 1992, nine payloads (geostationary meteorological satellites, communications satellites, and TV broadcasting satellites) had all been successfully launched by the H-I. The launch site for the H-I is Osaki Range of the Tanegashima Space Center (TNSC).
A1.3 Abstract of The H-II Launch Vehicle The H-II launch vehicle was designed to serve as NASDA’s main space transportation system in the 1990s to meet the demands for large satellites launching at lower cost yet maintaining a high degree of reliability. It is capable of sending a 4-ton class payload into geostationary transfer orbit. The H-II is a two-stage launch vehicle equipped with two large scale solid rocket boosters (SRBs) for thrust augmentation. A new liquid propellant (liquid hydrogen / liquid oxygen) engine, called LE-7, is developed for the first stage. The LE-7 is a high performance engine adopting a high-pressure staged-combustion cycle. The LE-5A engine, which is the improved version of the LE-5 engine developed for the H-I, is used in the second stage. A strapped-down inertial guidance system utilizing ring laser gyros is employed for the guidance system. The standard payload fairing is 4 meters in diameter and 12 meters in length so that it can encapsulate a payload up to 3.7 meters in diameter and 10 meters in length. Besides carrying satellites into low earth orbit and geostationary transfer orbit, the H-II is capable of launching planetary probes. The H-II is to be launched from Yoshinobu launch complex which has been newly constructed at the TNSC.
A1-1 Table A1.1.1 Launch results summary of NASDA launch vehicles (from 1975 to 2000) BAC
Number of Vehicles Mission Launch Success Launched GTO LEO SSO Vehicle Rate (%) Success Failure Success Failure Success Failure Success Failure
N-I 61862140Ð0*2 A
N-II 8 0 100 7 0 Ð 0 1 0
H-I 9 0 100 6 0 1 0 2 0
*1 H-II 5271324010*3 AB
Total 28 3 90 18 3 7 0 4 0
*1 : Include three dual launch missions. *2 : The third stage collided with the spacecraft after separation. A *3 : In the restart phase, the second stage’s engine was cutoff before scheduled plan (F#5). B : The first stage’s main engine was cut off before scheduled plan (F#8). C
A1-2 50 m
40 m
30 m
20 m
10 m
(*: Dual launch)
Solid Rocket Boosters (2 Iarge solid motors)
First stage: liquid propellant (LOX/LH2) engine (LE-7)
Second stage: liquid propellant (LOX/LH2) engine (LE-5A) Inertial Guidance System
Large fairing
H-ll
1994
5+(3*)
4,000 kg
Fairing
Third stage: solid rocket motor
Second stage: liquid propellant (LOX/LH2) engine (LE-5) Inertial Guidance System
First stage: liquid propellant (LOX/RJ-1) engine
Strap-on boosters (9 solid motors)
H-l
9
1986
1,100 kg
Fairing
Third stage: solid rocket motor
Second stage: improved liquid propellant engine (NTO/A-50) inertial guidance system First stage: liquid propellant (LOX/RJ-1) engine
Strap-on boosters (9 solid motors)
Configuration summary of NASDA launch vehicles
N-ll
8
1981
700 kg
Figure A1.1.1
Fairing
Third stage: solid rocket motor Second stage: liquid propellant engine (NTO/A-50) Radio Guidance System
First stage: liquid propellant engine (LOX/RJ-1)
Strap-on boosters (3 solid motors)
N-1
7
1975
260 kg
Launched Satellites First Flight GTO Mass
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E D IS L A N D A N D T A N E G A S H IMA R S P A C E C E N T E R APPENDIX 2. OUTLINE OF TANEGASHIMA ISLAND AND TANEGASHIMA SPACE CENTER A2.1 Tanegashima Island
A2.1.1 Location and topography Tanegashima Island is located about 40 km southeast of Sata Point, Kagoshima Prefecture, the southernmost part of Kyushu Island, and about 115 km due south of Kagoshima City. The entire island belongs to Kagoshima Prefecture and is composed of one city and two towns: Nishino-omote City, Nakatane-cho of Kumage District, and Minamitane-cho of Kumage District. The area is 445 km2 and the population is about 37,000. Topographically, Tanegashima is a long and narrow island lying from NNE to SSW. It is 58 km long and a maximum of 12 km wide, but only 6 km wide at the narrowest point near the center of the island. It is a flat island, which basically consists of coastal terraces, where the narrowest area is about 8 km long at the center and several stretches of hills of about 200 m above sea level run northeast and southeast of the narrowest area. Rocky beaches and small dunes are also found along the coast. In the southern part, there are a number of cliffs steeply rising from the sands along the coastline, creating scenic beauty at Kumano, Takesaki, and Kadokura Point. Yakushima Island has an area of about 505 km2 and lies to the west of Tanegashima Island. This island is a large round lump of land, most of which is mountainous . The central part of the island consists of a series of mountains including Miyanoura-dake Mountain (1,935 m), the highest in the region of Kyusyu. Figure A2.1.1 indicates the location of Tanegashima and Tanegashima Space Center.
A2.1.2 Climate Tanegashima island has a yearly mean temperature of 19.4°C, a maximum temperature of 35.9°C, and a minimum temperature of 0.0°C. The rainfall is heavy, about 2,240 mm yearly mean rainfall. About 1000 mm of this is concentrated from June to September. The yearly mean humidity is 76%.
A2.1.3 Traffic (1) Air route Tanegashima Airport, a Class 3 airport (1,500 m x 30 m wireway), is located in Nakatane-cho. Regular flight services are available; five flights to and from Kagoshima Airport and one flight to and from Osaka Airport (by YS-11 or SAAB).
A2-1 The flight time is 35 minutes between Kagoshima and Tanegashima Airports and 110 minutes between Osaka and Tanegashima Airports. It is about 3 hours from Tokyo to Tanegashima including transit time.
(2) Sea Route There are three regular cargo/passenger ferry boats and two high-speed jet boats in service between Tanegashima and the Kagoshima Prefecture. These regular liners connect Nishino-omote Port and Kagoshima Port. Tanegashima has another port at Shimama in Minamitane-cho, where chartered vessels disembark.
(3) Land transportation All national roads, prefectural roads and city roads in Tanegashima are well prepared and maintained; all roads used for transporting spacecraft and GSE from Tanegashima Airport, Nishino-omote Port, and Shimama Port to Kaminaka are well paved. Regular bus lines are available from main points of the island including Tanegashima Airport and Nishino-omote Port to Kaminaka. Taxis and rent-a-cars are also available. Time required by automobile, such as by taxi, between two main points is as follows:
Nishino-omote Port ⇔ TNSC : 80 minutes (47 km) Nishino-omote Port ⇔ Kaminaka : 70 minutes (43 km) Tanegashima Airport ⇔ TNSC : 40 minutes (26 km) Tanegashima Airport ⇔ Kaminaka : 30 minutes (17 km) Kaminaka ⇔ TNSC : 15 minutes (10 km)
A2-2 Kagoshima Airport
Kagoshima Port
to Osaka N
Sata Point Nishinoomote Port
Tanegashima Airport Shimama Port Tanegashima Island
Tanegashima 0 50km Yakushima Island Space Center
To Kagoshima Airport To Osaka Airport N
Nishinoomote Port Nishinoomote City
Nogi Radar Station Nogi Command Station Nogi
Nakawari Collimation Tower Masuda Tanegashima Airport
Masuda Tracking and Kamisato Nakatane Town Communication Station (MTCS) Collimation Tower
Shimama Port Shimama Kumano Port Kaminaka Tanegashima Space Center (TNSC) Minamitane Town Osaki Range (H-IIA, J-I) Takesaki Range (TR-1A)
Uchugaoka Radar Station Kadokura 0 10Km Point Third Optical Tracking Station
Figure A 2.1.1 Location of Tanegashima space center
A2-3 A2.2 The Tanegashima Space Center
A2.2.1 General Tanegashima Space Center (TNSC), located at the southeastern end of Tanegashima island, is the largest rocket range in Japan with a total site area of 8.6 million square meters. The site facilities include the Takesaki Range for small rockets, the Osaki Range for H-IIA and J-I launch vehicles, the Masuda Tracking and Communication Station (MTCS) 18 km north of Osaki, the Nogi Radar Station 6 km north of Masuda, the Uchugaoka Radar Station 6 km west of Osaki, and the optical tracking stations. Ground firing test facilities to verify liquid and solid propellant rockets and other necessary facilities are also included. The major tasks of TNSC are prelaunch preparation operations including various checks, adjustments, and assembly of satellites and launch vehicles; countdown operation for launch; and tracking and control after launch. Thus, TNSC plays a major role in launching application satellites. It also plays a part in space development by conducting ground firing tests for solid booster motors and liquid rocket engines as well as conducting material tests for space applications using small rockets. Figure 2.2.1 shows the layout of the major facilities in TNSC.
A2.2.2 Osaki Launch Range The Osaki Launch Range consists of the Yoshinobu Launch Complex for launching the H-IIA launch vehicle for large spacecraft, spacecraft-related test facilities, and hazardous material related test and storage facilities. The launch pad of the Yoshinobu Launch Complex is located at 30° 23’ 50” north latitude and 130° 58’ 47” east longitude. Figure A2.2.2 illustrates the layout of the major facilities in the Osaki Launch Range. Figure A2.2.3 shows an overview of the Yoshinobu Launch Complex. And Figure A2.2.4 shows a typical H-IIA launch operations process in the Osaki Launch Range. The following paragraphs describe the major facilities in the Osaki Launch Range.
A2.2.2.1 Yoshinobu Vehicle Assembly Building (VAB) The VAB is a facility for assembling, inspecting, and adjusting all components of the launch vehicle (first stage, second stage, and SRB-A) transported from factories. Inside this VAB, the Movable Launcher (ML) is housed, and the launch A vehicle is assembled on the ML and inspected.
A2-4 Then the spacecraft / fairing composite is transported to the VAB and mated to the second stage. After that, final preparations of the spacecraft as well as the launch vehicle are conducted prior to the launch day maintaining environments for the spacecraft. Details are shown in A 2.6.
A2.2.2.2 Movable Launcher (ML) A The ML, on which the spacecraft/launch vehicle assembly is loaded, moves to the launch pad on the launch day. The GSE of the spacecraft is set in the pressurized room inside the ML, and moved with the spacecraft. Details are shown in A2.7.
A2.2.2.3 Block House (B/H) The B/H, which is established under the ground, coordinates and supervises the launching operation for the spacecraft and the launch vehicle by instructing, operating, and monitoring the launch vehicle assembly, adjustment, checkout, propellant loading and other launch activities at the Launch Complex and by remotely conducting the launch control operation including data transmission to the Range Control Center. Start of the countdown and the automatic launch sequence system is instructed from B/H. Launch control operations and data monitoring of the spacecraft are not conducted in B/H.
A2.2.2.4 No. 2 Spacecraft Test and Assembly Building (STA2) Inside the No. 2 Spacecraft Test and Assembly Building, the transported large spacecraft is unpacked, visually inspected, and assembled. The spacecraft then undergoes various tests and inspections including radio characteristic tests, functional tests, and compatibility tests. In addition, the overall spacecraft control is conducted from this building, including remote control over the checkout during countdown operations. Details are shown in A 2.3.
A2.2.2.5 Spacecraft and Fairing Assembly Building (SFA) In this building, the high-pressure leak test of the propellant system is performed, the propellant is loaded and pressurized, and the pyrotechnics and the solid propellant motor are attached. Finally, the spacecraft is installed in the fairing. Details are shown in A2.4.
A2-5 A2.2.2.6 No. 1 Spacecraft Test and Assembly Building (STA1) The STA1 is mainly for testing and assembling small or medium size spacecraft in case of a dual launch. The transported spacecraft is unpacked, visually inspected, and assembled here. It then undergoes various tests and inspections including radio characteristic test, functional test, and compatibility test. Checkout of the spacecraft during launching is conducted in the STA2 (not STA1).
A2.2.2.7 Third stage and Spacecraft Assembly Building (TSA) The spacecraft assembled and tested using the STA1 is transferred to the TSA, where the high-pressure leak test of the propellant system is performed, the propellant is loaded and pressurized, and the pyrotechnics and the solid propellant motor are attached. The spacecraft is installed in the fairing in the SFA. Details are shown in A2.5. C
A2.2.2.8 Non-Destructive Test Facility (NDTF) A large X-ray CT system is installed inside the NDTF, for the X-ray inspection of the spacecraft solid propellant motor.
A2.2.2.9 Nakanoyama Telemetry Command Station The Nakanoyama Telemetry Command Station monitors and records in-flight launch vehicle conditions, receiving telemetry data on acceleration, pressure, temperature and so on, which are transmitted from each stage after lift-off. The Station then relays this data to the Takesaki Range Control Center for flight safety operation. It is also equipped with a flight safety command transmitter. A
A2.2.2.10 Other facilities (1) First solid propellant storage, second solid propellant storage and pyrotechnics storage Pyrotechnics for the launch vehicle, spacecraft system, and solid propellant motor are stored in these facilities. (2) Third solid propellant storage, fourth solid propellant storage The solid rocket motor is stored in these facilities. (3) Solid Booster Test Building (SBB) Various tests on the solid rocket motor and pyrotechnics are performed here. (4) Osaki Supporting Launch Building (SLB) Conducts analysis of gases and propellants, and flushing of cleanliness control
A2-6 components. (5) Yoshinobu Launch Building (LB) Operates the air-conditioning system of the launch vehicle including the fairing at the LP. (6) Osaki 80 m meteorological observation tower Observes the weather in Osaki Launch Range. (7) Osaki power station Generates power mainly for Osaki area in TNSC. This station switches between private and commercial power supply, monitors the power supply system, and monitors the air-conditioning systems in major buildings. (8) Osaki Office 1 An office for launch vehicle operators. (9) Osaki Office 2 An office for launch vehicle operators. (10) Osaki aerospace ground equipment (AGE) storage Tools, consumables and instruments used in TNSC are stored here and lent. (11) Yoshinobu combustion test facility Ground combustion test of the first-stage engine (LE-7A) for H-IIA is performed. (12) General accident prevention control office Responsible for accident prevention control in TNSC.
A2.2.3 Takesaki Area Relevant control facilities are located in the Takesaki area. The following are the main facilities.
A2.2.3.1 Administration Building (AB) General office building for TNSC workers; contains a reception area staffed by the Administration Department.
A2.2.3.2 Takesaki Range Control Center (RCC) The Takesaki RCC is responsible for information collection, analysis, instruction, coordination, and monitoring throughout the entire launch operations of the launch vehicles and spacecraft in Tanegashima, including prelaunch operations, ground safety operation, launching, and tracking, to ensure smooth progress in these operations. Details are shown in A 2.8.
A2-7 A2.2.3.3 Takesaki Observation Stand This stand is for press and broadcasting personnel and is equipped with an interview room for the reporters; waiting rooms for reportorial staff guests, and launch operation crew; TV-monitoring room; time indicating equipment; dark rooms for developing film; etc.
A2.2.3.4 Takesaki Space Exhibition Hall To further understanding of space development, interrelations between the space and mankind, the origination of space development, contribution of space development to mankind, the structure and functions of satellites and launch vehicles, exhibits for launching, tracking and controlling are exhibited and open to the public.
A2-8 to Hirayama First Optical Tracking Station
Non-Destructive Test Facility
A
Osaki Power Station 3 No. 2 Spacecraft Test and Assembly Building 1
Spacecraft and Fairing Assembly Building 4 16 2 5 Yoshinobu (H-IIA) Launch Complex 10 15 6 8 9 17 7 14 11 Osaki Range 12 13
Nakanoyama Telemetry Station Osaki Range Control Center Osaki Launch Complex to Kukinaga
1 Third stage and Spacecraft Assembly Building 2 No. 1 Spacecraft Test and Assembly Building Fourth Optical 3 Tracking Station Spin Test Building 4 Rocket Hill Solid Booster Test Building Observatory 5 Osaki Launch Support Building 6 Osaki Restaurant 7 Osaki AGE Storage 8 Osaki Office 1 9 Osaki Office 2 General Accident Prevention Control Office 10 Propellant Storage Facility Sea Observation Radar 11 Osaki 80m Meteorological Observation Tower Tanegashima Light House 12 Osaki Block House 13 Osaki Mobile Service Tower 14 Yoshinobu Engine Test Stand 15 Yoshinobu Block House Static Firing Test 16 Yoshinobu Vehicle Assembly Building Facility for Solid Motor 17 Yoshinobu Pad Service Tower Takesaki 22 Launch Complex 18 Administration Building 18 19 Takesaki Range Control Center
19 20 Guest House to Kukinaga TR-I A Rocket 20 21 Launcher 21 L band Radar Building 23 22 Takesaki Balloon Filling Building 24 25 23 Takesaki Power station 24 26 Takesaki Range Takesaki Gate 25 Takesaki Space Exhibition Hall Takesaki 26 Kamori Hill Observation Observation Station A
Figure A2.2.1 Location of Facilities of Tanegashima Space Center
A2-9 A
Non-Destructive Test Facility
N No.2 Spacecraft Test to Hirota and Assembly Building
Spacecraft and Fairing Spin Test Building Assembly Building Osaki Power Station A Spin Liquid Propellant Measument storage Building Liquid Oxydizer Osaki 5 roads storage crossing gate Third Stage and Satellite Solid Booster Building Yoshinobu Vehicle Assembly Building Assembly Building
No.1 Spacecraft Test Osaki No.1 Support Yoshinobu Block House and Assembly Building LSB gate Garage Osaki Supporting Liquid Hydrogen Storage Reserroir Launch Building Propellant New Yoshinobu Block House (underground) Osaki Restaurant Storage Facility Yoshinobu Second Camera Room Osaki office 2 Osaki office 1 Yoshinobu (H-IIA) Launch Complex LE-7 Engine Launch 0 100 200 300 400 500 Building Tool Room Test Stand Pad Service Tower (Launch Pad)
New Launch Pad 80m Meteorological Observation Tower Liquid Oxygen Storage
High-pressure Sea shore gate Gas Storage
to Kukinaga, Kaminaka
Pacific Ocean
to Takesaki Administration Building
Figure A2.2.2 Location of major facilities in Osaki Launch Range
A2-10 Figure A2.2.3 Overview of New Yoshinobu Launch Complex C
A2-11 STA 2 Lift-off
No.2 Spacecraft test and assembly Launch pad building (LP) (STA 2)
¥ Spacecraft integration ¥ Terminal Countdown B ¥ Spacecraft functional test
A
Spacecraft and fairing Vehicle assembly assembly building building ML (SFA) (VAB)
¥ Propellant loading
SFA VAB
Transporting A from SFA to VAB
Figure A2.2.4 H-IIA launch operations process A
A2-12 A2.3 No. 2 Spacecraft Test and Assembly Building (STA2)
A2.3.1 General The STA2 is used for a series of operations for a large spacecraft to be loaded into the H-IIA launch vehicle after delivery to the launch site. Operations include installation and adjustment of ground support equipment (GSE) validation, preparation for assembling the spacecraft after delivery, alignment measurement after assembly, functional tests, RF link test with Masuda Tracking and Communication Station, charging and discharging of batteries, weight measurement, and preparation for transportation into SFA. This building has two air lock rooms, two spacecraft preparation halls, two GSE storage rooms and three checkout rooms so that checkouts of two (or three A in maximum) spacecraft are conducted individually. During Phase 3 operations, the control and the checkout of the spacecraft, which is loaded onto the launch vehicle, are conducted by remote control from this building. Specifications of the first floor section are as follows. a) Cleanliness : class 100,000 b) Temperature : 22 ± 3 ºC c) Humidity : 50 ± 10 % RH e) Power supply : Frequency_60 Hz Panel capacity_Table A2.3.1 Location of outlets - shown in H-IIA Payload-Related Facilities A and GSE Manual
A2.3.2 Main Structure and Functions of STA2 The STA2 is a four-storied, steel-frame construction building of about 8,580 m2 floor space. (Total area of first floor is 4,550 m2) The first floor section contains mainly two air lock rooms, two spacecraft preparation halls, two GSE storage rooms, three checkout rooms, control room, A unpacking room, air conditioning equipment room 1, two cloak room, rest room, entrance, reception, stairs (first to fourth floor), outdoor ladder (first floor to the roof), and two high-pressure gas storages. The second floor section contains five office rooms, two meeting rooms, two storage rooms, kitchenette, rest room, tour path for visitors and so on. A The third floor section contains eight office rooms, two meeting rooms, two storage rooms, kitchenette, rest room, warehouse and two electric power rooms. The fourth floor section contains air-conditioning room (#2). Figure A2.3.1 shows the STA2.
A2-13 Figure A2.3.1 No. 2 Spacecraft Test and Assembly Building (STA2)
A2-14 Table A2.3.1 Summary of power supply (STA2) (1/2)
A Install place Name Power Quantity Remarks A3φWC 210 V 4A1100 A3φWC 210 V 3A510 P-2 A1φWC 100 V 2*580 A distribution board A1φWC 115 V 2*520 A A1φWC 115 V 2*240 A Uninterruptable A1φWC 210 V 2A550 *CVCF system Checkout room 1 A3φWC 480 V 4A1100 A3φWC 210 V 4*1100 A P-2-1 distribution A3φWC 210 V 4*520 A board A1φWC 115 V 2*330 A A1φWC 115 V 2*230 A WVall outlet A1φWC 100 22290 A x First class grounding terminal Grounding 4 Terminal box A3φWC 210 V 4A520 A3φWC 210 V 4A1200 A3φWC 100 V 3A1500 A3φWC 100 V 3A520 P-5 Uninterruptable distribution A3φWC 100 V 3A270 *CVCF system board A3φWC 480 V 4A715 Checkout room 2 A1φWC 100 V 2*580 A A1φWC 115 V 2*550 A A1φWC 115 V 2*290 A A1φWC 115 V 22260 A x Wall outlet A1φWC 100 V 22270 A x First class grounding terminal Grounding 3 Terminal box A3φWC 210 V 4A1200 A3φWC 210 V 4A520 A1φWC 100 V 3A1500 P-6 Uninterruptable distribution A3φWC 480 V 4A715 *CVCF system board A1φWC 100 V 2*580 A Checkout room 3 A1φWC 115 V 2*560 A A1φWC 115 V 2*280 A A1φWC 115 V 22250 A x Wall outlet A1φWC 100 V 22260 A x First class grounding terminal Grounding 3 Terminal box
A2-15 Table A2.3.1 Summary of power supply (STA2) (2/2)
Install place Name Power Quantity Remarks A
A3φWC 210 V 4A1100
A3φWC 210 V 3A1100
A3φWC 210 V 3A510 P-1 distribution A1φWC 210 V 3A520 board A1φWC 210 V 3A240
A1φWC 100 V 3A1200
Spacecraft A1φWC 100 V 3A590 Preparation A3φWC 210 V 4*620 A Hall 1 P-1-1 * For distribution A3φWC 115 V 4A260 refrigerator board A1φWC 100 V 3A520
A3φWC 210 V 3A340
Wall outlet A1φWC 115 V 22240 A x
A1φWC 100 V 22240 A x
First class grounding terminal Grounding 4 Terminal box
A3φWC 210 V 4A1100
A1φWC 210 V 3A520
A3φWC 115 V 4A520
P-7 A3φWC 115 V 4A280 distribution board A1φWC 100 V 3A570
Spacecraft A1φWC 100 V 3A210 Preparation A1φWC 100 V 3A1150 Hall 2 A1φWC 100 V 3A1200
A3φWC 210 V 3A350
Wall outlet A1φWC 115 V 22270 A x
A1φWC 100 V 22280 A x
First class grounding terminal Grounding 4 Terminal box
A2-16 A2.4 Spacecraft and Fairing Assembly Building (SFA)
A2.4.1 General The SFA is used for performing some Phase 2 and Phase 3 operations for a large spacecraft to be loaded into the H-IIA launch vehicle, following Phase 1 operations. Phase 2 and Phase 3 operations performed in the SFA include: unpacking and visual inspection of the spacecraft after transportation into the SFA; transportation of the GSE used in the SFA and its installation and adjustment; high-pressure leak tests of spacecraft propellant systems; preparation of filling equipment; filling and pressurization of propellant in the spacecraft; storage of filling equipment; installation of pyrotechnics in the spacecraft; final fitting out of solid propellant motor and its installation on the spacecraft; mating of the spacecraft and payload adapter, the spacecraft and fairing; and preparation for transporting the spacecraft and fairing composite. In addition to the above spacecraft-related operations, preparation and inspection of the fairing carried into this building are also performed. The walls between the spacecraft-fairing assembly hall, filling and assembly hall, and checkout room are explosion-proof. In addition, the spacecraft fairing assembly hall, filling and assembly hall, and filling equipment room are equipped A with emergency exhaust equipment, propellant washing and draining system, gas sensors, fire alarms, body shower and eye washer, and emergency exits for coping with accidents in handling the propellant. Specifications of the first floor are as follows. a) Cleanliness : class 100,000 b) Temperature : normal operation and after encapsulation 21 ± 3 ºC propellant loading 21 ± 1 ºC battery charging 18 ± 3 ºC c) Humidity : 45 ± 5 % RH A e) Power supply : Frequency 60 Hz Panel capacity - Table A2.4.1 Location of outlets - shown in H-IIA Payload-Related Facilities and GSE Manual A
A2.4.2 Main Structure and Functions of SFA The SFA is a three-storied steel-frame building of 4,010 m2 floor space. The first floor contains air locks (1 and 2), spacecraft - fairing assembly hall, A filling and assembly hall, filling equipment room, GSE storage room, checkout room, preparation room (#1 and #2), equipment storage room, pump room,
A2-17 entrance hall, rest rooms, showers, visitor's path (entrance), high-pressure gas A storage, and the open shed at the outside. The second floor contains air-conditioning room #1 and the transformation room. The third floor section contains the visitor's path and air-conditioning room #2. Figure A2.4.1 shows the spacecraft & fairing assembly building.
A2-18 Figure A.2.4.1 Spacecraft and Fairing Assembly Building (SFA)
A2-19 Table A2.4.1 Summary of power supply (SFA)
A Installed Name Power Quantity Remarks Place A3φWC200V 3A520
A1φWC100V 2*510 A P-1 Uninterruptable distribution A1φWC100V 2*240 A *CVCF system board A1φWC115V 2*520 A
A1φWC115V 2*240 A
A3φWC200V 3A520
Checkout A1φWC100V 2*510 A room P-2 A1φWC100V 2*240 A Uninterruptable distribution *CVCF system board A1φWC115V 2*520 A
A1φWC115V 2*240 A
A3φWC200V 3A1100
WVall outlet A1φWC100 22280A x First class grounding terminal Grounding 3 Terminal box A3φWC200V 4A340 Wall outlet Spacecraft- (Explosion- A1φWC100V 3A240 1 fairing proof socket) assembly A1φWC115V 3A230 hall First class grounding terminal Grounding 12 Terminal box
A2-20 A2.5 Third Stage and Spacecraft Assembly Building (TSA) A
A2.5.1 General The TSA is used for a series of operations for a small spacecraft and dual launch spacecraft to be loaded into the H-IIA launch vehicle after delivery to the launch site. Operations include filling and pressurization of the spacecraft propellant, pyrotechnics installation, solid propellant motor installation. Specifications of the first floor section are as follows. a) Cleanliness : class 100,000 b) Temperature : 21 ~ 25 ºC c) Humidity : 40 ~ 60 % RH e) Power supply : Frequency 60 Hz Panel capacity - Table A2.5.1 Location of outlets - shown in H-IIA Payload-Related Facilities and GSE Manual (in coordinating NASDA-HDBK-T.B.D.)
A2.5.2 Main Structure and Functions of TSA The TSA is a two storied reinforced concrete building. RF test antenna (directed to STA1) is provided on the rooftop. It consists of an airlock room, checkout room, spacecraft fairing assembly C hall, two cloak room, rest room, two air shower room, comunications equipment A room, outdoor high-pressure gas cylinder storage. Figure A2.5.1 shows the TSA.
A2-20-1A2-21 A
Figure A.2.5.1 Third Stage and Spacecraft Assembly Building (TSA)
A2-20-2A2-22 Table A2.5.1 Summary of power supply (TSA) A
Install place Name Power Quantity Remarks
A3φWC 200 V 4A370 Explosionproof Wall outlet spacecraft - fairing A1φWC 100 V 3A340 socket assembly hall First class grounding terminal Graounding 3 Terminal box
A3φWC 200 V 3A310 L-2 distribution A1φWC 100 V* 3A240 * Can switch to 200 V boad Checkout room A1φWC 200 V 2A210 A1φWC 100 V 22195 A x Wall outlet A3φWC 200 V 3A310
A2-20-3A2-23 A2.6 Yoshinobu Vehicle Assembly Building (VAB) A
A2.6.1 General A The VAB is a facility for assembling, inspecting and adjusting all components of the H-IIA vehicle (the first stage, the second stage, the SRB-A) transported from each factory. And the mating of the spacecraft/fairing composite onto the launch vehicle and final checkout of the spacecraft and the launch vehicle are also conducted in this building prior to the launch day. The spacecraft / fairing composite is carried into the VAB through the spacecraft carrying door loaded on the transportation dolly, and hoisted on the top of the second stage using the crane equipped in the VAB. To maintain environments inside the VAB, the air conditioning equipment is fully furnished. Especially, upper stories of the VAB, in which spacecraft operations are performed, are maintained in a cleanliness of the class 100,000. Environments inside the fairing are controlled at 10 – 25 ± 2 °C by supplying the air through the air conditioning duct. Cleanliness of this air is class 5,000. The GSE of the spacecraft is set in the pressurized room inside the ML, and control and monitor of the spacecraft is conducted by remote control from the STA2. On the launch day, the slide door is opened, and then the ML on which the spacecraft / vehicle is loaded is transferred to the LP. During transfer, conditioned air is supplied to the spacecraft. Figure A2.6.1 and Figure A2.6.2 show a general view and a sectional view of the A VAB respectively. The VAB has two pressurization rooms respectively at 9th, 10th and 11th A floor. Specifications of power supply for spacecraft operations in pressurization rooms at 9th and 10th floors are as follows. a) Frequency_60 Hz b) Panel capacity_Table A2.6.1 A Location of outlets - shown in H-IIA Payload-Related Facilities and GSE Manual (in coordinating NASDA-HDBK-T.B.D.) A
A2.6.2 Main Structure and Function of VAB A The elevator type floor is equipped in upper stories to access the spacecraft; it is adjustable to required heights. Each level is equipped with AGE storage space, air-conditioning equipment and so on. And for safety and accident prevention during operations, temperature sensors, smoke sensors, fire alarms, communications devices, announcement systems, face / body showers, and evacuation escape routes are provided at required places on each floor.
A2-24A2-21 B
Figure A2.6.1 Yoshinobu Vehicle Assembly Building (VAB) A
A2-25A2-22 RF
15FL
M15FL
14FL
M14FL
13FL
12FL
11FL
10FL
9FL
M9FL
8FL
7FL
6FL
M6FL
5FL
4FL
3FL
2FL
M2FL ML ML
GL
VAB1 VAB2 A
Figure A2.6.2 Yoshinobu Vehicle Assembly Building (VAB, sectional view) A
A2-26A2-23 Table A2.6.1 Summary of power supply (VAB 9th and 10th floor) A
A Install place Name Power Quantity Remarks
AAC210V 63φW0 41
AAC220V 31φW0 21 Spacecraft AGE Payload Power supply AAC220V 21φW0 21Uninterruptable Checkout equipment *CVCF system Room 1 AAC210V 31φW0 21 (VAB1) AAC115V 21φW0 22
AAC110V 21φW0 22
AAC210V 63φW0 41
AAC220V 31φW0 21 Spacecraft AGE Payload Power supply AAC220V 21φW0 21Uninterruptable Checkout equipment *CVCF system Room 2 AAC210V 31φW0 21 (VAB2) AAC115V 21φW0 22
AAC110V 21φW0 22
A2-27A2-24 A2.7 Movable Launcher (ML) A
A2.7.1 General A The ML is a movable facility, on which the spacecraft and the launch vehicle are loaded in the VAB, and which carries them to the launch pad on the launch day and functions as the launch table at the launch pad as it is. The GSE of the spacecraft is set in the pressurized room inside the ML and has the interface with the spacecraft through the umbilical line equipped in the umbilical mast of the ML. Figure A2.7.1 shows a general view of the ML. A
A2.7.2 Main structure and functions of ML A (1) Launch table The H-IIA launch vehicle is set on the launch settling table of the launch table. The first and second stages of the launch vehicle are loaded with propellants though the launch settling table and the umbilical mast respectively. For the spacecraft inside the payload fairing, electrical signals and power are sent through the umbilical cable and conditioned air is supplied through the air conditioning duct. The umbilical cable and the air conditioning duct are equipped in the umbilical mast on the launch table.
(2) Dolly (Movable Launcher Transporter) A The dolly is equipment which carries the launch table from the VAB to the C launch pad. The dolly has wheels and can move anywhere and rotate.
(3) Pressurization room The pressurization room is inside the ML and the GSE of the spacecraft is set in this room. Specifications of this room are as follows.
a) Area : 20 m2 (per each spacecraft) b) Temperature : 18 - 28 °C c) Humidity : 40 - 60 % RH d) Power supply : Frequency_50 Hz or 60 Hz C Panel capacity_Table A2.7.1 A Location of outlets - shown in H-IIA Payload-Related Facilities and GSE Manual (in coordinating NASDA-HDBK-T.B.D.) A
A2-28A2-25 B
Figure A2.7.1 Movable Launcher (ML) A
A2-29A2-26 Table A2.7.1 Summary of power supply (ML) A
A Place Power Quantity Remarks
AAC100V 51φW0 2z610 H
AAC100V 21φW0 2z620 H
AAC115V 51φW0 2z610 H
AAC115V 21φW0 2z620 H
AAC210V 51φW0 2z620 H ML Payload AAC208V 53φW0 4z610 H GSE Room 1 AAC208V 53φW0 4z610 H
AAC220V 31φW0 2z620 H
AAC220V 51φW0 2z610 H
AAC220V 11φW0 2z610 H
AAC220V 31φW0 2z520 H
AAC100V 51φW0 2z610 H
AAC100V 21φW0 2z620 H
AAC115V 51φW0 2z610 H
AAC115V 21φW0 2z620 H
AAC210V 51φW0 2z620 H
AAC208V 53φW0 4z610 H ML Payload GSE Room 2 AAC208V 53φW0 4z610 H
AAC220V 31φW0 2z620 H
AAC220V 51φW0 2z610 H
AAC220V 11φW0 2z610 H
AAC220V 31φW0 2z520 H
For Piggy-buck AAC100V 21φW0 2z610 H spacecraft
A2-30A2-27 A2.8 Takesaki Range Control Center (RCC) A
A2.8.1 General A The Takesaki RCC is responsible for control and planning of overall launch operations, communication with each launch site station, ground safety, flight safety, and weather observation. For these operations, the center is equipped with a range control system, flight safety processing system, flight safety control system, communication system (clock system, operational intercommunication telephone system (OIS), and optical transmission system), and weather observation system. In addition, the third floor is equipped with optical equipment (No. 2 optical observatory station) for tracking the rocket after lift-off.
A2.8.2 Main structure and functions of Takesaki RCC A Takesaki RCC is housed in a three-story reinforced concrete building. Its floor area is about 1,400 m2. The first floor of the building comprises the control room, communication A and computer room, flight safety command room, planning and combination room, meeting room, visitors observation room, and office rooms. The second floor contains the weather observation machine room, weather observation operations room, and office rooms. The third floor contains the optical observation room and observation camera room. Figure A2.8.1 shows the Takesaki RCC. A
A2-31A2-28 Figure A2.8.1 Takesaki Range Control Center (RCC) A
A2-29A2-32 A2.9 Down range stations A
A2.9.1 Ogasawara down range station A The Ogasawara down range station is located on Chichijima Island of Ogasawara village, Tokyo. The launch vehicle telemetry receiver facility, flight A safety command transmitter facility, and tracking radar facility are installed there. Figure A2.9.1 shows the Ogasawara down range station. A
(1) Objective of Ogasawara down range station The major objective of the Ogasawara down range station is to take over the tracking of the vehicle launched at Tanegashima outside the visible range of Tanegashima. For the H-IIA, it monitors the flight conditions during the period from combustion in the second stage to orbit injection using the vehicle telemetry and confirms the vehicle flight position by the tracking radar.
(2) Interface with Tanegashima The vehicle orbit prediction information (elevation angle, azimuth and line-of- sight distance) for the antenna and radar tracking system required to capture the vehicle is calculated in real time by the flight safety processing system at the Tanegashima Range Control Center Building, based on the vehicle tracking measurement data obtained at the Tanegashima Launch Site, then supplied to the Ogasawara down range station. The collected radar tracking data and specific vehicle telemetry data are transmitted to Tanegashima in real time. The data transmission interface between Tanegashima and Ogasawara is ensured by using telecommunication satellites and domestic leased lines.
A2.9.2 Christmas down range station A The Christmas down range station is located at 157 deg. W. longitude and 2 deg. N. latitude on Christmas Island of the Republic of Kiribati about 1,000 km south of Hawaii. The launch vehicle telemetry receiver facility, INMARSAT earth station facilities and power facilities are installed there. Figure A2.9.2 shows the Christmas down range station. A
(1) Objective The major objective of the Christmas down range station is to monitor reignition of the vehicle second stage engine (SEIG #2) and to confirm separation between the vehicle and spacecraft using telemetry. The received data are recorded and sent to Tanegashima via the INMARSAT line after tracking.
A2-33A2-30 (2) Initial capturing of launch vehicle The vehicle orbit prediction information (elevation angle and azimuth) for the antenna required to capture the vehicle is supplied from Tanegashima in real time and is based on the measurements of the vehicle from the Ogasawara down range station.
A2-34A2-31 Figure A2.9.1 Ogasawara down range station A
Figure A2.9.2 Christmas down range station A
A2-35A2-32 A P P E N D IX 3 .
P A Y L O A D A D A P T E R
P
A A Y
P
L
P O A
E D
N
A
D D
I A
X P
T
3 E
. R APPENDIX 3. PAYLOAD ADAPTER
11 types of payload adapter are offered. C These are; a) Model 937M b) Model 937MH c) Model 937M-SPIN d) Model 937M-SPIN-A e) Model 1194M f) Model 1666M g) Model 1666S h) Model 2360S i) Model 3470S j) Model 1666MA C k) Model 239M In the following sections, more information on these adapters is provided.
A3-1 APPENDIX 3.1 937M ADAPTER A
The main characteristics are as follows. C (1) Interface diameter : 945 mm A (2) Height : 900 mm (3) Material a. Cone : co-cured graphite epoxy b. Adapter ring : Aluminum B (4) Attached system : Clamp bands (5) Separation system : 4 springs (6) Clamp band A Maximum tension : 24.1 kN (7) Maximum load per spring : 1670 N (8) Adapter mass : 76 kg
Figure A3.1.1 shows the photograph of the 937M adapter. C Figure A3.1.2 shows a general view of the 937M adapter . Figure A3.1.3 shows details of the 937M adapter. Figure A3.1.4 shows stay-out zone around the 937M adapter. A Figure A3.1.5 shows the limit load of the 937M adapter. Figure A3.1.6 shows the spacecraft separation shock spectrum with the 937M adapter. Figure A3.1.7 shows the limit load at separation plane of the 937M adapter.
C
Figure A3.1.1 Photograph of the 937M adapter
A3.1-1 (0°) IV A Spacecraft Separation Spring (4 pl) θ = 45.0° θ = 315.0°
θ = 67.5°
.5
825 φ 2 φ1219. (90°) ° I III (270 )
Spacecraft Umbilical Connector (2 pl) ° θ = 135.0 θ = 225.0°
° II (180 ) Launch Vehicle Separation Switch (2pl, Inside of separation spring)
φ945.26
Al ring C 211 structure
Co-cured graphite 900 epoxy A cone structure
φ2190 (PCD) Unit : mm B
Figure A3.1.2 General view of the 937M adapter CA
A3.1-2 A
φ945.26 ± 0.076
H 0.05 Spacecraft frame 0.33 A
R1 ~ 2.3 Detail A
± ° φ876.30 0.25 0
3 0.076
± ' 5 1 ± ° 5.84 5 Spacecraft 1
3.56 Separation R1 ~ 2.3 Plane 3.2 1.6 1.6
+0.13 0.254 -0 (2 pl)
.13 0 + -0
+0.13 R0.13 3.56 -0 +0.13 φ940.94 -0 9.53
+0.15 φ945.26 0
0 φ939.97 -0.2
0 2.65 -0.1 9.5 ' 5 C0.2 2pl 1 ± 0.076 ° 0.03 R0.13 2pl ± 63 ± 5 63 1 φ891.50 5.84 63 1.53
C1
° 1
5 R2.3 5.4 Payload adapter
Unit : mm
Figure A3.1.3 Details of the 937M adapter C
A3.1-3 A
φ3650
φ3280
φ3144 21°
φ1908
φ1398
φ1270
φ940
φ724 ° 15 350
12° 50.8 Spacecraft Separation Plane 120 210 380
° 0 3
R1050
Stay-out zone
Figure A3.1.4 Stay-out zone around the 937M adapter C
A3.1-4 A
4000
3600 3500
3000
2500
2000
1500 Allowable range
Spacecraft mass (kg) 1000
500
0 0.0 1.01.3 2.0 3.0
Spacecraft C.G. height from separation plane (m)
Figure A3.1.5 Limit load of the 937M adapter C
A3.1-5 10000 CA
1500Hz 3000Hz
3000G
1000 Acceleration (G) 100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 (m/s2) B
Figure A3.1.6 Spacecraft separation shock spectrum with the 937M adapter CA
A3.1-6 B
60
47.36 40
20
5.94 0 Allowable area
Axial load (ton) -20
-35.48 -40 Tension Compression -60 024689.73 10 Bending moment (ton¥m)
Figure A3.1.7 Limit loads at separation plane of the 937M adapter C
A3.1-7 A APPENDIX 3.2 937MH ADAPTER
The main characteristics are as follows. C (1) Interface diameter : 958 mm (2) Height : 900 mm (3) Material B a. Cone : co-cured graphite epoxy A b. Adapter ring : Aluminum (4) Attached system : Clamp bands B (5) Separation system : 4 springs A (6) Clamp band Maximum tension : 14.9 kN (7) Maximum load per spring : 1670 N (8) Adapter mass : 80 kg
Figure A3.2.1 shows the photograph of the 937MH adapter. CA Figure A3.2.2 shows a general view of the 937MH adapter . Figure A3.2.3 shows details of the 937MH adapter. Figure A3.2.4 shows stay-out zone around the 937MH adapter. Figure A3.2.5 shows the limit load of the 937MH adapter. Figure A3.2.6 shows the spacecraft separation shock spectrum with the 937MH adapter. B Figure A3.2.7 shows the limit load at separation plane of the 937MH adapter.
C
Figure A3.2.1 Photograph of the 937MH adapter
A3.2-1 ° IV (0.0 ) CA θ = 345.5° Spacecraft � Separation Spring � (4 pl)
θ = 75.5.0°
25.5 8 φ1219.2 φ
(90.0°) ° I III (270.0 )
θ = 255.5°
Key way Spacecraft� Umbilical Connector � (2 pl) θ = 135.0° Launch Vehicle� C Separation Switch� A (2 pl) ° θ = 165.5 ° II (180.0 )
φ958.85 BC A
Al ring C
211 structure B A
Co-cured graphite 900 epoxy cone structure
φ2190 (PCD) Unit : mm A
Figure A3.2.2 General view of the 937MH adapter C
A3.2-2 939.800 φ A
Spacecraft separation plane
R0.254 R0.127
3.302 3.150 1.905 1.778 R2.286 11.049
2 0 ° ± Spacecraft frame 0 .2 5°
4 7 ° 0.381 0.254 φ925.592
3.683 φ937.997 3.556 φ937.870 φ955.294 φ955.167 φ958.926 φ958.774
φ958.05 ± 0.08
+0 φ954.26 - 0.1 0
+0 φ937.62 - 0.13 +0 2.65 - 0.1 φ899 ± 0.25 ° 5 .2 C0.2 (TYP) +0.13 0 - 0 ± ° 4 5 0 R0.13 (TYP) ° 2
0.03 ± ± 0 . 5
° 1.53
R0.5 0.25 0.25 ±
± B ± 0.25 +0.05 - 0 C 0.07 0.25 ± ± 0 2.59
R2.28 1.60
3.22 1.60 1.80 ± 0.25 (const) A Payload adapter
φ944 (REF)
Unit : mm Figure A3.2.3 Details of the 937MH adapter C
A3.2-3 A
φ1880
φ1560
φ940
φ595.4 Spacecraft Rear 50.8 Frame Spacecraft Separation Plane 258 325 Payload Adapter 500
Stay-out zone
Unit : mm
Figure A3.2.4 Stay-out zone around the 937MH adapter C
A3.2-4 A
2500
2000
1700 1500
1000 Spacecraft mass (kg) 500 Allowable range
0 0.0 1.0 2.0 3.0
Spacecraft C.G. height from separation plane (m)
Figure A3.2.5 Limit load of the 937MH adapter C
A3.2-5 A
10000 C 1500Hz 3000Hz
4100G
1000 Acceleration (G) 100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2 B A
Figure A3.2.6 Spacecraft separation shock spectrum with the 937MH adapter C
A3.2-6 B
20
10 8.46
2.80 0 Allowable area Axial load (ton)
-10 -12.85 Tension Compression -20 0.0 1.0 2.0 3.03.68 4.0 5.0 Bending moment (ton¥m)
Figure A3.2.7 Limit loads at separation plane of the 937MH adapter C
A3.2-7 APPENDIX 3.3 937M-SPIN ADAPTER A
The main characteristics are as follows. C (1) Interface diameter : 959 mm (2) Height : 605 mm (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 4 springs (6) Clamp band Maximum tension : 13.5 kN (7) Maximum load per spring : 1670 N B (8) Adapter mass : 130 kg The plume shield can be installed if necessary. The adapter has a spin-table section, which can spin the spacecraft up to 100*1 rpm with an angular acceleration of less than 10 rad/sec2. Figure A3.3.1 shows the photograph of the 937M-SPIN adapter. A Figure A3.3.2 shows a general view of the 937M-SPIN adapter. Figures A3.3.3 and A3.3.4 show details of the 937M-SPIN adapter. Figure A3.3.5 shows the stay-out zone around the 937M-SPIN adapter. Figure A3.3.6 shows the limit load of the 937M-SPIN adapter. Figure A3.3.7 shows the spacecraft separation shock spectrum of the 937M- SPIN adapter.
Figure A3.3.1 Photograph of the 937M-SPIN adapter A
A3.3-1 Spin rocket motor
Plume shield
φ 1920
2600 φ
Separation spring (4 pl) φ958.85 Clamp band S/C Separation plane 605
φ2190 (PCD) φ2220
Unit : mm
Figure A3.3.2 General view of 937M-SPIN adapter A
A3.3-2 B IV CB A
58.5 4 pl B 90°
45°
840 φ φ825.5 I III
S / C separation monitor II Switch pad (4 pl)
Unit : mm
Figure A3.3.3 Details of 937M-SPIN adapter #1 A
A3.3-3 ° ° + 0 15' R0.38 60 0° R0.13
30° +0.08 -0 ±0°30' 5.54 8.74
17.48 ± 0.25 +0.08 - 0
φ958.85 (REF)
Detail A
φ958.85±0.08 0 φ937.62 -0.13 φ899±0.25
45°±0°30' 30° 0°30'
± 7.11 Separation plane 0.03 0.07 0.25 R3 (TYP) ± ± 1.60 2.59
Separation switch
Section Ｂ Unit : mm
Figure A3.3.4 Details of 937M-SPIN adapter #2 A
A3.3-4 C φ940
φ632 Spacecraft 50.8 Spacecraft Rear Separation Frame Plane
φ406 101
Payload Adapter 203
Stay-out zone
Unit : mm
Figure A3.3.5 Stay-out zone around the 937M-SPIN adapter A
A3.3-5 2000 C
1800
1600
1400
1200
1000
800
Spacecraft mass (kg) 600 Allowable range
400
200
0 1.060 1.737 0.0 1.0 2.03.0 4.0
Spacecraft C.G. height from separation plane (m)
Figure A3.3.6 Limit load of the 937M-SPIN adapter A
A3.3-6 10000 C
1500Hz 3000Hz
2000G
1000 Acceleration (G) 100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2
Figure A3.3.7 Spacecraft separation shock spectrum of the 937M-SPIN adapter A
A3.3-7 A APPENDIX 3.4 937M-SPIN-A ADAPTER
The main characteristics are as follows. C (1) Interface diameter : 958 mm B (2) Height : 1,000 mm A (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 4 springs (6) Clamp band Maximum tension : 14.9 kN B (7) Maximum load per spring : 1670 N A (8) Adapter mass : 150 kg
The plume shield can be installed if necessary. A The adapter has a spin-table section, which can spin the spacecraft up to 50*1 C rpm with an angular acceleration of less than 5 rad / sec2. A C Figure A3.4.1 shows the photograph of the 937M-SPIN-A adapter. A Figure A3.4.2 shows the general view of the 937M-SPIN-A adapter. Figure A3.4.3 shows the details of the 937M-SPIN-A adapter. Figure A3.4.4 shows the stay-out zone around the 937M-SPIN-A adapter. Figure A3.4.5 shows the limit load of the 937M-SPIN-A adapter. Figure A3.4.6 shows the spacecraft separation shock spectrum of the 937M- SPIN-A adapter. B Figure A3.4.7 shows the limit load at separation plane of the 937M-SPIN-A adapter.
Figure A3.4.1 Photograph of the 937M-SPIN-A adapter
A3.4-1 ° IV (0.0 ) A θ = 345.5° C Spacecraft Separation Spring (4 pl)
θ = 75.5° .5 9.2 5 21 φ1 82 φ
(90.0°) ° I III (270.0 )
θ = 255.5°
Key way
Spacecraft Umbilical Connector (2 pl) θ = 135.0° C Launch Vehicle Separation Switch A (2 pl) ° θ = 165.5 ° II (180.0 )
φ1160
φ958.85 BC A
211 C A 1000 258
φ2190 (PCD) C
Unit : mm Figure A3.4.2 General view of 937M-SPIN-A adapter A
A3.4-2 φ939.800 A
Spacecraft separation plane
R0.254 R0.127
3.302 3.150 1.905 1.778 R2.286 11.049
2 0 ° ± Spacecraft frame 0 .2 5° B
4 7 A ° 0.381 0.254 φ925.592
3.683 φ937.997 3.556 φ937.870 φ955.294 φ955.167 φ958.926 φ958.774
φ958.05 ± 0.08
+0 φ954.26 - 0.1 0
+0 φ937.62 - 0.13 +0 2.65 - 0.1 φ899 ± 0.25 ° 5 .2 C0.2 (TYP) +0.13 0 - 0 ± ° 4 5 0 R0.13 (TYP) ° 2
0.03 ± ± 0 . 5
° 1.53
R0.5 0.25 0.25 ±
± CB ± 0.25 +0.05 - 0 0.07 0.25 ± ± 0 2.59
R2.28 1.60
3.22 1.60 1.80 ± 0.25 (const) Payload adapter A
φ944 (REF)
Unit : mm Figure A3.4.3 Details of 937M-SPIN-A adapter C
A3.4-3 A
φ1880
φ1560
φ940
φ595.4 Spacecraft Rear 50.8 Frame Spacecraft Separation Plane 258 325 Payload Adapter 500
Stay-out zone
Unit : mm
Figure A3.4.4 Stay-out zone around the 937M-SPIN-A adapter C
A3.4-4 A
2500
2000
1700 1500
1000 Spacecraft mass (kg) 500 Allowable range
0 0.0 1.0 2.0 3.0
Spacecraft C.G. height from separation plane (m)
Figure A3.4.5 Limit load of the 937M-SPIN-A adapter C
A3.4-5 A
10000 C 1500Hz 3000Hz
4100G
1000 Acceleration (G) 100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2 B
Figure A3.4.6 Spacecraft separation shock spectrum of the 937M-SPIN-A adapter CA
A3.4-6 B
20
10 8.46
2.80 0 Allowable area Axial load (ton)
-10 -12.85 Tension Compression -20 0.0 1.0 2.0 3.03.68 4.0 5.0 Bending moment (ton¥m)
Figure A3.4.7 Limit loads at separation plane of the 937M-SPIN-A adapter C
A3.4-7 APPENDIX 3.5 1194M ADAPTER A
The main characteristics are as follows. C (1) Interface diameter : 1,215 mm (2) Height : 480 mm (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 4 or 8 springs (6) Clamp band BA Maximum tension : 36.8 kN (7) Maximum load per spring : 1670 N (8) Adapter mass : 100 kg
This adapter has two ventholes of 45 mmø (1590.4 mm2) assuming that the BA internal volume of the spacecraft is less than 2 m3. When interface connectors are installed on the separation plane, connectors are located 789.125 mm from the center of the vehicle axis. A C Figure A3.5.1 shows the photograph of the 1194M adapter. A Figure A3.5.2 shows the general view of the 1194M adapter. Figures A3.5.3 to A3.5.5 show the details of the 1194M adapter. Figure A3.5.6 to A3.5.7 show the stay-out zone around the 1194M adapter. Figure A3.5.8 shows the limit load of the 1194M adapter. Figure A3.5.9 shows the spacecraft separation shock spectrum of the 1194M adapter. Figure A3.5.10 shows the limit loads at separation plane of the 1194M adapter. B
Figure A3.5.1 Photograph of the 1194M adapter A
A3.5-1 I
MICRO SWITCH (2 pl)
F
° 22.5° A 22.5 F ° A 45
22.5
° II IV °
22.5
SPRING HOUSING (8 pl)
SPACECRAFT III SEPARATION PLANE
//0.5 A 0.3 0.03/10 φ1215 (REF) 480
φ2190 (PCD) A
Unit : mm
Figure A3.5.2 General view of 1194M adapter A
A3.5-2 φ1161 A
ML MICRO SWITCH SECTION FÐF� �
SPACECRAFT SEPARATION PLANE C
B
SPRING HOUSING
ML
SECTION AÐA� � Unit : mm
Figure A3.5.3 Details of 1194M adapter #1 (Details of separation plane) A
A3.5-3 φ1211.2 ± 0.15 0.3 φ≦1194 CA φ≦1132
¡ R6 ¡ 5 R 4.5 0 .2 IXX 0 - ¡ 5 10 1 IYY ≧ L (See Note) e 1.6 1.6 D 1184 ± 0.5 φ 0.125 C 5.72 φ1215 ± 0.5 Spacecraft frame 1.3 ± 0.3 0.3
A
1215 ± 0.15 0.3 φ 0.05 A 1184.3 ± 0.15 0.03/10 φ 0.1 B 0.3 φ1161 A E 1.6
IXX
IYY L (See Note)
φ1196 ± 0.5 Adapter
Detail B Unit : mm
Figure A3.5.4 Details of 1194M adapter #2 (Cross section of frames) A
A3.5-4 0.125 C A C ° °± 0.5 10 + 0.12 φ1209.52 0 3.07 ± 0.03 0.05 ± 0.76 1.52 ¥ R 0.3 C 0.2 C 0.4 Spacecraft 0.5×0.5 One engraved� mark on one spacecraft lateral axis Detail D
0 1209.17 - 0.13 B φ 2.54 ± 0.03
R 0.5 ± 0.13 C 0.2 A (TYP) (TYP) C 0.2
1 5 °
-
0 0 0.03
°
. ±
2 5.72
5 ° 1.27
R 0.5 1.6
φ1211.2 ± 0.15 0.3
Detail E Unit : mm
Figure A3.5.5 Details of 1194M adapter #3 (Details of frames) A
A3.5-5 A
[I] φ3700 mm ±20°
φ1815 mm [II] [IV]
φ1215 mm ±20° [III]
φ1194 mm Spacecraft rear frame φ1000 mm Usable volume φ820 mm
φ134 mm 200 mm 100 mm 60 mm
Spacecraft 46° ° Separation 46 110 mm Plane Stay out zone 450 mm 336 mm 480 mm
15° (REF)
This area must be coordinated with NASDA.
Figure A3.5.6 Stay-out zone around the 1194M adapter (I/III)
A3.5-6 A
φ3700 mm
[I] φ1815 mm ±70°
φ1215 mm [II] [IV] Spacecraft 1194 mm ±70° φ rear frame [III] φ1000 mm Usable volume φ820 mm � φ134 mm 200 mm 60 mm
Spacecraft
separation 110 mm plane Stay out zone 46° 450 mm 336 mm 480 mm
15° 12° (REF)
This area must be coordinated with NASDA.
Figure A3.5.7 Stay-out zone around the 1194M adapter (II/IV)
A3.5-7 A
5000
4500
4000
3500
3000
2500 Allowable range 2000 Spacecraft mass (kg) 1500
1000
500
0 0.0 1.0 2.0 3.0 4.0 5.0 Spacecraft C.G. height from separation plane (m)
Figure A3.5.8 Limit load of the 1194M adapter
A3.5-8 A
10000 1500Hz 3000Hz
4100G
1000 Acceleration (G)
100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2 B
Figure A3.5.9 Spacecraft separation shock spectrum of the 1194M adapter A
A3.5-9 B
100
70.20
50
7.59 0 Allowable area Axial load (ton)
-50 -55.02 Tension Compression -100 0 4 8 12 1618.69 20
Bending moment (ton¥m)
Figure A3.5.10 Limit loads at separation plane of the 1194M adapter
A3.5-10 APPENDIX 3.6 1666M ADAPTER A A
The main characteristics are as follows. C (1) Interface diameter : 1,666 mm (2) Height : 1,000 mm (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 4 or 6 springs (6) Clamp band BA Maximum tension : 36.3 kN A (7) Maximum load per spring : 1670 N (8) Adapter mass : 100 kg
2 This adapter has 12 vent holes of 83 mm φ (5410.6 mm ) assuming that the BA internal volume of the spacecraft is less than 2 m3. When interface connectors are installed on the separation plane, connectors are located 942.45 mm from the center of the vehicle axis. A C Figure A3.6.1 shows the general view of the 1666M adapter. A Figure A3.6.2 to A3.6.4 show details of the 1666M adapter. Figure A3.6.5 shows the stay-out zone around the 1666M adapter. Figure A3.6.6 shows the limit load of the 1666M adapter. Figure A3.6.7 shows the spacecraft separation shock spectrum of the 1666M C adapter. A Figure A3.6.8 shows the limit load at separation plane of the 1666M adapter. B
A3.6-1 A +X n locatio (S/C) pring S ¡ 337.3 ¡ 247.3 ¡ 202.3 ¡ 157.3 67.3¡ 22.3¡
Spring housing (4 or 6 pl)
Micro switch (2 pl) 6 7. 3¡ ¡ .3 2 1 1
-Y +Y φ160 (S/C) 0 (S/C)
-X (S/C)
Unit : mm 0.5 A
φ 1666 (REF) 0.3 CB A
C A 1000
φ2190 (PCD) A
Figure A3.6.1 General view of the 1666M adapter
A3.6-2 φ1588 A
Micro switch
φ1600 (REF)
Spring housing
Unit : mm
Figure A3.6.2 Details of the 1666M adapter #1
A3.6-3 φ1666.1 ± 0.1 0.3 A 0 φ1644.2 - 0.5 B φ1626 ± 0.2 0.3 A R2 5.0
3.0 0.03/10
B R3 1666.1 ± 0.1 (REF) R6 φ 0.1 8.0 ± 0
67.2 R0.2(TYP) φ1548.0 φ1644.2 - 0.5 (REF)
3.0 1 5° φ1633.3 0.07 ±
± 0 ° ° 1 3 1.6 5 4 ' φ1663.6 1.6 3.22
SECT A-A 0.5 R2.28 R3 (REF) 1.6 0.03 ±
0 ± 5 0.2 R1. 5 (TYP) 9.02
DETAIL B 0.8 0.025 3 A +X φ0.33 3 B 3 A (S/C) 13° + 0.08 50 (26 ' 46 9.19 - 0.05 pl) ' DETAIL C
± 0.1 (REF)
1666.1 φ A A C -Y -Y (S/C) (S/C)
Unit : mm -X (S/C) Figure A3.6.3 Details of the 1666M adapter #2
A3.6-4 Y A
X ≤ φ1642.3
R6.3 2.2
≥ 0.3 e 4.6 R A 0.03/10 2.29 3.23 R 0.05 B R3 1.6
° 1.6 6.35 7 4 ' 5 1 10.4 ° 0 ° ± φ1666.1 ± 0.1 (REF) B 15
SECT A-A
+0.08
9.19 - 0.05
03
.
0 ±
5 2 1. R P)
Y 9.0 (T
0.8 0.025 3 A
φ0.33 3 B 3 A
"
6 4
'
0
5
°
3
1
Stiffness : A A S = T.B.D. mm2 Ixx ≧ T.B.D. mm4 ± T.B.D. % Iyy ≧ T.B.D. mm4 } C Applicable length = T.B.D. mm A
φ 1666 ± 1 B A
Unit : mm
Figure A3.6.4 Details of the 1666M adapter #3 (Spacecraft frame)
A3.6-5 1396 1396 BA 1033 1033
A
A 300 300 300 300
Unit : mm
3250 1396 (REF) 1935 1033 (REF) 1820 Separation plane 27 32 495 100 200 200 100 4 5 45 775
1000 1205 1375
1776 1892 : Acceptable spacecraft envelope 2100 : Subject to discuttion : Stay-out zone SECT A-A
Figure A3.6.5 Stay-out zone around the 1666M adapter A
A3.6-6 A
5300
5000
4500
4000
3500
3000
2500 Allowable range
2000 Spacecraft mass (kg) 1500
1000
500
0 0.0 1.0 2.02.4 3.0 4.0
Spacecraft C.G. height from separation plane (m)
Figure A3.6.6 Limit load of the 1666M adapter
A3.6-7 A
Figure A3.6.7 Spacecraft separation shock spectrum of the 1666M adapter
A3.6-8 B
100
71.93
50
8.91 0 Allowable area Axial load (ton)
-50 -54.44 Tension Compression -100 0.0 5.0 10.0 15.0 20.0 25.026.81 30.0
Bending moment (ton¥m)
Figure A3.6.8 Limit loads at separation plane of the 1666M adapter
A3.6-9 A APPENDIX 3.7 1666S ADAPTER
The main characteristics are as follows. C (1) Interface diameter : 1,664 mm (PCD) (2) Height : 1,000 mm (3) Material : Aluminum semi-monocoque (4) Attached system : 4 places 19.1 mm (3/4") Bolts and Separation nuts C (5) Separation system : 4 springs A (6) Separation nuts B Maximum tension : 178 kN / Bolt A (7) Maximum load per spring : 2390 N (8) Adapter mass : 100 kg
This adapter has 8 vent holes (100 x 155 mm 4 places, 50 x 100 mm x 4 places) B assuming that the internal volume of the spacecraft is less than 2 m3. When interface connectors are installed on the separation plane, connectors A are located 928.9 mm from the center of the vehicle axis. C Figure A3.7.1 shows the general view of the 1666S adapter. A Figure A3.7.2 to A3.7.4 show details of the 1666S adapter. Figure A3.7.5 shows the stay-out zone around the 1666S adapter. Figure A3.7.6 shows the limit load of the 1666S adapter. Figure A3.7.7 shows the spacecraft separation shock spectrum of the 1666S adapter. Figure A3.7.8 shows the limit load at separation plane of the 1666S adapter.
A3.7-1 +X A Spring housing (S/C) 64.21 (4 pl)
B
A
1500
4 5 +Y -Y
(S/C) (S/C) 4 5 1176.41 (TYP) (REF) 588.21
Micro switch
178.79 B (2 pl)
A
588.21
1176.41 (TYP) (REF) -X (S/C) Unit : mm 1638.3 (REF)
C
1000 A
2190 (PCD) A
Figure A3.7.1 General view of the 1666S adapter
A3.7-2 A
φ1758
φ1663.7 (PCD)
Spacacraft separation plane
Separation nut (4 pl) 1000 (REF)
Unit : mm
Figure A3.7.2 Details of the 1666S adapter #1
A3.7-3 A
φ1663.7 (REF)
+ 0.0 φ34.04 Ð 0.04 + 0.0 Ð 0.05 ± 0.25 φ22.28 0.5 ± 11 2.67
0.01
15°
Separation nut
Unit : mm
Figure A3.7.3 Details of the 1666S adapter #2
A3.7-4 B
2138
1900
Separation plane 448 495
1030
1200
Stay-out Zone
Unit : mm
Figure A3.7.4 Stay-out zone around the 1666S adapter #1
A3.7-5 B
1176.41 (TYP) (REF) 1176.41
0
0
.
5 4
458.98 588.21 (REF)
1140.36 114.3 139.7
Separation plane
Unit : mm
DETAIL of Spacecraft corner fitting
: Stay-out zone
Figure A3.7.5 Stay-out zone around the 1666S adapter #2
A3.7-6 A
5300
5000
4500
4000
3500
3000
2500 Allowable range
2000 Spacecraft mass (kg) 1500
1000
500
0 0.0 1.0 2.02.4 3.0 4.0
Spacecraft C.G. height from separation plane (m)
Figure A3.7.6 Limit load of the 1666S adapter
A3.7-7 A
Figure A3.7.7 Spacecraft separation shock spectrum of the 1666S adapter
A3.7-8 B
100
71.93
50
8.91 0 Allowable area Axial load (ton)
-50 -54.44 Tension Compression -100 0.0 5.0 10.0 15.0 20.0 25.026.81 30.0
Bending moment (ton¥m)
Figure A3.7.8 Limit loads at separation plane of the 1666S adapter
A3.7-9 APPENDIX 3.8 2360S ADAPTER A
The main characteristics are as follows. C (1) Interface diameter : 2,360 mm (PCD) (2) Height : 300 mm (3) Material : co-cured graphite epoxy (4) Attached system : 8 places 15.9 mm (5/8") Bolts and Separation nuts B (5) Separation system : None (6) Separation nuts Maximum tension : 118 kN / Bolt (7) Maximum load per spring : None (8) Adapter mass : 40 kg
Figure A3.8.1 shows the photograph of the 2360S adapter. A Figure A3.8.2 shows a general view of the 2360S adapter. Figures A3.8.3 to A3.8.5 show details of the 2360S adapter.
Figure A3.8.1 Photograph of the 2360S adapter A
A3.8-1 1180 * 2420 (REF) 834
AA 834 1180
Separation switch (2 pl) (180°opposite direction) 834 834 1180 1180
*φ2420 φ2360 (PCD) 300.0 294.0
φ2190 (PCD) *φ2220
* Outside diameter Unit : mm
Figure A3.8.2 General view of 2360S adapter A
A3.8-2 φ2420
φ2360 (PCD)
Spacecraft separation plane
Separation nut (8 pl) 294 (REF) 300 (REF)
Section A-A
Unit : mm
Figure A3.8.3 Details of 2360S adapter #1 A
A3.8-3 φ2360 (REF)
0 33.34 -0.05
18.35±0.05 0.25 ± 0 -0.05 1 11
60° +0.05 0 Separation nut
Details of separation nut (Adapter side)
Unit: mm
Figure A3.8.4 Details of 2360S adapter #2 A
A3.8-4 Bolt catcher envelope
35 min
φ50 134.1 min
Spacecraft Separation plane
φ18.5±0.05 0.1 ± +0.05 20 φ35.5 0
60° 0° -0.25°
Unit: mm Details of bolt catcher (spacecraft side)
Figure A3.8.5 Details of 2360S adapter #3 A
A3.8-5 APPENDIX 3.9 3470S ADAPTER A
The main characteristics are as follows. C (1) Interface diameter : 3,472 mm (PCD) (2) Height : 1,089 mm (3) Material : Aluminum semi-monocoque (4) Attached system : 4 places 15.9 mm (5/8") Bolts and Separation nuts B (5) Separation system : None (6) Separation nuts Maximum tension : 147 kN / Bolt B (7) Maximum load per spring : None (8) Adapter mass : 350 kg
Figure A3.9.1 shows a general view of the 3470S adapter. A Figures A3.9.2 to A3.9.4 show details of the 3470S adapter.
A3.9-1 B 3472
Separation bolt� B attach point 4pl
φ4030 B
φ2930 B
B 1089
φ2190 (PCD)
Unit: mm
Figure A3.9.1 General view of 3470S adapter A
A3.9-2 Separation plane
Separation nut (4 pl)
Figure A3.9.2 Details of 3470S adapter #1 A
A3.9-3 φ4030 φ2930 1080 (REF)
++ + + + +
536
(
R
59
E °
F ++ ) ++ + + + + -A-
Unit: mm
Figure A3.9.3 Details of 3470S adapter #2 A
A3.9-4 φ61 (REF) (Interface range)
0 φ37.84 -0.05
φ17.10±0.05 11 1 ± 0.25 -0.05 0
+15« 30° 0
Spacecraft connecting point details
17.1 (REF)
Unit : mm
Figure A3.9.4 Details of 3470S adapter #3 A
A3.9-5 APPENDIX 3.10 1666MA ADAPTER C
The main characteristics are as follows. (1) Interface diameter : 1,666 mm (2) Height : 480 mm (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 4 or 8 springs (6) Clamp bands Maximum tension : 38.9 kN (7) Maximum load per spring : 1670 N (8) Adapter mass : 100 kg
This adapter has 4 vent holes of 45 mm ø (1590.4 mm2) assuming that the internal volume of the spacecraft is less than 1.5 m3. When interface connectors are installed on the separation plane, connectors are located 942.5 mm from the center of the vehicle axis.
Figure A3.10.1 shows the photograph of the 1666MA adapter. Figure A3.10.2 shows the general view of the 1666MA adapter. Figures A3.10.3 to A3.10.5 show the details of the 1666MA adapter. Figures A3.10.6 to A3.10.7 show the stay-out zone around the 1666MA adapter. Figure A3.10.8 shows the limit loads of the 1666MA adapter. Figure A3.10.9 shows the spacecraft separation shock spectrum of the 1666MA adapter. Figure A3.10.10 shows the limit loads at separation plane of the 1666MA adapter.
Figure A3.10.1 Photograph of the 1666MA adapter
A3.10-1 C
+Y （S/C） Micro switch (2pl)
F .5 22 22 ° .5°
45 F ° A
A
2
2
.
5
° -X +X
（S/C） 2 （S/C）
2
.
5 °
Spring housing (8pl)
-Y （S/C）
Spacecraft separation plane
0.5 A
0.3 0.03/10 φ1666（REF） 480
A φ2190（PCD）
Unit : mm
Figure A3.10.2 General view of the 1666MA adapter
A3.10-2 C
φ1600
Micro switch
Sect F-F �
Spacecraft separation plane
C
B
Spring housing
Sect A-A � Unit : mm
Figure A3.10.3 Details of the 1666MA adapter #1
A3.10-3 C
0.25 6± >90 120
° > α >69 R6 ° >8 0.25
± R 4.5 2 D
1.6 0.3 φ1538±0.5 0.03 / 10 φ1633±0.5 1666 0.5 φ ± 0.3
1666 0.15 0.05 A φ ± 0.3
0.1 B 47.5±0.25
φ1600 0.03 / 10 A 0.3
1.6 E
φ1538±0.5
Figure A3.10.4 Details of the 1666MA adapter #2
A3.10-4 C
φ1666±0.5（REF）�
φ1662.2±0.15 0.3 + 0.12 0.125 C φ1660.52 0 3.07±0.03 10°±0.5°� 1.6
0 R 0.5
15°-0.25°�
� 5.72
TYP） 0.76
C 0.2±0.1
R 0.3±0.1（ 1.52±0.05
C 0.4±0.1（TYP）�
Detail D
0 φ1660.17 -0.13 B
C0.2±0.1 2.54±0.03 （TYP）�
R0.5±0.13
1.27±0.03 R3
R0.5±0.13C0.2±0.1
R0.5 5.72
1.6 15°-0.25°� φ1662.2±0.15 0.3
0°�
Detail E Unit : mm
Figure A3.10.5 Details of the 1666MA adapter #3
A3.10-5 C
+Y （S/C）
±20° ±20° -X +X （S/C） （S/C）
-Y （S/C）
φ4600 φ2266 φ1666 φ1626 Spacecraft φ1375 rear Usable 1195 frame φ volume Spacecraft φ820 separation plane
569 100
φ 60 φ194 200 φ134 Stay out
zone 336
° 440 450 110 46 480 46°
15 （REF° ）
: This area must be coordinated with NASDA. Unit : mm
Figure A3.10.6 Stay-out zone around the 1666MA adapter (+X / -X)
A3.10-6 C
+Y （S/C）
±70°
-X +X （S/C） （S/C）
±70°
-Y （S/C）
φ4600 φ2266
φ1666 φ1626 Spacecraft φ1375 rear Usable 1195 frame φ volume Spacecraft φ820 separation 569 plane
φ 60 φ194 200 φ134 Stay out
zone 336 ° 440 12 450 110 46° 480
15 （REF° ）
: This area must be coordinated with NASDA. Unit : mm
Figure A3.10.7 Stay-out zone around the 1666MA adapter (+Y / -Y)
A3.10-7 7000 C
6000
5000
4000
3000 Allowable range Spacecraft mass (kg) 2000
1000
0 0.0 0.5 1.0 1.5 2.0 2.2 2.5 3.0 3.5 4.0 4.5 5.0 Spacecraft C.G. height from separation plane (m)
Figure A3.10.8 Limit loads of the 1666MA adapter
A3.10-8 10000 C 1500Hz 3000Hz
4100G
1000 Acceleration (G) 100
Q=10 10 100 1000 10000 Frequency (Hz) G = 9.80665 m/s2
Figure A3.10.9 Spacecraft separation shock spectrum of the 1666MA adapter
A3.10-9 C 100
80.6
50 Allowable area
12.1 0 Axial load (ton) -50 -56.5
Tension-100 Compression 0.0 5.0 10.0 15.0 20.0 25.028.6 30.0
Bending moment (ton¥m)
Figure A3.10.10 Limit loads at separation plane of the 1666MA adapter
A3.10-10 C APPENDIX 3.11 239M ADAPTER
The main characteristics are as follows. (1) Interface diameter : 239 mm (2) Height : 100 mm (3) Material : Aluminum semi-monocoque (4) Attached system : Clamp bands (5) Separation system : 3 springs (6) Clamp band Maximum tension : 7.00 kN (7) Maximum load per spring : 500 kN (8) Adapter mass : 5.0 kg
This adapter is for the piggyback satellite, and is mounted on the piggyback satellite support structure which is installed at the side of the PSS. The mass of the spacecraft shall be 50 to 100 kg. When interface connector is installed on the separation place, the connector shall be located at the center of adapter axis.
Figure A3.11.1 shows the general view of the 239M adapter . Figure A3.11.2 shows the details of the 239M adapter. Figure A3.11.3 shows the spacecraft envelope of the piggyback satellite for the 239M adapter.
A3.11-1 C
Interface point of spacecraft separation switch (2pl)
） P Y T （ ゜ R77.5 0 6
Separation spring (3pl)
）
P
Y
T （
）�
゜
0
2 1 TYP （ 58
Interface connector mounting hole
φ239 100
(Unit : mm)
Figure A3.11.1 General view of the 239M adapter
A3.11-2 φ225±0.25 C φ219±0.5
R4 R4 6±0.25 2±0.25
+0¼15' R 1.6 45¼�0 4
1.6 5±0.25
0.3 +0¼15'
0.03/10 45¼�0 φ135±0.5
+0.13 0.3 φ220 0� A � 0.1 φ239.0±0.13 0.3
A Spacecraft side
φ239.0±0.13 0.3
0 0.1 φ220 -0.13� B 0.3 � 0 ¼ ¼ 2±0.25 45 �-0 15' 0.03/10 0.3 A B 3.0±0.25 1.6 2±0.25 2±0.25
1.6 R0.5
(type (m a x) ) +0¼15' 45¼�0
Adapter side (Unit:mm)
Figure A3.11.2 Details of the 239M adapter
A3.11-3 C 500 500
Spacecraft envelope 450 50
239M adapter
(Unui:mm)
Figure A3.11.3 Spacecraft envelope of the piggyback satellite for the 239M adapter
A3.11-4