长征三号甲系列运载火箭用户手册-2011

# 长征三号甲系列运载火箭用户手册（LM-3A Series Launch Vehicle User's Manual, Issue 2011）

**来源**：pdf/长征三号甲系列运载火箭用户手册-2011.pdf（25.6MB / 254页 / 有文本层）
**发布方**：CALT（中国运载火箭技术研究院）/ CGWIC（中国长城工业集团）/ CLTC（中国卫星发射测控系统部）
**语言**：英文
**版本**：Issue 2011
**页数**：254
**提取方式**：文本层直提（_pdf_text.py / PyMuPDF）
**提取日期**：2026-08-22
**校对状态**：🌱 未校对（文本层直提，质量高）
**世界线**：[OTL]
**备注**：长征三号甲系列（CZ-3A / CZ-3B / CZ-3C）运载火箭官方用户手册——总体参数/性能/发射服务/接口要求/环境条件/任务规划
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LM-3A Series Launch Vehicle User’s Manual
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LM-3A Series Launch Vehicle

User’s Manual

Issue 2011

APPROVALS

CEN Zheng
GAO Ruofei
SHEN Lei
Director General
LM-3A Series
Executive Vice President
President
China Academy of Launch
Vehicle Technology
(CALT)
China Great Wall Industry
Corporation
(CGWIC)
China Satellite Launch &
Tracking Control General
(CLTC)

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[ Intentionally Blank ]

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FOREWORD
This Long March 3A (LM-3A) Series Launch Vehicle User’s Manual is intended to provide
the essential technical and programmatic information on the LM-3A, LM-3B, LM-3B
Enhanced (LM-3BE) and LM-3C launch vehicles for the customers’ preliminary mission
planning and spacecraft design. The Manual also provides information of the launch site
facilities, the documentation required, typical launch processing, the mission analysis
requirements and details of the additional engineering support available to the customer’s
mission. This Manual is to be used as the baseline for using the LM-3A Series launch
vehicle, and the detailed requirements will be defined in the relevant technical documents of
the launch services contract.
This Manual will be updated as necessary. The updated information will be provided to
existing customers. The current version and all updates will also be posted to the
www.cgwic.com and www.calt.com web site for download as needed.
This Issue will supersede any information given in the previous issues or occasions.
Users of this Manual are encouraged to contact the offices listed below to discuss the LM-3A
Series launch vehicle and how the Long March (LM) launch vehicles can meet the user’s
requirements.
For general issues:
Launch Services Division
China Great Wall Industry Corporation
Building 2, No. 18A, Xing Shi Kou Road, Haidian District
Beijing 100195, P. R. China
Tel:
+86 10 8810 2168
Fax:
+86 10 8810 2109
E-mail:
launch@cgwic.com
www.cgwic.com
For technical issues:
Space Division
China Academy of Launch Vehicle Technology
No.1, Nan Da Hong Men Road, Fengtai District
Beijing 100076, P. R. China
Tel:
+86 10 6876 9837, +86 10 6875 3701
Fax:
+86 10 6875 7398
E-mail:
caltinfo@calt.com.cn or launch@cgwic.com
www.calt.com

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REVISION HISTORY
Issue
Title
Change Description
Approval
2011
LM-3A Series Launch
Vehicle User’s Manual
Completely revised compared to
Issue 1999.
CEN Zheng
Director General
LM-3A Series
2003
LM-3A Series Launch
Vehicle User’s Manual
(Chinese Version)
Three manuals in Issue 1999 were
integrated into one manual in Issue
2003. (Only Chinese version was
published for Issue 2003.)
LONG Lehao
Director General
LM-3A Series
1999
LM-3A User’s Manual
LM-3B User’s Manual
LM-3C User’s Manual
There were three independent
manuals in Issue 1999, for LM-3A,
LM-3B and LM-3C respectively.
LONG Lehao
Director General
LM-3A Series

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PREFACE
China Aerospace Science and Technology Corporation (CASC) evolved from China
Aerospace Corporation, which was created from the Fifth Academy of the Ministry of
National Defense established on October 8, 1956. The Fifth Academy was expanded into
the Seventh Ministry of Machinery Industry and underwent further changes through the
Ministry of Space Industry, the Ministry of Aerospace Industry and China Aerospace
Industry Corporation, from which CASC was formally founded on July 1, 1999 with the
approval of the State Council of the People’s Republic of China. CASC is a large
state-owned enterprise group with its own internationally recognized brands such as Long
March and Shenzhou. CASC has developed its own intellectual properties in the space
industry, and it has an outstanding record for its core business and is highly competitive in
the international market. As an institutional investor authorized by the State Council, CASC
has eight large R&D and production entities in addition to China Great Wall Industry
Corporation, plus many other institutions and companies directly subordinated to it. CASC’s
primary business includes research, design, manufacture and launch of space systems and
products, such as launch vehicles, satellites and manned spaceships.
China Great Wall Industry Corporation (CGWIC) is a launch services provider offering low
risk and reliable solutions to deliver the customers’ satellites into a wide range of orbits. The
launch services are provided directly through CGWIC, which is the exclusive commercial
organization authorized by the Chinese Government to offer Long March launch services to
the international market. As the prime contractor for launch services, CGWIC is responsible
for marketing launch services, negotiation and implementation of the launch services
contract. CGWIC provides the customer with a single point of contact for all communications
related to launch services contract.
China Academy of Launch Vehicle Technology (CALT) is the designer and manufacturer of
LM-3A Series launch vehicles and is a subcontractor for launch services. CALT is
responsible for the launch vehicle design, development, manufacture, testing, mission
analyses, launch vehicle technical interface coordination and flight safety engineering.
CGWIC and CALT are both wholly – owned subsidiaries of CASC.
China Satellite Launch and Tracking Control General (CLTC) is in charge of China’s TT&C
network and launch sites, including XSLC, JSLC and TSLC. As the subcontractor of Long

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March launch services, CLTC is responsible for the TT&C and launch site technical and
interface coordination, launch site operations, launch control and launch site safety.
In cooperation with CALT and CLTC, CGWIC has established an outstanding launch record
that clearly demonstrates the high reliability of Long March launch vehicles and their ability
to deliver the customers’ satellite into designed orbit in an accurate and reliable manner.

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TABLE OF CONTENTS

APPROVAL PAGE .............................................................................................................0-1
FOREWORD ......................................................................................................................0-3
REVISION HISTORY..........................................................................................................0-4
PREFACE...........................................................................................................................0-5
TABLE OF CONTENTS .....................................................................................................0-7
LIST OF FIGURES...........................................................................................................0-15
LIST OF TABLES.............................................................................................................0-20
ABBREVIATIONS AND ACRONYMS..............................................................................0-22
CHAPTER 1 INTRODUCTION...........................................................................................1-1
1.1 Long March Family of Launch Vehicles - Development History......................................1-3
1.2 Long March Launch Services....................................................................................1-4
1.2.1 Overview.............................................................................................................1-4
1.2.2 Customer Interfaces ...........................................................................................1-6
1.3 Overview of the Launch Centers ...............................................................................1-6
1.3.1 Xichang Satellite Launch Center (XSLC)............................................................1-7
1.3.2 Taiyuan Satellite Launch Center (TSLC).............................................................1-8
1.3.3 Jiuquan Satellite Launch Center (JSLC).............................................................1-9
1.4 Launch Record of Long March..................................................................................1-9
1.5 Long March Launch Services Advantages..............................................................1-10
CHAPTER 2 GENERAL DESCRIPTION............................................................................2-1
2.1 Summary...................................................................................................................2-1
2.2 LM-3A Launch Vehicle ..............................................................................................2-2
2.2.1 Technical Parameters.........................................................................................2-2
2.2.2 LM-3A Description ..............................................................................................2-4
2.3 LM-3B and LM-3BE Launch Vehicles .....................................................................2-15
2.3.1 Technical Parameters.......................................................................................2-15
2.3.2 LM-3B and LM-3BE Description .......................................................................2-17
2.4 LM-3C Launch Vehicle............................................................................................2-21
2.5 Launch Vehicle Coordinate System ........................................................................2-22
2.6 Launch Missions of LM-3A Series Launch Vehicles................................................2-23
2.7 Satellites Launched.................................................................................................2-24
2.8 Future Enhancements for LM-3A Series Launch Vehicles......................................2-26

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CHAPTER 3 LAUNCH PERFORMANCE AND INJECTION ACCURACY ........................3-1
3.1 Summary...................................................................................................................3-1
3.2 Standard Mission Profile .......................................................................................... 3-1
3.2.1 Coordinates of Launch Pads ............................................................................. 3-1
3.2.2 Geostationary Transfer Orbit Missions and Other Orbit Missions ...................... 3-4
3.2.3 GTO Mission Injection Optimization................................................................... 3-5
3.2.4 Minimum Residual Shutdown (MRS) Missions .................................................. 3-5
3.2.5 Orbit Injection Accuracy..................................................................................... 3-6
3.2.6 Pointing Accuracy .............................................................................................. 3-6
3.2.7 Separation and Spin-up Accuracy...................................................................... 3-7
3.2.8 Launch Windows ............................................................................................... 3-8
3.2.9 Separation Attitude ............................................................................................ 3-9
3.2.10 Separation Velocity.......................................................................................... 3-9
3.2.11 Collision Avoidance Maneuver......................................................................... 3-9
3.3 LM-3A Launch Vehicle ............................................................................................. 3-9
3.3.1 Typical Orbit Parameters ................................................................................... 3-9
3.3.2 Launch Capability for LM-3A Launch Vehicle .................................................. 3-13
3.4 LM-3B Launch Vehicle ........................................................................................... 3-15
3.4.1 Typical Orbit Parameters ................................................................................. 3-15
3.4.2 Launch Capability for LM-3B Launch Vehicle .................................................. 3-19
3.5 LM-3BE Launch Vehicle......................................................................................... 3-23
3.5.1 Typical Orbit Parameters ................................................................................. 3-23
3.5.2 Launch Capability for GTO Missions ............................................................... 3-26
3.6 LM-3C Launch Vehicle........................................................................................... 3-30
3.6.1 Typical Orbit Parameters ................................................................................. 3-30
3.6.2 Launch Capability for GTO Missions ............................................................... 3-33
CHAPTER 4 PAYLOAD FAIRING......................................................................................4-1
4.1 Introduction ...............................................................................................................4-1
4.2 3350 Fairing ..............................................................................................................4-2
4.3 3700Z Fairing............................................................................................................4-5
4.4 4000F Fairing............................................................................................................4-8
4.5 4200F Fairing..........................................................................................................4-11
4.6 4200Z Fairing..........................................................................................................4-12
4.7 Static Envelope .......................................................................................................4-13
4.8 Dynamic Envelope ..................................................................................................4-14
4.9 Thermal Protection..................................................................................................4-14
4.10 Fairing Separation.................................................................................................4-14

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4.10.1 Lateral Release Mechanism ...........................................................................4-17
4.10.2 Longitudinal Release Mechanism...................................................................4-17
4.10.3 Fairing Separation Mechanism .......................................................................4-17
4.11 RF Windows and Access Doors............................................................................4-17
4.12 Customer Logo......................................................................................................4-18
CHAPTER 5 MECHANICAL AND ELECTRICAL INTERFACES.......................................5-1
5.1 Introduction ...............................................................................................................5-1
5.2 Satellite Reference Axes...........................................................................................5-1
5.3 Mechanical Interface .................................................................................................5-2
5.3.1 Introduction.........................................................................................................5-2
5.3.2 Payload Adapter .................................................................................................5-2
5.3.3 Separation System .............................................................................................5-8
5.4 Electrical Interfaces.................................................................................................5-13
5.4.1 Summary ..........................................................................................................5-13
5.4.2 In-Flight-Disconnectors.....................................................................................5-15
5.4.3 In-Flight-Disconnectors Characteristics ............................................................5-15
5.4.4 Umbilical Cable Onboard Launch Vehicle.........................................................5-16
5.4.5 Ground Umbilical Cables..................................................................................5-17
5.4.6 Shielding and Grounding ..................................................................................5-19
5.5 Satellite Constraints ................................................................................................5-19
5.6 RF Links..................................................................................................................5-19
5.6.1 RF Relay Path ..................................................................................................5-20
5.6.2 Characteristics and Interface of RF Link...........................................................5-20
5.7 Post Encapsulation Interfaces.................................................................................5-20
5.7.1 Payload Usable Volume....................................................................................5-20
5.7.2 Satellite Access ................................................................................................5-21
5.8 Interface Verification ...............................................................................................5-21
5.8.1 Satellite AIT..........................................................................................................5-21
5.8.2 Launch Site Interface Verification .....................................................................5-21
5.8.3 Satellite EGSE Interface Verification.................................................................5-22
CHAPTER 6 ENVIRONMENTAL CONDITIONS ................................................................6-1
6.1 Summary...................................................................................................................6-1
6.2 Environment at XSLC................................................................................................6-1
6.3 Satellite Processing Environment..............................................................................6-2
6.3.1 Environment in SC Processing Building BS2 and BS3.......................................6-2
6.3.2 Environment during Transportation to Launch Pad.............................................6-3
6.3.3 Air-conditioning inside Fairing.............................................................................6-3

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6.4 Electromagnetic Environment ...................................................................................6-5
6.4.1 Radio Equipment onboard LM-3A Series launch vehicles and Ground Test Equipment...6-5
6.4.2 RF Equipment and Radiation Strength at XSLC.................................................6-6
6.4.3 Launch Vehicle Electromagnetic Radiation and Susceptibility............................6-6
6.4.4 EMC Analysis .....................................................................................................6-8
6.5 Mechanical Environments .........................................................................................6-9
6.5.1 Steady State Acceleration...................................................................................6-9
6.5.2 Vibration Environments.......................................................................................6-9
6.5.3 Acoustic Noise ..................................................................................................6-10
6.5.4 Shock Environment...........................................................................................6-10
6.6 Thermal Environment..............................................................................................6-12
6.6.1 Thermal Environment prior to Fairing Jettison ..................................................6-12
6.6.2 Thermal Environment Post Fairing Jettison ......................................................6-12
6.6.3 Launch Vehicle Generated Heat Flux ...............................................................6-13
6.7 Pressure Environment.............................................................................................6-13
6.8 Contamination and Cleanliness Control ..................................................................6-13
6.8.1 Contamination Control......................................................................................6-14
6.8.2 Cleanliness Control...........................................................................................6-14
CHAPTER 7 LAUNCH SITE ..............................................................................................7-1
7.1 General Description ..................................................................................................7-1
7.1.1 Xichang Airport ...................................................................................................7-2
7.1.2 Transfer to XSLC ................................................................................................7-4
7.2 Xichang Satellite Launch Center...............................................................................7-4
7.3 Technical Center .......................................................................................................7-4
7.4 LV Processing (Test) Building (BL) ...........................................................................7-5
7.4.1 Transit Building (BL1) .........................................................................................7-5
7.4.2 Testing Building (BL2).........................................................................................7-6
7.5 SC Processing Buildings...........................................................................................7-7
7.5.1 Non-Hazardous Operation Building (BS2)........................................................7-12
7.5.2 Hazardous Operation Building (BS3)................................................................7-13
7.5.3 Solid Propellant Motor Checkout and Processing Building (BM) ......................7-15
7.5.4 SPM X-ray Detection Building (BMX) ...............................................................7-16
7.5.5 Storehouse of Hazardous Substances..............................................................7-18
7.5.6 Power Supply, Grounding, Lightning Protection, Fire-De tection & Alarm .............7-18
7.5.7 Launch Control Center (LCC) ...........................................................................7-19
7.6 Launch Complexes .................................................................................................7-19
7.6.1 General Description..........................................................................................7-20

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7.6.2 No.3 Launch Complex (LC-3) ...........................................................................7-21
7.6.3 No.2 Launch Complex (LC-2) ...........................................................................7-21
7.6.4 Power Supply for Launch Pad ..........................................................................7-23
7.6.5 SC Fore-positioned Equipment Room (FER)....................................................7-23
7.7 Mission Command & Control Center (MCCC).........................................................7-24
7.7.1 General Description..........................................................................................7-24
7.7.2 Functions of MCCC ..........................................................................................7-25
7.7.3 Configuration of MCCC.....................................................................................7-26
7.8 Tracking, Telemetry and Control System (TT&C) ...................................................7-26
7.8.1 General Description..........................................................................................7-26
7.8.2 Main Functions of TT&C...................................................................................7-26
7.8.3 Composition of TT&C .......................................................................................7-26
7.8.4 Tracking Sequence of TT&C System................................................................7-27
7.9 Transportation and Handling...................................................................................7-28
7.10 Propellants and Gases..........................................................................................7-31
CHAPTER 8 LAUNCH OPERATIONS...............................................................................8-1
8.1 Overview of Launch Site Operations.........................................................................8-1
8.2 SC Operations at the Launch Site.............................................................................8-1
8.3 Launch Vehicle Launch Operations ..........................................................................8-1
8.4 SC/LV Combined Operations....................................................................................8-3
8.4.1 General Description............................................................................................8-3
8.4.2 SC/LV Combined Operations for Encapsulation-on-Pad ....................................8-3
8.4.3 SC/LV Combined Operations for Encapsulation-in-BS3 ...................................8-13
8.4.4 SC Preparation and Checkouts ........................................................................8-16
8.5 Launch Limitation....................................................................................................8-17
8.6 Pre-launch Countdown Procedure ..........................................................................8-17
8.7 Post-launch Activities ..............................................................................................8-18
CHAPTER 9 RANGE SAFETY CONTROL........................................................................9-1
9.1 Responsibility and Requirements of Range Safety ...................................................9-1
9.2 Range Safety Issues for a LV Anomaly.....................................................................9-1
9.2.1 Safety Control Plan.............................................................................................9-1
9.2.2 Safety Control Procedure ...................................................................................9-2
9.3 Composition of the Range Safety Control System ....................................................9-3
9.4 Safety Criteria ...........................................................................................................9-4
9.4.1 Approval Procedure of Range Safety Criteria.....................................................9-5
9.4.2 Common Criteria of Range Safety Control .........................................................9-5
9.4.3 Special Criteria ...................................................................................................9-5

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9.5 Emergency Measures ...............................................................................................9-5
CHAPTER 10 SATELLITE DESIGN REQUIREMENTS...................................................10-1
10.1 Introduction ...........................................................................................................10-1
10.2 Load Conditions for Satellite Design .....................................................................10-1
10.2.1 Frequency Requirement .................................................................................10-1
10.2.2 Loads Applied for SC Structure Design...........................................................10-1
10.3 Coupled Load Analysis..........................................................................................10-2
10.4 Safety Requirements.............................................................................................10-3
10.5 Satellite Mass and CoG Requirements .................................................................10-4
10.6 Qualification and Acceptance Test Specifications.................................................10-4
10.6.1 Static Loads Test (Qualification)......................................................................10-4
10.6.2 Sinusoidal Vibration Test.................................................................................10-4
10.6.3 Acoustic Test...................................................................................................10-5
10.6.4 Shock Test ......................................................................................................10-6
10.6.5 Proto-flight Test...............................................................................................10-8
CHAPTER 11 MISSION INTEGRATION AND MANAGEMENT.......................................11-1
11.1 Introduction ...........................................................................................................11-1
11.2 Management Overview .........................................................................................11-1
11.2.1 Management Organization..............................................................................11-1
11.2.2 Approach for Successful Implementation of the Program ...............................11-2
11.2.3 Program Management Reviews......................................................................11-2
11.2.4 SC/LV Interface Control ..................................................................................11-3
11.2.5 Launch Services Schedule .............................................................................11-3
11.2.6 Launch Campaign Schedule...........................................................................11-5
11.2.7 Launch License and Permits...........................................................................11-5
11.2.8 Satellite Technology Safeguard.......................................................................11-6
11.3 Launch Services Program.....................................................................................11-6
11.4 Interface Control Document (ICD).........................................................................11-6
11.5 Mission Analysis....................................................................................................11-7
11.5.1 Introduction .....................................................................................................11-7
11.5.2 Preliminary Mission Analysis...........................................................................11-7
11.5.3 Final Mission Analysis...................................................................................11-10
11.6 Verification of Satellite Design.............................................................................11-12
11.6.1 Satellite Environmental Test Plan..................................................................11-12
11.6.2 Satellite Environmental Test Program ...........................................................11-13
11.6.3 Satellite Environmental Test Report ..............................................................11-13
11.7 Interface Meetings and Reviews .........................................................................11-13

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11.7.1 Kick-Off Meeting (KOM)................................................................................11-14
11.7.2 Phase 1/2/3 Safety Submission ....................................................................11-15
11.7.3 Preliminary Mission Analysis Review (PMAR) ..............................................11-15
11.7.4 Technical Interface Meeting (TIM).................................................................11-15
11.7.5 Final Mission Analysis Review (FMAR).........................................................11-15
11.7.6 Launch Site Survey.......................................................................................11-15
11.7.7 Launch Vehicle Pre-shipment Review (PSR)................................................11-16
11.7.8 Launch Site Operation Meetings...................................................................11-16
11.7.9 Combined Operations Procedure Review.....................................................11-16
11.7.10 Launch Readiness Review (LRR)...............................................................11-16
11.8 Launch Reports after SC/LV Separation .............................................................11-16
11.8.1 Orbit Injection Parameters Report.................................................................11-17
11.8.2 Orbit Tracking Report....................................................................................11-17
11.8.3 Launch Mission Evaluation Report................................................................11-17
11.9 The Customer Commitment ................................................................................11-17
11.9.1 Documentation..............................................................................................11-17
11.9.2 Hardware for Integration and Launch............................................................11-18
11.10 Launch Campaign .............................................................................................11-18
11.10.1 Introduction .................................................................................................11-18
11.10.2 Launch Campaign Preparation ...................................................................11-19
11.10.3 Launch Campaign Documentation..............................................................11-19
11.10.4 Launch Campaign Management.................................................................11-22
11.10.5 Launch Campaign Working Meetings .........................................................11-23
11.11 Safety Requirements.........................................................................................11-24
11.11.1 Safety Training ............................................................................................11-24
11.11.2 Safety Requirements for Hazardous Operations.........................................11-25
11.12 Launch Services Documentation.......................................................................11-25

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LIST OF FIGURES

FIGURE 1-1 LONG MARCH FAMILY OF LAUNCH VEHICLES ......................................................1-1
FIGURE 1-2 LONG MARCH FLIGHTS .....................................................................................1-3
FIGURE 1-3 THE ORGANIZATION OF LONG MARCH LAUNCH SERVICES...................................1-5
FIGURE 1-4 LONG MARCH LV ASSEMBLY AND TEST WORKSHOP............................................1-6
FIGURE 1-5 SATELLITE LAUNCH CENTERS............................................................................1-7
FIGURE 1-6 XICHANG SATELLITE LAUNCH CENTER ...............................................................1-8
FIGURE 1-7 TAIYUAN SATELLITE LAUNCH CENTER ................................................................1-8
FIGURE 1-8 JIUQUAN SATELLITE LAUNCH CENTER................................................................1-9
FIGURE 2-1 LM-3A/3B/3BE/3C CONFIGURATION.................................................................2-2
FIGURE 2-2 LM-3A LAUNCH VEHICLE ..................................................................................2-3
FIGURE 2-3 CALT ASSEMBLY AREA .....................................................................................2-5
FIGURE 2-4 LM-3A VEB CONFIGURATION............................................................................2-6
FIGURE 2-5 SECOND STAGE ENGINES .................................................................................2-7
FIGURE 2-6A SCHEMATIC DIAGRAM OF STAGE-1 PROPULSION SYSTEM .................................2-8
FIGURE 2-6B SCHEMATIC DIAGRAM OF STAGE-2 PROPULSION SYSTEM .................................2-8
FIGURE 2-6C SCHEMATIC DIAGRAM OF STAGE-3 PROPULSION SYSTEM.................................2-9
FIGURE 2-7A SCHEMATIC DIAGRAM OF LM-3A GUIDANCE & CONTROL SYSTEM ...................2-10
FIGURE 2-7B SCHEMATIC DIAGRAM OF LM-3A ATTITUDE CONTROL SYSTEM........................2-10
FIGURE 2-7C SCHEMATIC DIAGRAM OF LM-3A GUIDANCE SYSTEM......................................2-10
FIGURE 2-8 SCHEMATIC DIAGRAM OF LM-3A TELEMETRY SYSTEM...................................... 2-11
FIGURE 2-9 SCHEMATIC DIAGRAM OF LM-3A TRACKING & RANGE SAFETY SYSTEM.............2-12
FIGURE 2-10 LM-3A PROPELLANT MANAGEMENT & ATTITUDE CONTROL SYSTEM ................2-12
FIGURE 2-11 LM-3A SERIES LAUNCH VEHICLES PROPELLANT UTILIZATION SYSTEM.............2-13
FIGURE 2-12 LM-3A/3B/3C SEPARATION EVENTS..............................................................2-14
FIGURE 2-13 LM-3B LAUNCH VEHICLE ..............................................................................2-15
FIGURE 2-14 LM-3B & LM-3C FAIRING FOR ENCAPSULATION-ON-PAD ................................2-18
FIGURE 2-15 LM-3B & LM-3C FAIRING FOR ENCAPSULATION-IN-BS3.................................2-19
FIGURE 2-16 LM-3B/3BE/3C PROPELLANT MANAGEMENT & ATTITUDE CONTROL SYSTEM...........2-20
FIGURE 2-17 LM-3C LAUNCH VEHICLE..............................................................................2-22
FIGURE 2-18 COORDINATE SYSTEM OF LM-3A/3B/3BE/3C................................................2-23
FIGURE 2-19 SCHEMATIC DRAWING OF GTO AND GEO ......................................................2-24
FIGURE 3-1 LAUNCH AZIMUTH OF LM-3A.............................................................................3-2

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FIGURE 3-2 LAUNCH AZIMUTH OF LM-3B AND LM-3BE.........................................................3-3
FIGURE 3-3 LAUNCH AZIMUTH OF LM-3C.............................................................................3-4
FIGURE 3-4 PERIGEE COORDINATE SYSTEMS.......................................................................3-7
FIGURE 3-5 DEFINITION OF SEPARATION ACCURACY.............................................................3-8
FIGURE 3-6 LM-3A FLIGHT PROFILE.................................................................................. 3-11
FIGURE 3-7A LM-3A LOAD FACTOR & VELOCITY VS. FLIGHT TIME........................................3-12
FIGURE 3-7B LM-3A ALTITUDE & MACH VS. FLIGHT TIME....................................................3-12
FIGURE 3-8 LM-3A GTO CAPABILITY.................................................................................3-14
FIGURE 3-9 LM-3A SSO CAPABILITY.................................................................................3-15
FIGURE 3-10 LM-3A EARTH-ESCAPE MISSION CAPABILITY .................................................3-15
FIGURE 3-11 LM-3B GTO FLIGHT PROFILE........................................................................3-17
FIGURE 3-12A LM-3B LOAD FACTOR & VELOCITY VS. FLIGHT TIME .....................................3-19
FIGURE 3-12B LM-3B FLIGHT MACH & ALTITUDE VS. FLIGHT TIME ......................................3-18
FIGURE 3-13 LM-3B GTO PERFORMANCE (ENCAPSULATION-ON-PAD)................................3-20
FIGURE 3-14 LM-3B GTO PERFORMANCE (ENCAPSULATION-IN-BS3).................................3-21
FIGURE 3-15 LM-3B SUN SYNCHRONOUS CAPABILITY ........................................................3-22
FIGURE 3-16 LM-3B EARTH ESCAPE MISSION CAPABILITY..................................................3-22
FIGURE 3-17 LM-3BE FLIGHT PROFILE .............................................................................3-24
FIGURE 3-18A LM-3BE LOAD FACTOR & VELOCITY VS. FLIGHT TIME...................................3-26
FIGURE 3-18B LM-3BE ALTITUDE & MACH VS. FLIGHT TIME ...............................................3-26
FIGURE 3-19 LM-3BE GTO PERFORMANCE (ENCAPSULATION-ON-PAD)..............................3-27
FIGURE 3-20 LM-3BE GTO PERFORMANCE (ENCAPSULATION-IN-BS3) ..............................3-28
FIGURE 3-21 LM-3B ENHANCED SSO CAPABILITY .............................................................3-29
FIGURE 3-22 LM-3BE EARTH ESCAPE MISSION CAPABILITY ...............................................3-29
FIGURE 3-23 LM-3C FLIGHT PROFILE................................................................................3-31
FIGURE 3-24A LM-3C LOAD FACTOR & VELOCITY VS. FLIGHT TIME .....................................3-33
FIGURE 3-24B LM-3C ALTITUDE & MACH VS. FLIGHT TIME..................................................3-33
FIGURE 3-25 LM-3C GTO PERFORMANCE (ENCAPSULATION-ON-PAD)................................3-34
FIGURE 3-26 LM-3C GTO PERFORMANCE (ENCAPSULATION-IN-BS3).................................3-35
FIGURE 3-27 LM-3C SSO CAPABILITY...............................................................................3-36
FIGURE 3-28 LM-3C EARTH ESCAPE MISSION CAPABILITY .................................................3-36
FIGURE 4-1 PAYLOAD FARING ENCAPSULATION ON LAUNCH PAD ...........................................4-1
FIGURE 4-2A LM-3A FAIRING...............................................................................................4-3
FIGURE 4-2B 3350 FAIRING STATIC ENVELOPE.....................................................................4-4
FIGURE 4-3 3350 FAIRING ENCAPSULATION-ON-PAD ............................................................4-5
FIGURE 4-4A 3700Z FAIRING...............................................................................................4-6
FIGURE 4-4B 3700Z FAIRING STATIC ENVELOPE...................................................................4-7

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FIGURE 4-5A 4000F FAIRING...............................................................................................4-8
FIGURE 4-5B 4000F FAIRING STATIC ENVELOPE...................................................................4-9
FIGURE 4-6 4000F FARING ENCAPSULATION-ON-PAD.........................................................4-10
FIGURE 4-7 4000F FAIRING ENCAPSULATION-ON-PAD ........................................................4-10
FIGURE 4-8A 4200F/4200Z FAIRING.................................................................................. 4-11
FIGURE 4-8B 4200F FAIRING STATIC ENVELOPE.................................................................4-12
FIGURE 4-8C 4200Z FAIRING STATIC ENVELOPE ................................................................4-13
FIGURE 4-9A STRUCTURE & SEPARATION DEVICE OF LM-3A FAIRING..................................4-15
FIGURE 4-9B STRUCTURE & SEPARATION DEVICE OF LM-3B/3BE/3C FAIRING ....................4-15
FIGURE 4-10A ACCESS DOOR & RF WINDOW ALLOWABLE OF 3350 FAIRING........................4-19
FIGURE 4-10B ACCESS DOOR & RF WINDOW ALLOWABLE OF 4000F/4200F/4200Z FAIRING............4-20
FIGURE 4-10C ACCESS DOOR & RF WINDOW ALLOWABLE OF 3700Z FAIRING .....................4-21
FIGURE 5-1 LM-3A SERIES LAUNCH VEHICLE COORDINATE SYSTEM.....................................5-1
FIGURE 5-2 937B INTERFACE (ENCAPSULATION-ON-PAD) .....................................................5-3
FIGURE 5-3 1194 INTERFACE (ENCAPSULATION-ON-PAD)......................................................5-4
FIGURE 5-4 PAYLOAD ADAPTER WITH 1194 INTERFACE........................................................5-5
FIGURE 5-5 1194A INTERFACE (ENCAPSULATION-ON-PAD)....................................................5-6
FIGURE 5-6 1194A INTERFACE (ENCAPSULATION-IN-BS3) ....................................................5-7
FIGURE 5-7A C100 SEPARATION MECHANISM.......................................................................5-9
FIGURE 5-7B C60 SEPARATION MECHANISM ......................................................................5-10
FIGURE 5-7C C60 SEPARATION MECHANISM ......................................................................5-10
FIGURE 5-7D SC SEPARATION SPRINGS............................................................................. 5-11
FIGURE 5-7E 1194/1194A CLAMPBAND..............................................................................5-12
FIGURE 5-8 SC UMBILICAL CABLE CONFIGURATION............................................................5-14
FIGURE 5-9 SC RF LINKS & UMBILICAL CABLES.................................................................5-15
FIGURE 5-10 AN EXAMPLE OF SEPARATION SIGNALS & BREAK-WIRES.................................5-17
FIGURE 6-1 PRE-LAUNCH AIR-CONDITION SYSTEM OF LM-3A SERIES LVS.............................6-4
FIGURE 6-2 LOCATIONS OF ON-BOARD RADIO EQUIPMENTS..................................................6-6
FIGURE 6-3 ELECTROMAGNETIC RADIATION FROM LV AND XSLC ..........................................6-7
FIGURE 6-4 LAUNCH VEHICLE ELECTROMAGNETIC RADIATION SUSCEPTIBILITY......................6-8
FIGURE 6-5 SHOCK RESPONSE SPECTRUM AT SC/LV SEPARATION PLANE - C100 CLAMPBAND.. 6-11
FIGURE 6-6 SHOCK RESPONSE SPECTRUM AT SC/LV SEPARATION PLANE - C60 CLAMPBAND.... 6-11
FIGURE 6-7 FREE MOLECULAR HEATING FLUX OF LM-3A/3B/3BE/3C.................................6-12
FIGURE 6-8 FAIRING INTERNAL PRESSURE .........................................................................6-13
FIGURE 7-1 XICHANG SATELLITE LAUNCH CENTER ...............................................................7-1
FIGURE 7-2 LOCAL MAP OF XSLC .......................................................................................7-2
FIGURE 7-3 ANTONOV-124 LANDING AT XICHANG AIRPORT ...................................................7-3

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FIGURE 7-4 UNLOADING OPERATION AT XICHANG AIRPORT ...................................................7-3
FIGURE 7-5 SATELLITE TRANSFER FROM XICHANG AIRPORT TO XSLC...................................7-4
FIGURE 7-6 TECHNICAL CENTER OF XSLC...........................................................................7-5
FIGURE 7-7 TRANSIT BUILDING (BL1) ..................................................................................7-6
FIGURE 7-8 TESTING BUILDING (BL2)..................................................................................7-6
FIGURE 7-9 SC PROCESSING BUILDING (BS).......................................................................7-7
FIGURE 7-10 LAYOUT OF BS FIRST FLOOR...........................................................................7-8
FIGURE 7-11 LAYOUT OF BS SECOND FLOOR.......................................................................7-8
FIGURE 7-12 NON-HAZARDOUS OPERATION BUILDING (BS2)..............................................7-12
FIGURE 7-13 NON-HAZARDOUS OPERATION BUILDING (BS2)..............................................7-12
FIGURE 7-14 HAZARDOUS OPERATION BUILDING (BS3)......................................................7-14
FIGURE 7-15 SOLID PROPELLANT MOTOR PROCESSING BUILDINGS (BM)............................7-15
FIGURE 7-16 SOLID PROPELLANT MOTOR X-RAY BUILDING (BMX)......................................7-17
FIGURE 7-17 LAUNCH COMPLEXES OF XSLC.....................................................................7-20
FIGURE 7-18 TWO LAUNCH PADS IN XSLC (LEFT: LC-2, RIGHT: LC-3) ................................7-20
FIGURE 7-19 NO.3 LAUNCH COMPLEX (LC-3) ....................................................................7-21
FIGURE 7-20 NO.2 LAUNCH COMPLEX (LC-2) ....................................................................7-21
FIGURE 7-21 SC TEST ROOMS..........................................................................................7-23
FIGURE 7-22 XSLC MISSION COMMAND & CONTROL CENTER (MCCC) ..............................7-24
FIGURE 7-23 XSLC MISSION COMMAND & CONTROL CENTER (MCCC) ..............................7-24
FIGURE 7-24 MISSION COMMAND & CONTROL CENTER (MCCC) ........................................7-25
FIGURE 7-25 ANTENNAS OF TT&C SYSTEM .......................................................................7-27
FIGURE 7-26 TT&C OPERATIONS OF LM-3A/3B/3BE/3C GTO LAUNCH MISSIONS..............7-28
FIGURE 8-1 SC CONTAINER ................................................................................................8-4
FIGURE 8-2 SC CONTAINER TRANSFER................................................................................8-4
FIGURE 8-3 SC/LV INTEGRATION PROCESS..........................................................................8-6
FIGURE 8-4 TRANSFER SC CONTAINER TO THE LAUNCH PAD ................................................8-8
FIGURE 8-5 SC CONTAINER TRANSFER TO THE LAUNCH PAD ................................................8-9
FIGURE 8-6 SC CONTAINER LIFT .........................................................................................8-9
FIGURE 8-7 LAYOUT ON THE 8TH FLOOR OF THE SERVICE TOWER........................................8-10
FIGURE 8-8 FARING ENCAPSULATION-ON-PAD.................................................................... 8-11
FIGURE 8-9 COMBINED OPERATIONS OF FAIRING ENCAPSULATION-ON-PAD ..........................8-12
FIGURE 8-10 SC/LV INTEGRATION PROCESS IN BS3...........................................................8-14
FIGURE 8-11 TRANSIT OF SC/PLA/PLF FROM BS3 TO THE LAUNCH PAD.............................8-15
FIGURE 8-12 SC/LV MATING PROCESS...............................................................................8-16
FIGURE 9-1 FLOWCHART OF RANGE SAFETY CONTROL SYSTEM ...........................................9-3
FIGURE 9-2 RANGE SAFETY CONTROL SYSTEM AT XSLC......................................................9-4

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FIGURE 10-1 SC SINUSOIDAL VIBRATION TEST SPECIFICATION ...........................................10-5
FIGURE 10-2 SC ACOUSTIC TEST SPECIFICATION...............................................................10-6
FIGURE 10-3 SHOCK RESPONSE SPECTRUM AT SC/LV SEPARATION PLANE - C100 CLAMPBAND...........10-7
FIGURE 10-4 SHOCK RESPONSE SPECTRUM AT SC/LV SEPARATION PLANE - C60 CLAMPBAND.............10-8
FIGURE 11-1 PROGRAM ORGANIZATION ..................................................................................... 11-1
FIGURE 11-2 TYPICAL LAUNCH SERVICES PROGRAM SCHEDULE.................................................. 11-4
FIGURE 11-3 THE TYPICAL SCHEDULE ........................................................................................ 11-5

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LIST OF TABLES

TABLE 1-1 THE MAIN CHARACTERISTICS OF LONG MARCH FAMILY.........................................1-2
TABLE 1-2 RESPONSIBILITIES OF THE LAUNCH SERVICES RELATED ORGANIZATIONS...............1-5
TABLE 1-3 FLIGHT RECORD OF LONG MARCH (AS OF JANUARY 1, 2011) ..............................1-10
TABLE 2-1 SC LAUNCHED BY LM-3A SERIES LV .................................................................2-25
TABLE 3-1 LAUNCH CAPABILITY VS. COMMAND SHUTDOWN PROBABILITY...............................3-6
TABLE 3-2A INJECTION ACCURACY FOR STANDARD GTO MISSION (1 SIGMA)..........................3-6
TABLE 3-2B COVARIANCE MATRIX OF INJECTION FOR STANDARD GTO MISSION.....................3-6
TABLE 3-3 LM-3A TYPICAL FLIGHT SEQUENCE ...................................................................3-10
TABLE 3-4 CHARACTERISTIC PARAMETERS OF TYPICAL TRAJECTORY ..................................3-12
TABLE 3-5 LM-3B TYPICAL FLIGHT SEQUENCE...................................................................3-16
TABLE 3-6 CHARACTERISTIC PARAMETERS OF TYPICAL TRAJECTORY ..................................3-18
TABLE 3-7 LM-3BE TYPICAL FLIGHT SEQUENCE ................................................................3-23
TABLE 3-8 LM-3BE CHARACTERISTIC PARAMETERS OF TYPICAL TRAJECTORY.....................3-25
TABLE 3-9 LM-3C TYPICAL FLIGHT SEQUENCE...................................................................3-30
TABLE 3-10 LM-3C CHARACTERISTIC PARAMETERS OF TYPICAL TRAJECTORY.....................3-32
TABLE 4-1 FAIRING TYPE.....................................................................................................4-2
TABLE 4-2 INSERTION LOSS AND TRANSPARENCY RATE FOR THE RF TRANSPARENT WINDOWS.............4-17
TABLE 5-1 IFDS................................................................................................................5-15
TABLE 5-2 UMBILICAL CABLE.............................................................................................5-16
TABLE 5-3 GROUND UMBILICAL CABLE FUNCTIONS.............................................................5-18
TABLE 5-4 GROUND INTERFACES FOR SATELLITE TEST .......................................................5-18
TABLE 5-5 SIGNAL LEVEL INTERFACE BETWEEN RF SYSTEM AND STE..................................5-20
TABLE 6-1 AVERAGE MONTHLY TEMPERATURE IN XSLC........................................................6-1
TABLE 6-2 AVERAGE MONTHLY GROUND WIND SPEED IN XSLC ............................................6-2
TABLE 6-3 AIR-CONDITIONING PARAMETERS INSIDE FAIRING..................................................6-4
TABLE 6-4 CHARACTERISTICS OF ON-BOARD RADIO EQUIPMENT AND GROUND TEST EQUIPMENT............6-5
TABLE 6-5 TYPICAL MAXIMUM STATIC LONGITUDINAL ACCELERATION.....................................6-9
TABLE 6-6 SINUSOIDAL VIBRATION LEVELS...........................................................................6-9
TABLE 6-7 ACOUSTIC NOISE IN THE FAIRING.......................................................................6-10
TABLE 7-1 ROOM AREA AND ENVIRONMENT IN BS2...............................................................7-9
TABLE 7-2 ROOM AREA AND ENVIRONMENT IN BM..............................................................7-16
TABLE 7-3 ROOM AREA AND ENVIRONMENT IN BMX............................................................7-17

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TABLE 7-4 SPECIFICATION OF VEHICLE AND LOADING/UNLOADING EQUIPMENT ......................7-28
TABLE 7-5 SPECIFICATION OF GASES.................................................................................7-31
TABLE 8-1 TYPICAL LAUNCH VEHICLE WORK FLOW (ENCAPSULATION-ON-PAD) AT XSLC ...............8-2
TABLE 8-2 TYPICAL PRE-LAUNCH COUNTDOWN PROCEDURE ON THE LAUNCH DAY ..............8-18
TABLE 10-1 LM-3A LIMIT LOADS........................................................................................10-2
TABLE 10-2 LM-3B/3BE/3C LIMIT LOADS ..........................................................................10-2
TABLE 10-3 SATELLITE MASS AND COG REQUIREMENT.......................................................10-4
TABLE 11-1 MEETING SCHEDULE..................................................................................... 11-14
TABLE 11-2 TYPICAL LAUNCH CAMPAIGN FLOWCHART (ENCAPSULATION-ON-PAD).............. 11-20
TABLE 11-3 LAUNCH SERVICES DOCUMENTATION SUBMISSION DATE................................. 11-26

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ABBREVIATIONS AND ACRONYMS

Acronym
Meaning

Right Ascension of Ascending Node

Argument of perigee
A

A
Ampere(s)
a
Semi-major Axis
A/D
Analog to Digital
ADS
Autonomous Destruction System
AIT
Assembly, Integration and Test
B

BACC
Beijing Aerospace Control Center
BL
Launch Vehicle Processing Building
BM
Solid Propellant Motor Testing and Processing Building
BMX
X-ray Building
BS
SC Test and Fueling Building
BS2
SC Non-Hazardous Operation Building
BS3
SC Hazardous Operation Building

°C
Degree(s) Celsius
CALT
China Academy of Launch Vehicle Technology
CASC
China Aerospace Science and Technology Corporation
CAST
China Academy of Space Technology
CDR
Critical Design Review
CDS
Command Destruction System
CGWIC
China Great Wall Industry Corporation
CLA
Coupled Loads Analysis
CLTC
China Satellite Launch and Tracking Control General
CoG
Center of Gravity
COTE
Check-Out Test Equipment
CS
Command Shutdown
CTS
LM-2C Top Stage
D

D/A
Digital to Analog
dB
Decibel(s)
dBm
Decibel(s) Relative to 1 Miliwatt
dBW
Decibel(s) Relative to 1 Watt
dBpT
Decibel(s) Relative to 1 Picotesla
dBV
Decibel(s) Relative to 1 Microvolt
Deg, °
Degree(s)
Dia
Diameter
E

e
Eccentricity

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Acronym
Meaning
EDC
Effective Date of the Contract
EGSE
Electrical Ground Support Equipment
EIRP
Equivalent Isotropic Radiated Power
EMC
Electromagnetic Compatibility
EMI
Electromagnetic Interference
EPKM
LM-2E Perigee Kick Motor
F

f
Frequency
FER
Fore-positioned Equipment Rooms
FM
Frequency Modulation
FMAR
Final Mission Analysis Review
FMH
Free Molecular Heating
G

g
Acceleration of Gravity
GEO
Geostationary Earth Orbit
GNSS
Global Navigation Satellite System
GTO
Geostationary Transfer Orbit
H

Ha
Height of Apogee
Hp
Height of Perigee
Hr
Hour(s)
Hz
Hertz
I

i
Inclination
ICD
Interface Control Document
IFD
In-Flight Disconnector
IMU
Inertial Measurement Unit
IRD
Interface Requirement Document
J

JSLC
Jiuquan Satellite Launch Center
K

kg
Kilogram(s)
km
Kilometer(s)
kN
Kilo-Newton(s)
kPa
Kilo-Pascal(s)
kW
Kilo-Watt(s)
L

LC
Launch Complex
LCC
Launch Control Center
LEO
Low Earth Orbit
LEOP
Launch and Early Orbit Phase
LH2
Liquid Hydrogen
LM
Long March
LM-3BE
LM-3B Enhanced
LOX
Liquid Oxygen
LRR
Launch Readiness Review
LSC
Launch Services Contract
LV
Launch Vehicle
M

m
Meter(s)
mA
Mili-Ampere(s)

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Acronym
Meaning
ms
Mili-second(s)
Max
Maximum
MCCC
Mission Command and Control Center
MDP
Maximum Dynamic Pressure
MEO
Medium Earth Orbit
min
Minute(s)
MM, mm
Millimeter(s)
MMH
Monomethyl Hydrazine
MRS
Minimum Residual Shutdown
m/s
Meter(s) per second
N

N
Newton(s)
N2
Nitrogen
N2O4
Nitrogen Tetroxide
N/A
Not Applicable
O

Oct.
Octave
P

Pa
Pascal(s)
PCM
Pulse Code Modulation
PFD
Power Flux Density
PLA
Payload Adapter
PLF
Payload Fairing
PMAR
Preliminary Mission Analysis Review
PPB
Payload Preparation Building
PSR
Pre-Shipment Review
PUS
Propellant Utilization System
Q

Q
Quality Factor
R

RAAN
Right Ascension of the Ascending Node
rad
Radian(s)
Rx
Receive
RF
Radio Frequency
RHCP
Right Hand Circular Polarization
rpm
Round(s) per minute
S

s, sec
Second(s)
SAST
Shanghai Academy of Spaceflight Technology
SC, S/C
Spacecraft / Satellite
SGTO
Super Geostationary Transfer Orbit
SOW
Statement of Work
SPM
Solid Propellant Motor
sq
Square
SSO
Sun Synchronous Orbit
STE
Satellite Test Equipment
T

t
Ton(s)
TAS
Thales Alenia Space
TIM
Technical Interchange Meeting
Tx
Transmit

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Acronym
Meaning
TSLC
Taiyuan Satellite Launch Center
TT&C
Telemetry, Tracking and Control
U

UDMH
Unsymmetrical Dimethyl Hydrazine
UPS
Uninterrupted Power Supply
V

V
Volt, Velocity
VEB
Vehicle Equipment Bay
vs
Versus
W

W
Watt(s)
W/m2
Watt(s) Per Square Meter
WSLC
Wenchang Spacecraft Launch Center
X

XSCC
Xi’an Satellite Control Center
XSLC
Xichang Satellite Launch Center

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CHAPTER 1 INTRODUCTION
The development of Long March (LM) launch vehicle family can be traced back to the 1960s.
Up to now, the Long March family of launch vehicles has included the LM-2C Series, the
LM-2D, the LM-4 Series, the LM-3A Series and the LM-2F launch vehicles for LEO, SSO
and GTO missions. With the further expanded launch capability developed over the last few
years, the Long March launch vehicle family has successfully launched a series of manned
spacecraft.
The LM launch vehicle family is shown in Figure 1-1.

Figure 1-1 Long March Family of Launch Vehicles
The LM-1, LM-2, LM-2E, LM-2E/EPKM and LM-3 are no longer in service. Currently, the
LM-3A Series launch vehicles, including LM-3A, LM-3B, LM-3BE and LM-3C, are the
workhorse launch vehicles for China’s GTO launch missions.
The main characteristics of Long March launch vehicle family in service are shown in Table
1-1.

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Table 1-1 The Main Characteristics of Long March Family
Items
LM-2C
LM-2C/
CTS1
LM-2C/
CTS2
LM-2D
LM-4B/
LM-4C
LM-2F
LM-3A
LM-3B
LM-3BE
LM-3C
Height (m)
43.0
43.0
43.0
41.0
48.0
58.3/
52.0
52.5
54.8
56.3
54.8
Lift-off
Mass (t)
245
245
245
250
250
497.9
241
425.8
456
345
Lift-off
Thrust (kN)
2,962
2,962
2,962
2,962
2,962
5,923
2,962
5,923
5,923
4,443
Fairing
Diameter (m)
3.35
3.35
3.35
3.35/3.80
2.90/3.35/
3.80
3.80/
4.20
3.35
4.00
4.00/4.20
4.00
Stage-1 Propellant
N2O4 / UDMH
Stage-2 Propellant
N2O4 / UDMH
Stage-3 Propellant
N/A
Solid
propellant
Solid
propellant
N/A
N2O4 /
UDMH
N/A
LOX / LH2
Main Mission
LEO/SSO
SSO
GTO
LEO/
SSO
SSO
LEO
GTO
Launch Capability
(kg)
3,850/
900
2,100
1,250
4,000/
1,150
2,230/
2,950
8,080/
8,600
2,600
5,100
5,500
3,800
Launch Site
JSLC/
TSLC/
XSLC
JSLC/
TSLC/
XSLC
JSLC/
TSLC/
XSLC
JSLC/
TSLC
JSLC/
TSLC
JSLC
XSLC

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1.1 Long March Family of Launch Vehicles - Development History
The development of the Long March (LM) launch vehicles began in the mid-1960s and has
resulted in the establishment of a family of launch vehicles suitable for a full range of
missions and payloads. The LM launch vehicles have a long list of successful launches and
this record is attributed to the fact that the launch vehicle development has been based on
mature and proven technologies, which includes using the same product at subsystem level,
wherever possible among the members of the LM family. This has resulted in an improved
reliability and a high launch success rate as demonstrated by the whole LM launch vehicle
family. The members of LM family that have been used for international commercial launch
services include LM-2C and its enhanced version LM-2C/CTS1, the LM-2E, the
LM-2E/EPKM, the LM-3, the LM-3B and the LM-3BE.
Figure 1-2 Long March Flights
The LM-1 performed its first flight successfully in April 1970, sending the first Chinese
satellite into a low earth orbit. After the second flight in 1971, the LM-1 was phased out.
The LM-2 is a two-stage launch vehicle developed based on LM-1. Its upgraded version, the
LM-2C successfully made its first flight in November 1975. The enhanced versions of the
LM-2C are designated as the LM-2C/CTS1 and LM-2C/CTS2, which are three-stage launch
vehicles capable of delivering heavier payloads into the orbits that require greater launch
capability. The LM-2C, LM-2C/CTS1 and LM-2C/CTS2 have a record of 32 consecutive
successful flights from the maiden flight in 1975 as of the end of 2010.
The LM-2E and LM-2E/EPKM use the LM-2C as their core stage, around which there are
four strap-on boosters. The LM-2E and LM-2E/EPKM has been phased out since the end of
1995.
The LM-3 is a three-stage launch vehicle, of which the first and second stages were derived

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from the LM-2C. The third stage has a cryogenic engine, using LOX and LH2 as the
propellants, which gives the third stage engine a re-start capability, thus providing greater
flexibility in the mission. The first flight of the LM-3 was in January 1984, and in the middle of
2000 after 13 flights, the LM-3 launch vehicle has been phased out.
The LM-3A launch vehicle is also a three-stage launch vehicle based on the mature
technologies of the LM-3. A newly designed third stage, which also uses LOX and LH2 as
propellants, was used by the LM-3A. The LM-3A third stage can also perform attitude
adjustment maneuvers to orient the payloads and to provide adjustable spin-up operations.
The first LM-3A launch took place in February 1994 and as of the end of 2010 had flown 19
times, all of which were successful.
The LM-3B employs the LM-3A vehicle configuration as the core stage plus four strap-on
boosters on the first stage. The first LM-3B launch was conducted in February 1996, and it
has been launched 10 times as of the end of 2010. Besides the failure of the maiden flight
and the anomaly that occurred in the Palapa-D mission in 2009, all other LM-3B flights were
successful. (In Palapa-D mission, the satellite was sent into a GTO with an apogee altitude
lower than specified due to an anomaly on the third stage engine and as a result the
expected satellite in orbit life was reduced from 15 years to approximately 11 years).
The LM-3BE was developed based on LM-3B with a lengthened first core stage and
strap-on boosters. The launch capability was improved for LM-3BE and the first flight took
place in May 2007. As of the end of 2010, the LM-3BE had made three successful fights.
The LM-3C is a simplified version of LM-3B with only two strap-on boosters on the first stage.
The first LM-3C launch was conducted in April 2008 and there have been six LM-3C
launches as of the end of 2010, all successful.
The LM-2D and LM-4 Series launch vehicles are both developed and manufactured by
Shanghai Academy of Spaceflight Technology (SAST). As of the end of 2010, all the 14
LM-2D flights and 22 LM-4 Series flights are successful.
Please refer to summary Table 1-3 on page 1-10.
1.2 Long March Launch Services
1.2.1
Overview
CGWIC, a wholly-owned subsidiary of CASC, was established in 1980. CGWIC is the sole

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commercial organization authorized by the Chinese government to provide commercial
launch services to international customers. Over the past decades, CGWIC has provided
the international launch services with support from CALT and CLTC. The relationship among
these organizations is illustrated in Figure 1-3. To implement the launch services for the
customer, CGWIC organizes and coordinates the joint team comprising CALT and CLTC for
the execution of the launch service contract. Table 1-2 summarizes the main responsibilities
of these organizations.
China Aerospace Science
& Technology Corp. (CASC)
China Academy of Launch
Vehicle Technology
(CALT)
Subcontractor for
launch vehicle development
China Great Wall
Industry Corp.
(CGWIC)
Prime Contractor

China Satellite Launch &
Tracking Control General
(CLTC)
Subcontractor for launch
operations, TT&C services
Customer

Figure 1-3 The Organization of Long March Launch Services
Table 1-2 Responsibilities of the Launch Services Related Organizations
Subcontractor
Prime Contractor
Subcontractor
CLTC
CGWIC
CALT
 Launch Site Technical
Interface Coordination
 Launch Campaign Planning
& Organization
 Launch Site Operations
 TT&C
 Launch Site & Launch
Safety
 Flight Safety Engineering
 Marketing and Sales of
Long March launch services
 Program management
 Interaction with customers
 Insurance and financing
services on case by case
basis
 Launch vehicle
- Design,
- Development,
- Manufacture, and
- Testing
 Mission Analyses
 LV Technical Interface
Coordination
 LV Quality Control
CALT is one of the main subsidiaries of CASC, and is the leading company engaged in the
development and production of launch vehicles. CALT is located in Beijing and has a total
staff of 25,000, which includes 8,000 engineers and 1,800 senior engineers and scientists.
The LM-1, LM-2, LM-2C Series, LM-3A Series and LM-2F are all designed and

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manufactured by CALT.

Figure 1-4 Long March LV Assembly and Test Workshop
CLTC is the organization that performs the launch operations and provides the launch
facilities including the telemetry and tracking ground station network for the launch. CLTC
manages three independent satellite launch centers, a global TT&C network that includes
tracking ships, and two research institutes. CLTC has a team of 5,000 technical staff to
support the launch and tracking operations.
1.2.2
Customer Interfaces
CGWIC is the single point of contact with the customer. When a launch services contract is
signed, CGWIC nominates a Program Manager who is responsible for implementing the
launch services contract.
1.3 Overview of the Launch Centers
There are three launch centers in China, i.e. Xichang Satellite Launch Center (XSLC),
Jiuquan Satellite Launch Center (JSLC) and Taiyuan Satellite Launch Center (TSLC). Figure
1-5 shows their locations. Wenchang Spacecraft Launch Center (WSLC) in Hainan Province
is under construction.

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Figure 1-5 Satellite Launch Centers
1.3.1 Xichang Satellite Launch Center (XSLC)
XSLC is located in Sichuan Province, southwestern China, and is primarily used for GTO
missions. The launch center comprises processing buildings for the pre-launch preparation
of the launch vehicles, satellites and solid motors as well as the ground support equipment.
XSLC has two launch complexes (LC), one of which (LC-3) is primarily used for LM-3A
launch operations and the other one (LC-2) is for LM-3B, LM-3BE and LM-3C launch
operations.

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Figure 1-6 Xichang Satellite Launch Center 1.3.2 Taiyuan Satellite Launch Center (TSLC) TSLC is located in Shanxi Province, northern China, and is primarily used for LEO missions utilizing the LM-2C Series and SSO missions utilizing the LM-4 Series.

Figure 1-7 Taiyuan Satellite Launch Center

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1.3.3 Jiuquan Satellite Launch Center (JSLC)
JSLC is located in Gansu Province, northwestern China. It is the first launch site in China. It
is mainly used for LEO launch missions and MEO launch missions using the LM-2C Series,
LM-2D and LM-4 Series launch vehicles. It is also used for China’s manned space missions
with the LM-2F launch vehicles.

Figure 1-8 Jiuquan Satellite Launch Center

1.4 Launch Record of Long March
The Long March series of launch vehicles has been undergoing continuous development in
China since the mid-1960s. Different launch vehicles with various capabilities have been
developed, which has enabled China's launch business to enter into both domestic and
international markets and to accomplish different LEO, GTO and SSO satellite launch
missions.
The Long March launch vehicle spectrum and their track record are summarized in Table
1-3 below:

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Table 1-3 Flight Record of Long March (As of January 1, 2011)

Launch Vehicle
Success/
Flight
Total Success/
Total Flight
Status
LM-1
02/02
LM-2
00/01
LM-2E
03/05
LM-4A
02/02
07/10
Out of Service
LM-2C & LM-2C/CTS
32/32
LM-2D
14/14
LM-4B & LM-4C
20/20
LEO & SSO Missions
LM-2F
07/07
73/73
In Service
LM-2E/EPKM
02/02
LM-3
10/13
12/15
Out of Service
LM-3A
19/19
LM-3B
11/13
GTO Missions
LM-3C
06/06
36/38
In Service
Total
128/136

1.5 Long March Launch Services Advantages
The selection of the Long March launch services provides the customer with some key
advantages:
a)
Long March launch vehicles are flight proven, highly reliable and compatible with all
major satellite platforms.
b)
The system, sub-systems and major equipment production of the Long March family is
wholly controlled within CASC group, which facilitates the launch vehicle hardware and
software development process.
c)
The Long March launch is conducted on a dedicated basis with no risk of launch delay
by co-passengers and scheduling flexibility can be arranged to the customer’s benefit.
d)
As it is a dedicated launch, any residual margin in the launch capability can be used to
optimize the injection parameters in order to extend the satellite’s in-orbit life. This is
without any additional cost to the customer. If the additional margin in launch vehicle

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capability is identified early in the launch services program, it could provide the
customer with greater flexibility in the satellite design.
e)
The state-of-the-art ground support facilities are purpose built, and the satellite
processing facility at the launch site is handed over to the customer for the duration of
the launch campaign.
f)
The security of the customer’s satellite technology is regarded as paramount and all
necessary security measures are implemented and reviewed with the customer. The
security measures include an exemption for the customer’s satellite from the Chinese
customs inspection and the full control of access to the satellite processing facility by
the customer throughout the launch campaign.
g)
CGWIC procures the launch vehicles in bulk which allows the lead time to meet the
customers launch schedule requirements. The launch vehicle is subject to
comprehensive tests and checkouts at the launch vehicle manufacturing facility before
shipment to the launch site, which allows the launch vehicle processing at the launch
site to be streamlined for completion in approximately 24 days. This provides additional
schedule flexibility for the launch vehicle up to combined operations. In addition, the use
of two launch complexes and two satellite processing facilities makes it possible for two
launch campaigns to be conducted in parallel, which further increases the schedule
flexibility of Long March launch services.
h)
CGWIC has full support from the Chinese government for the provision of launch
services to international customers, and there are no obstacles for CGWIC in obtaining
the Chinese government’s approval to perform the launch services for an international
customer.
i)
A fully equipped modern hotel is located at the launch site providing comfortable living
conditions and easy access to the processing facilities.
j)
CGWIC demonstrated its abilities and proved its competence in over 20 years of
space-related activities and has a dynamic and experienced engineering and
management team.

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CHAPTER 2 GENERAL DESCRIPTION
2.1 Summary
The China Academy of Launch Vehicle Technology (CALT) started to develop the LM-3A
Series launch vehicles in 1986, and they are three-stage liquid propellant launch vehicles
used mainly for Geostationary Transfer Orbit (GTO) missions. LM-3A Series launch vehicles
include the LM-3A, LM-3B, LM-3B Enhanced (LM-3BE) and LM-3C. The configuration of
LM-3A Series launch vehicles is shown in Figure 2-1. The GTO capability of LM-3A is 2,600
kg.
The LM-3B is a three-stage launch vehicle, which utilizes the proven LM-3A as its core stage
and has been upgraded by the addition of four liquid strap-on boosters to the first stage. The
second stage tanks were also increased in volume, which increased the GTO capability up
to 5,100 kg for the LM-3B version.
The LM-3BE is based on the LM-3B but has a lengthened first stage and boosters, which
increases the GTO launch capability up to 5,500 kg.
The LM-3C is a three-stage launch vehicle but with two liquid strap-on boosters. The LM-3C
variant is capable of delivering a satellite with the mass of 3,800 kg into GTO. The LM-3C
and LM-3B variants were developed in parallel using the same modular design as the
LM-3A. The LM-3C uses the same SC/LV interface and payload fairing as is used by the
LM-3B.
All the LM-3A Series launch vehicles are flight proven.

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Figure 2-1 LM-3A/3B/3BE/3C Configuration
2.2 LM-3A Launch Vehicle
2.2.1 Technical Parameters
LM-3A is 52.52 m long with lift-off mass of 242 tons and uses a fairing that is 8.887 m long
with diameter of 3.35 m.

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Figure 2-2 LM-3A Launch Vehicle

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Technical Parameters of the First Stage:
Length
23.272 m
Diameter
3.35 m
Propellant
N2O4 / UDMH
Mass of Propellant
171,800 kg
Engine
YF-21C
Ground Thrust
2,961.6 kN
Ground Specific Impulse
2,556.5 N·s/kg
Technical Parameters of the Second Stage:
Length
11.276 m
Diameter
3.35 m
Propellant
N2O4 / UDMH
Mass of Propellant
32,600 kg
Engine
YF-24E
Thrust ( in vacuum)
Main Engine: 742 kN

Vernier Engine: 47.1 kN
Specific Impulse (in vacuum)
Main Engine: 2,922.57 N·s/kg

Vernier Engine: 2,910.5 N·s/kg
Technical Parameters of the Third Stage:
Length
12.375 m
Diameter
3.0 m
Propellant
LOX / LH2
Mass of Propellant
18,200 kg
Engine
YF-75
Thrust (at vacuum)
167.17 kN
Specific Impulse (in vacuum)
4,295 N·s/kg
2.2.2 LM-3A Description
The LM-3A launch vehicle consists of the following major subsystems: vehicle structure,
propulsion system, control system, measurement system (telemetry system and tracking &
range safety system), propellant management system & reaction control system, propellant
utilization system, separation system and auxiliary system.
Vehicle Structure
The function of the vehicle structure is to withstand the internal and external loads on the
launch vehicle during ground transportation, hoisting and flight, in addition to housing all the
sub-systems. The vehicle structure comprises the first stage, second stage, third stage and
payload fairing.

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Figure 2-3 CALT Assembly Area
The first stage includes inter-stage section, oxidizer tank, inter-tank section, fuel tank, rear
transit section, tail section, engine, valves and pipes.
The second stage is comprised of oxidizer tank, inter-tank section, fuel tank, and engine.
The third stage includes the Payload Adapter (PLA), vehicle equipment bay (VEB),
cryogenic propellant tanks and engine. The PLA mates the satellite with the LM-3A and
transfers the loads between them. The PLA can be one of the international standard
interfaces designated as 937B, 1194 or 1194A. The VEB is a circular structure made of
metal honeycomb and trusses (see Figure 2-4), which is mounted on top of the propellant
tanks. The propellant tank of third stage is thermally insulated with a common bulkhead,
convex upward in the middle. The common bulkhead is thermally insulated with dual-layer
honeycomb vacuum thermal insulation. The tanks are arranged with the liquid hydrogen
(LH2) in the upper tank and liquid oxygen (LOX) in the lower tank.
The payload fairing consists of the dome, forward cone section, cylindrical section, reverse
cone section, and fairing separation mechanisms (see Chapter 4 for details).

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Figure 2-4 LM-3A VEB Configuration
Propulsion System
The propulsion system, which consists of the engines and pressurization system, generates
the flight thrust and vehicle control forces.
There are four engines symmetrically installed in the first stage, each of which can swing in
tangentially to the thrust vector. The propellant tanks are pressurized by the autogenous
pressurization systems.
There are one main engine and four vernier engines on the second stage, and as with the
first stage the propellant tanks are pressurized by the autogenous pressurization systems.
The first stage and second stage employ storable propellants, nitrogen tetroxide (N2O4) and
unsymmetrical dimethyl hydrazine (UDMH).

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The third stage is equipped with two universal gimbaled cryogenic engines with restart
capability. The propellants are liquid hydrogen (LH2) and liquid oxygen (LOX). The LH2 tank
is pressurized with helium by the autogenous pressurization system, and the LOX tank is
pressurized with heated helium by the autogenous pressurization system.

Figure 2-5 Second Stage Engines

See Figure 2-6a, Figure 2-6b and Figure 2-6c for propulsion systems of the first, second and
third stages.

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Figure 2-6a Schematic Diagram of Stage-1 Propulsion System

Figure 2-6b Schematic Diagram of Stage-2 Propulsion System

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Figure 2-6c Schematic Diagram of Stage-3 Propulsion System

Control System
The function of control system is to maintain the flight stability, perform the navigation and
guidance functions to deliver the satellite into the predetermined orbit. The control system
uses a four-axis inertial platform and laser IMU, combined with computer-controlled
guidance and digital attitude control. The control system uses advanced digital technologies
and redundancy design to provide enhanced reliability and flexibility for a wide range of
missions for the LM-3A. Figure 2-7a, Figure 2-7b and Figure 2-7c illustrate the schematic
diagrams of the LM-3A Series launch vehicles control system.

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Figure 2-7a Schematic Diagram of LM-3A Guidance & Control System

Figure 2-7b Schematic Diagram of LM-3A Attitude Control System

Figure 2-7c Schematic Diagram of LM-3A Guidance System

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Measurement System
The measurement system includes the telemetry system and the tracking and range safety
system.
(a) Telemetry System
The functions of telemetry system are to measure and transmit parameters of the launch
vehicle in flight. Some measured data can be processed in real time. The data acquisition
and encoding units in the telemetry system are powered in group based on sensors’ location.
The measurements of the command signals are digitized. The powering and check-up are
performed automatically. The on-board digital converters are intelligent. Totally about 560
parameters are measured in flight. Refer to Figure 2-8 for the diagram of LM-3A Telemetry
System.

Figure 2-8 Schematic Diagram of LM-3A Telemetry System
(b) Tracking and Range Safety System
The tracking and range safety system is used to measure the trajectory data and final
injection parameters, and provide safety assessment information. Self-destruction or
destruction by remote control will be conducted if flight anomaly occurs. See Figure 2-9.

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Figure 2-9 Schematic Diagram of LM-3A Tracking & Range Safety System
Propellant Management and Attitude Control System
This system is used for the attitude control and propellant management during the coast
phase of the third stage, final velocity adjustment after engines shutdown and attitude
reorientation prior to satellite separation. It adopts a set of reaction control thrusters fueled
by pressure-fed hydrazine, which can be started for multiple times according to the
command. See Figure 2-10 for the system diagram.

Figure 2-10 LM-3A Propellant Management & Attitude Control System
Propellant Utilization System
The propellant utilization system on third stage is to measure in real time the level of
propellants inside the third stage tanks and to adjust the consuming rate of liquid oxygen
(LOX) to achieve the residual propellants in an optimum proportion. The adjustment of

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propellant mixture ratio is used to compensate the deviation due to engine performance and
propellant loading to enhance the launch capability. The system contains a processor,
propellant level sensors and adjusting valves. Refer to Figure 2-11 for the system
configuration.

Figure 2-11 LM-3A Series Launch Vehicles Propellant Utilization System
Separation System
Separation system is to unlock the connecting mechanism and separate the connected
sections. For LM-3A, four separations take place in sequence during the flight phase, i.e.
first/second stage separation, fairing jettisoning, second/third stage separation and SC/LV
separation as illustrated in Figure 2-12.
(a) First/Second Stage Separation
First/second stage separation employs “hot separation”, which means the first stage is
pushed away by the jet of the second stage engine after explosive bolts connecting the first
stage and the second stage are unlocked.
(b) Fairing Jettisoning
During payload fairing separation, lateral explosive bolts are unlocked first and then followed

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by longitudinal explosive bolts, and finally the fairing turns outward around the hinges under
the force of separation springs.
(c) Second/Third Stage Separation
The second/third stage separation is a “cold separation”. The explosive bolts are unlocked
first and then the retro-rockets on the second stage generate separation force and push the
second stage away.
(d) SC/LV Separation
The SC is bound together with the LV through a clampband tightened by two explosive bolts.
The clampband is released immediately after either one or both of the two explosive cutters
are ignited. Then the separation springs provide separation force, pushing satellite and the
third stage away from each other.

Figure 2-12 LM-3A/3B/3C Separation Events
Auxiliary System
In addition to the main subsystems as mentioned above, there are auxiliary systems to
perform necessary functions before the launch vehicle lift-off, which includes measuring the

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propellant level and temperature, air-conditioning, water-proofing and dehumidifying for
payload fairing.
2.3 LM-3B and LM-3BE Launch Vehicles
2.3.1 Technical Parameters
The total lift-off mass for the LM-3B and LM-3BE is 426 tons and 456 tons respectively. The
total length is 54.838 m ~ 57.056 m depending on the type of payload fairing to be used and
fairing encapsulation method selected.

Figure 2-13 LM-3B Launch Vehicle
There are four types of payload fairing available for launch services with the nominal
diameter of 3.7 m (for dual launch), 4.0 m and 4.2 m respectively. The option selected is
dependent on whether the fairing is encapsulated on the pad or in BS3 (SC Hazardous
Operation Building). The two encapsulation ways are as follows:
Encapsulation-on-Pad
When the satellite is encapsulated on the pad, the satellite and fairing are transported to
launch pad separately; then the fairing is encapsulated following the mate of satellite to the
launch vehicle.

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Encapsulation-in-BS3
When the satellite is encapsulated in BS3, the satellite is mated to the Payload Adapter
(PLA) and encapsulated in the fairing in BS3. The encapsulated satellite is shipped to the
launch pad in the fairing and the complete assembly is mated to the launch vehicle on the
launch pad.
Technical Parameters of Booster:

LM-3B
LM-3BE
Length
15.326 m
16.094 m
Diameter
2.25 m
2.25 m
Propellant
N2O4 / UDMH
N2O4 / UDMH
Mass of Propellant
37,700 kg 41,100 kg
Engine Designation
YF-25
YF-25
Ground Thrust
740.4 kN
740.4 kN
Ground Specific Impulse
2,556.2 N·s/kg
2,556.2 N·s/kg
Technical Parameter of First Stage:

LM-3B
LM-3BE
Length
23.272 m
24.76 m
Diameter
3.35 m
3.35 m
Propellant
N2O4 / UDMH
N2O4 / UDMH
Mass of Propellant
171,800 kg
186,200 kg
Engine
YF-21C
YF-21C
Ground Thrust
2,961.6 kN
2,961.6 kN
Ground Specific Impulse
2,556.5 N·s/kg
2,556.5 N·s/kg
Technical Parameter of Second Stage:
Length
12.92 m
Diameter
3.35 m
Propellant
N2O4 / UDMH
Mass of Propellant
49,400 kg
Engine
YF-24E
Thrust (in vacuum)
Main Engine: 742 kN

Vernier Engine: 47.1 kN
Specific Impulse (in vacuum)
Main Engine: 2,922.57 N·s/kg

Vernier Engine: 2,910.5 N·s/kg
Technical Parameter of Third Stage
Length
12.375 m
Diameter
3.0 m
Propellant
LH2 / LOX
Mass of Propellant
18,200 kg
Engine
YF-75
Thrust (in vacuum)
167.17 kN
Specific Impulse (in vacuum)
4,295 N·s/kg

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2.3.2 LM-3B and LM-3BE Description
The LM-3B and LM-3BE launch vehicles are comprised of the vehicle structure, propulsion
system, control system, measurement system (telemetry system and tracking & range
safety system), propellant management and reaction control system, propellant utilization
system, separation system and auxiliary system, etc.
Vehicle Structure
The vehicle structure consists of the core stage and boosters, with the core stage being
similar to that of LM-3A but with an enhanced tank structure to meet the flight load
requirements of LM-3B. The tanks of the second stage have been extended by a total of
1.65 m to allow more propellants to be loaded (See Figure 2-1).
The booster consists of nose, oxidizer tank, inter-tank section, fuel tank, rear transit section,
tail section, fins, engine, valves and pipes, etc.
The third stage includes the Payload Adapter (PLA), vehicle equipment bay (VEB),
cryogenic propellant tanks and engines. The PLA mates the satellite to LM-3B and bears the
mechanical loads. The PLA can be one of the international standard interfaces designated
as 937B, 1194 or 1194A. The satellite can be encapsulated on the pad or in BS3 (See Figure
2-14 and Figure 2-15).
The payload fairings consist of the dome, bi-conic section, cylindrical section, and reverse
cone section and separation mechanisms. See Chapter 4 for details.

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Figure 2-14 LM-3B & LM-3C Fairing for Encapsulation-on-Pad

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Figure 2-15 LM-3B & LM-3C Fairing for Encapsulation-in-BS3
Propulsion System
The propulsion system consists of the engines and pressurization system, and generates
the flight thrust, which includes the dynamic control system and is identical to that of the
LM-3A.
There are four strap-on boosters attached to the first stage, each equipped with a single
engine that is identical to that of the core stage and each booster has an autogenous
pressurization system.
Engine systems for the first stage, second stage and third stage are identical to those of
LM-3A, with one main engine and four vernier engines on the second stage and two
universal gimbaled cryogenic engines on the third stage. The schematic diagrams of the
propulsion system are shown in Figure 2-6a, Figure 2-6b and Figure 2-6c.

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Control System
The function and composition of the Control System is basically similar to that of the LM-3A,
except that an additional control circuit has been added to control the engine ignition,
shutdown and separation of the strap-on boosters. See Figure 2-7a, Figure 2-7b and Figure
2-7c.
Measurement System
(a) Telemetry System
The telemetry system is basically similar to that of LM-3A with telemetry circuit added to
measure parameters of the boosters. There are altogether about 800 parameters measured
in flight. See Figure 2-8.
(b) Tracking and Range Safety System
The Tracking and Range Safety System has the similar function and composition to that of
LM-3A. See Figure 2-9.
Propellant Management and Attitude Control System during Coast Flight
There is a Propellant Management and Attitude Control System on the third stage. The
system has the same function and composition with those of LM-3A. To meet the
requirements of attitude control of LM-3B, two small reaction nozzles each with thrust of 40N
are added for pitch and yaw. The thrust of propellant settling nozzle is increased from 40N in
LM-3A to 45N in LM-3B /3C due to the mass increase during the coast flight phase of
LM-3B/3C. See Figure 2-16.

Figure 2-16 LM-3B/3BE/3C Propellant Management & Attitude Control System

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Separation Systems
There are five separation actions during the flight of LM-3B: booster separation, first/second
stage separation, fairing jettisoning, second/third stage separation and SC/LV separation as
illustrated in Figure 2-12.
(a) Booster Separation.
The boosters are “thrown-away” upon separation. Before separation, boosters are mounted
to the core stage through three pairs of explosive bolts at the front section and a separation
nut at the rear section. Four small separation nozzles generate outward separation force to
push the boosters away from the first stage following the simultaneous unlocking of the
explosive bolts and the separation nut.
(b) First/Second Stage Separation
The same as that of LM-3A.
(c) Fairing Jettisoning
During the payload fairing separation, the lateral explosive bolts unlock first and then the
longitudinal detonation cord and explosive bolts unlock. The fairing turns outward around
the hinges under the force of separation springs.
(d) Second/Third Stage Separation
The same as that of LM-3A.
(e) Satellite Separation System
The same as that of LM-3A.
Propellant Utilization System and Auxiliary System
The propellant utilization system is identical to that of LM-3A. See Figure 2-11. The auxiliary
system is same as that of LM-3A.
2.4 LM-3C Launch Vehicle
Total lift-off mass of the LM-3C (Figure 2-17) is approximately 345 tons and its overall length
is 54.838 m with a fairing of 4 m in diameter (Encapsulation-on-Pad). The fairings available
and their encapsulation methods are the same as that of LM-3B. See Section 2.3 covering
the LM-3B for details.

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The function and composition of the LM-3C launch vehicle is the same as the LM-3B with
the exception that only two boosters are attached on the first stage.

Figure 2-17 LM-3C Launch Vehicle
2.5 Launch Vehicle Coordinate System
The origin (O) of coordinate (OXYZ) of the Launch Vehicle is located at the instantaneous
mass center, i.e. the integrated mass center of SC/LV combination including the Payload
Adapter (PLA), propellants and Payload Fairing (PLF), etc. The OX axis coincides with the
longitudinal axis of the launch vehicle. The OY axis is perpendicular to the OX axis and falls
in the plane of the launching direction. The OX, OY and OZ axes form a right-handed
orthogonal system. The flight attitude of the launch vehicle axis is defined as shown in
Figure 2-18. The satellite manufacturer will define the spacecraft coordinate system, and the
relationship or clocking orientation between the launch vehicle and spacecraft coordinate
systems will be determined through technical coordination meetings for the specific project.

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Figure 2-18 Coordinate System of LM-3A/3B/3BE/3C
2.6 Launch Missions of LM-3A Series Launch Vehicles
LM-3A Series launch vehicles are primarily used for GTO missions, although LEO and SSO
missions can also be conducted if required by customers.
GTO Missions
The main mission for the LM-3A Series launch vehicles is to deliver the satellite into
Geostationary Transfer orbit (GTO). The transfer of the satellite from GTO to the
Geostationary Earth Orbit (GEO) will be conducted by the engine carried on the satellite.
GTO refers to an elliptical orbit with an apogee altitude equal or even higher than the altitude

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of GEO. GEO is the satellite working orbit, on which the SC has the same orbital period as
the rotation period of the Earth, with orbit plane coinciding with the equatorial plane of the
Earth. Typical GTO mission profiles for the LM-3B/3BE and LM-3C launch vehicles are
shown in Figure 2-19.

Figure 2-19 Schematic Drawing of GTO and GEO
LEO Missions
The LM-3A Series launch vehicles can also be used to inject satellite into low earth orbit
(LEO), which refers to orbits with average altitude lower than 2,000km.
The LM-3A Series launch vehicles can also be used to inject satellite into sun synchronous
orbit (SSO). SSO refers to an orbit that maintains a fixed angle with respect to the earth –
sun direction. That is, the orbital plane has a fixed orientation with respect to the earth – sun
direction and the angle between the orbital plane and the earth – sun direction is constant.
Deep Space Missions
The LM-3A Series launch vehicles can also be used to launch space probes into escape
orbit beyond the earth gravitational field, for example for mission to the Moon and Mars.
2.7 Satellites Launched
LM-3A Series launch vehicles have conducted a total of 38 launches as of January 1, 2011.
See Table 2-1 for details.

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Table 2-1 SC launched by LM-3A Series LV (As of January 1, 2011)
Designed Orbit
LV
SC
Manufacture
SC Name
Launch Date
i (deg)
Hp (km)
Ha (km)
SC Mass
(Kg)
F01
CAST/CALT
SJ-4/KF-1
1994.02.08
28.5
200
36,311
396/1,342
F02
CAST
ChinaSat-5
1994.11.30
28.5
200
36,185
2,232
F03
CAST
ChinaSat-6
1997.05.12
28.5
200
36,065
2,267
F04
CAST
ChinaSat-22
2000.01.26
25.0
200
41,991
2,320
F05
CAST
BD-1A
2000.10.31
25.0
200
41,991
2,315
F06
CAST
BD-1B
2000.12.21
25.0
200
41,991
2,311
F07
CAST
BD-1C
2003.05.25
25.0
200
41,991
2,325
F08
CAST
ChinaSat-20
2003.11.15
25.0
200
41,991
2,307
F09
SAST
FY-2C
2004.10.19
27.1
280
36,056
1,400
F10
CAST
ChinaSat-22A
2006.09.13
25.0
200
41,991
2,330
F11
SAST
FY-2D
2006.12.08
24.9
200
36,407
1,390
F12
CAST
BD-1D
2007.02.03
25.0
200
41,991
2,299
F13
CAST
Compass-01
2007.04.14
55.0
200
21,650
2,320
F14
CAST
SinoSat-3
2007.06.01
25.0
200
41,991
2,320
F15
CAST
CE-1
2007.10.24
31.0
200
51,000
2,352
F16
SAST
FY-2E
2008.12.23
24.1
200
35,700
1,390
F17
CAST
Compass-05
2010.08.01
55.0
200
35,991
2,300
F18
CAST
ChinaSat-20A
2010.11.25
25.0
200
41,991
2,325
LM-3A
F19
CAST
Compass-07
2010.12.18
55.0
200
35,991
2,300
F01
SS/Loral
Intelsat-708
1996.02.15
24.5
200
35,786
4,594
F02
SS/Loral
MABUHAY
1997.08.20
24.5
200
47,924
3,775
F03
SS/Loral
APSTAR-2R
1997.10.17
24.5
200
47,924
3,746
F04
Lockheed Martin
ChinaStar-1
1998.05.30
24.5
200
85,000
2,917
F05
Aerospatial
SinoSat-1
1998.07.18
19.0
600
35,786
2,832
F06
Alcatel Space
APSTAR-VI
2005.04.12
26.0
200
50,000
4,576
F07
CAST
SinoSat-2
2006.10.29
28.5
200
35,786
5,107
F08*
CAST
NigComSat-1
2007.05.14
25.2
200
41,991
5,086
F09
TAS
ChinaSat-6B
2007.07.05
24.3
200
50,289
4,516
F10
TAS
ChinaSat-9
2008.06.09
24.2
200
50,289
4,466
F11*
CAST
VeneSat-1
2008.10.30
24.8
200
41,991
5,050
F12
TAS
Palapa-D
2009.08.31
21.2
200
50,291
4,100
LM-3B / LM-3BE (marked with *)
F13*
CAST
SinoSat-6
2010.09.06
25.2
200
41,991
5,100
F01
CAST
TL-1
2008.04.25
18.0
200
41,991
2,475
F02
CAST
Compass-02
2009.04.15
20.5
200
35,974
3,060
F03
CAST
Compass-03
2010.01.17
20.5
200
35,974
3,060
F04
CAST
Compass-04
2010.06.02
20.5
200
35,974
3,060
F05
CAST
CE-2
2010.10.01
28.5
200
364,160
2,480
LM-3C
F06
CAST
Compass-06
2010.11.01
20.5
200
35,974
3,060

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2.8 Future Enhancements for LM-3A Series Launch Vehicles
The following enhancements are being planned for the LM-3A Series launch vehicles:
a)
To further improve the reliability and robustness of LM-3A Series launch vehicles.
b)
To study the feasibility to improve the launch capability of LM-3B to 6 tons for the
standard GTO mission.
c)
To introduce a newly developed upper stage, so that payloads can be launched into
10,000~20,000 km circular orbit or directly injected into GEO. The new upper stage has
the capability to readjust attitude and reorient so that multiple satellites can be launched
with one launch vehicle. Launch vehicles for deep space exploration will be developed
on the basis of LM-3A Series launch vehicles, which are the basic carriers for space
exploration flight in China to send the payloads into transfer orbit to the Moon or Mars.
d)
Please refer to our web site for announcements and updates.

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CHAPTER 3

LAUNCH PERFORMANCE AND INJECTION ACCURACY
3.1 Summary
The LM-3A Series launch vehicles’ performance data given in this Chapter are based on the
following conditions:
a)
The launch is from XSLC, with initial launch azimuth 97.5 deg for the LM-3B, LM-3BE
and LM-3C, 104 deg for the LM-3A.
b)
In the process of launch orbit design the following requirements have been taken into
consideration: limitations of the impact area, range safety and flight tracking.
c)
The mass of the Payload Adapter (PLA) is included in the third stage. So the launch
capability refers to the mass above the SC/LV interface.
d)
The third stage of LM-3A Series launch vehicles is loaded with sufficient propellant to
reach the predetermined orbit with a probability of no less than 99.73%.
e)
At fairing jettisoning, the aerodynamic flux by free molecular is no greater than 1,135
W/m2.
f)
Radius of the earth equator is 6,378.14 km.
3.2 Standard Mission Profile
3.2.1 Coordinates of Launch Pads
The LM-3A Series launch vehicles are launched from Launch Complex No.2 (LC-2) (for
LM-3B, LM-3BE and LM-3C) and Launch Complex No.3 (LC-3) (for LM-3A) of Xichang
Satellite Launch Center (XSLC), the geographic coordinates of LC-2 are as follows:
Latitude
28.25 deg N
Longitude
102.025 deg E
Elevation
1,825 m
The geographic coordinates of the LC-3 are as follows:
Latitude
28.25 deg N
Longitude
102.029 deg E
Elevation
1,825 m

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The launch coordinate system for LM-3A, LM-3B, LM-3BE and LM-3C is illustrated in Figure
3-1, Figure 3-2 and Figure 3-3.

Figure 3-1 Launch Azimuth of LM-3A

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Figure 3-2 Launch Azimuth of LM-3B and LM-3BE

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Figure 3-3 Launch Azimuth of LM-3C
3.2.2 Geostationary Transfer Orbit Missions and Other Orbit Missions
The delivery of a satellite to GTO is the prime mission for LM-3A Series launch vehicles. The
mission can be an injection to a standard GTO or a GTO tailored to the customer
requirements. The standard GTO is the baseline mission with following injection parameters
from XSLC.
Perigee Altitude
Hp
200 km
Apogee Altitude
Ha
35,958* km
Inclination
i
28.5 Deg
Perigee Argument
ω
179.6 Deg
Note: * The parameters in the table represent the SC’s immediate orbit upon separation,
where Ha corresponds to the altitude of 35,786km when the SC arrives at its first apogee
after a period of unpowered flight, taking into account the perturbation caused by the earth’s
oblateness.

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LM-3A Series launch vehicles can inject the satellite into a Super Geostationary Transfer
Orbit (SGTO). It results in a smaller delta velocity required for satellite to reach the final orbit
location consuming less propellant, which will increase the satellite in-orbit life.
The LM-3A Series launch vehicles can be also used for special missions tailored to the
customers’ specific requirements, which include SSO missions, LEO missions, MEO
missions and deep space exploration missions. In addition, dual-launch and multiple-launch
missions can also be performed with the LM-3A Series launch vehicles.
3.2.3 GTO Mission Injection Optimization
The third stage of LM-3A Series launch vehicle operates in either of the following two
shutdown modes, Command Shutdown (CS) and Minimum Residual Shutdown (MRS).
The Command Shutdown (CS) means that the third stage of LM-3A Series launch vehicle
carries sufficient propellant allowing the satellite to enter the predetermined orbit with
probability no less than 99.73%, which is the standard mission profile for the LM-3A Series
launch vehicles.
The Minimum Residual Shutdown (MRS) means that the propellants in the third stage are
burned to minimum acceptable residual level, which allows a significant increase in the
nominal performance with a decrease in injection accuracy. The third stage of LM-3A Series
launch vehicle has a Propellant Utilization System (PUS), which can adjust the mixture ratio
to optimize the consumption of propellant and assure the reliability under Minimum Residual
Shutdown (MRS).
The launch capability defined in this User's Manual are based on the Command Shutdown
(CS) mechanism unless stated otherwise.
3.2.4 Minimum Residual Shutdown (MRS) Missions
The MRS mission will provide a higher apogee for the satellite transfer orbit upon separation
from launch vehicle, resulting in a reduction of the propellant used during the Launch and
Early Orbit Phase (LEOP), which increases the satellite operational life. The MRS mission
allows heavier payloads to be launched and the mission is therefore optimized to maximize
the satellite life, but with a relative reduction in injection accuracy.
Customers can opt for a trade-off between satellite mass, injection altitude and injection
accuracy with the MRS offering the biggest lift-off mass. The relationship between CS
probability up to MRS and the corresponding gains of launch capability are shown in Table

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3-1.
Table 3-1 Launch Capability vs. Command Shutdown Probability
Command Shutdown
Probability
Gains of Launch
Capability (kg)
99.7%
0
95.5%
33
68.3%
67
50.0%
78
3.2.5 Orbit Injection Accuracy
The orbit injection accuracy and covariance matrix for standard GTO mission is shown in
Table 3-2a and Table 3-2b.
Table 3-2a Injection Accuracy for Standard GTO Mission (1σ)
Symbol
Parameters
Deviation
a

Semi-major Axis
40 km
i
Inclination
0.07°
ω
Perigee Argument
0.20°
Ω
Right Ascension of Ascending Node
-0.20° *
Hp

Perigee Altitude
10 km
Note: * the error of launch time is not considered in determining Ω

Table 3-2b Covariance Matrix of Injection for Standard GTO Mission

a
e
i
ω
Ω

Semi-major
axis
Eccentricity
Inclination
Argument of
perigee
Right
Ascension of
Ascending node
a
1524
0.02492
0.5266
3.2344
-0.09688
e

0.52706E-6
0.8615E-5
0.6146E-4
0.5314E-8
i

0.4752E-2
0.1237E-3
-0.4212E-2
ω

0.03897
-0.01780
Ω

0.03927
3.2.6 Pointing Accuracy
The attitude control system on the third stage initiates a reorientation maneuver for the
SC/LV combination to the required orientation from 20 seconds after the final third stage
shutdown to separation. It takes 80 seconds to complete the reorientation maneuver. The
attitude accuracy following the maneuver up to separation of the satellite is less than 2 deg.

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The pointing requirements are defined by the customer with reference to the perigee
coordinates system (OUVW).
The perigee coordinate system (OUVW) is defined as follows:
OU is a radial vector with the origin at the earth center, pointing to the intended
perigee.
OV is perpendicular to OU in the intended orbit plane and points to the intended
direction of the perigee velocity.
OW is perpendicular to OV and OU, and OUVW forms a right-handed orthogonal
coordinate system.
The coordinate system is shown in Figure 3-4.

Figure 3-4 Perigee Coordinate Systems
3.2.7 Separation and Spin-up Accuracy
For missions where the satellite does not require spin-up before separation, the
post-separation pointing parameters are as follows: if the lateral angular rate ω <1 deg/s the
instant deviation of geometry axis is δX<3 deg.
For missions where the satellite requires a spin rate between 5 rpm to 10 rpm along the
launch vehicle X axis after separation, the post-separation pointing parameters are as
follows: the deviation of longitudinal angular rate is ±0.6 rpm, and if the lateral angular rate
ω<2.5 deg/s, the deviation of angular momentum pointing direction is δH<8 deg.
For missions where the satellite needs a spin rate below 3 deg/s around the satellite lateral
axes after separation, the post-separation pointing parameters are as follows: if the lateral

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angular rate ω<0.7 deg/s, the deviation of angular momentum pointing direction δH<15 deg.
The parameters for separation are shown in Figure 3-5.

Figure 3-5 Definition of Separation Accuracy
3.2.8 Launch Windows
3.2.8.1 Definitions
The following definitions apply to the launch windows for each of the LM-3A Series launch
vehicles.
Launch Period
The Launch Period is a period of three consecutive calendar months during which the
customer wants to launch their satellite and there is a launch opportunity on each day of the
period.
Launch Slot
The Launch Slot has a duration of one calendar month that falls within the Launch Period.
Launch Day
The Launch Day is the calendar day within the Launch Slot on which the launch is
scheduled to take place.
Launch Window
The launch window is the period during which the launch can take place and is based on the

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injection parameters. A launch window of at least 45 minutes duration is required on every
calendar day of the launch period.
3.2.8.2 GTO Launch Window Requirements
The launch window is required to be a minimum of 45 minutes on each calendar day of the
launch period. If a 45 minutes launch window is not available during a period, the customer
will be required to review the potential options with CALT.
3.2.9 Separation Attitude
The attitude at separation should be stipulated in the Interface Control Document.
3.2.10 Separation Velocity
The separation system is designed to provide a relative separation velocity between the
spacecraft and launch vehicle upon separation. Typically, the minimum relative separation
velocity provided is 0.5 m/s.
3.2.11 Collision Avoidance Maneuver
The mission analysis will verify that the relative separation velocity is sufficient to prevent
re-contact between the third stage and the spacecraft. In addition, the third stage will
activate a collision avoidance maneuver to ensure that no re-contact can occur. Two
seconds after separation, the third stage helium thrusters will automatically initiate a reverse
thrust maneuver to slow down the launch vehicle and ensure an adequate separation
distance to the spacecraft.
3.3 LM-3A Launch Vehicle
3.3.1 Typical Orbit Parameters
3.3.1.1 Flight Sequence and Events
The typical flight sequence for the LM-3A launch vehicle in a standard GTO mission is
shown in Table 3-3 and Figure 3-6. The characteristic parameters of the typical trajectory
are shown in Table 3-4.

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Table 3-3 LM-3A Typical Flight Sequence
Events
Time (s)
Liftoff
0.0
Pitch Over
12.0
Stage-1 Shutdown
146.4
Stage-1/Stage-2 Separation
147.9
Fairing Jettisoning
236.9
Stage-2 Main Engine Shutdown
258.3
Stage-2 Vernier Engine Shutdown
263.3
Stage-2/Stage-3 Separation/Stage-3 First Start
264.3
Stage-3 First Shutdown
617.3
Coast Flight Phase Beginning
620.8
Coast Phase Ending/Stage-3 Second Start
1252.5
Stage-3 Second Shutdown/Velocity Adjustment Beginning
1374.4
Velocity Adjustment Ending
1394.4
SC/LV Separation
1474.4

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Figure 3-6 LM-3A Flight Profile

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Table 3-4 Characteristic Parameters of Typical Trajectory
Events
Flight
Altitude (km)
Ground
Range (km)
Projection
Latitude (deg)
Projection
Longitude (deg)
Liftoff
1.8
0.0
28.2
102.0
Stage-1 Shutdown
55.6
79.1
27.9
102.8
S-1/S-2 Separation
56.8
82.3
27.9
102.8
Fairing Jettisoning
119.0
324.9
27.3
105.2
S-2 Main Engine
Shutdown
134.1
403.3
27.1
106.0
S-2 Vernier Engine
Shutdown
137.8
423.0
27.1
106.2
S-2/S-3 Separation
138.6
427.0
27.1
106.2
Stage-3 First
Shutdown
195.3
2291.5
21.4
123.5
Coast Phase
Beginning
195.2
2316.6
21.3
123.8
Stage-3 Second
Start
194.9
6853.7
2.1
160.8
Stage-3 Second
Shutdown
212.9
7855.1
-2.4
168.5
Velocity Adjust.
Ending
222.7
8044.3
-3.3
170.0
SC/LV Separation
288.0
8793.0
-6.6
175.9
3.3.1.2 Main Parameters Curves of Typical Orbit
The flight acceleration, velocity, Mach numbers and altitude vs. time are shown in Figure
3-7a and Figure 3-7b.

Figure 3-7a LM-3A Load Factor & Velocity vs. Flight Time

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Figure 3-7b LM-3A Altitude & Mach vs. Flight Time
3.3.2 Launch Capability for LM-3A Launch Vehicle
3.3.2.1 Standard GTO Mission
LM-3A launch capability is 2,600 kg for standard GTO mission.
3.3.2.2 Non-Standard GTO
The LM-3A GTO launch capability varies with apogee altitudes for a range of inclinations, as
shown in Figure 3-8.

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Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
100,000
14
1,154
1,011
892
795
16
1,395
1,238
1,110
1,003
18
1,637
1,463
1,323
1,206
20
1,877
1,689
1,534
1,407
22
2,101
1,899
1,729
1,593
24
2,334
2,106
1,920
1,780
26
2,523
2,282
2,083
1,928
28.5
2,600
2,354
2,153
1,993
Figure 3-8 LM-3A GTO Capability
3.3.2.3 Low-Earth Orbit (LEO) Mission
The launch capability of the LM-3A for standard LEO mission into a 200 km circular orbit at
28.5 deg. inclination is 6,000 kg.
3.3.2.4 Sun-Synchronous Orbit (SSO) Mission
The LM-3A has the capability to inject a satellite directly into a Sun-Synchronous Orbit
(SSO). The launch performance of LM-3A for a Sun Synchronous Orbit mission is shown in
Figure 3-9.

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Figure 3-9 LM-3A SSO Capability
3.3.2.5 Deep Space Exploration Mission
The LM-3A has the capability to inject a satellite into an Earth-Escape orbit for deep space
exploration and interplanetary missions. The launch capability of LM-3A for Earth-Escape
mission is shown in Figure 3-10, where C3 is the square of the velocity at infinity in km2/s2.

Figure 3-10 LM-3A Earth-Escape Mission Capability
3.4 LM-3B Launch Vehicle
3.4.1 Typical Orbit Parameters
3.4.1.1 Flight Sequence and Events

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The typical flight sequence for the LM-3B launch vehicle in a standard GTO mission is
shown in Table 3-5 and Figure 3-11. The characteristic parameters of typical trajectory are
shown in Table 3-6.

Table 3-5 LM-3B Typical Flight Sequence
Events
Time (s)
Liftoff
0.0
Pitch Over
10.0
Boosters Shutdown
127.2
Boosters Separation
128.7
Stage-1 Shutdown
144.7
Stage-1/Stage-2 Separation
146.2
Fairing Jettisoning
215.2
Stage-2 Main Engine Shutdown
325.5
Stage-2 Vernier Engine Shutdown
330.5
Stage-2/Stage-3 Separation, and Stage-3 First Start
331.5
Stage-3 First Shutdown
615.7
Coast Phase Beginning
619.2
Coast Phase Ending, and Stage-3 Second Start
1258.4
Stage-3 Second Shutdown, Velocity Adjustment Beginning
1437.7
Velocity Adjustment Ending
1457.7
SC/LV Separation
1537.7

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Figure 3-11 LM-3B GTO Flight Profile

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Table 3-6 Characteristic Parameters of Typical Trajectory
Event
Flight
Altitude (km)
Ground
Range (km)
Projection
Latitude (deg)
Projection
Longitude (deg)
Liftoff
1.8
0.0
28.2
102.0
Booster Shutdown
53.9
68.7
28.2
102.7
Boosters Separation
55.4
71.8
28.2
102.8
Stage-1 Shutdown
71.0
108.2
28.1
103.1
S-1/S-2 Separation
72.5
112.0
28.1
103.2
Fairing Jettisoning
131.5
307.2
17.8
105.1
S-2 Main Engine
Shutdown
190.3
744.8
27.1
109.5
S-2 Vernier Engine
Shutdown
192.1
769.8
27.0
109.7
S-2/S-3 Separation
192.5
774.8
27.0
109.8
Stage-3 First
Shutdown
204.3
2466.2
22.8
125.9
Coast Phase
Beginning
204.3
2491.2
22.7
126.1
Stage-3 Second
Start
200.1
7061.3
4.4
164.1
S-3 Second
Shutdown
219.9
8531.1
-2.3
175.5
Velocity Adjust.
Ending
231.6
8720.0
-3.2
177.0
SC/LV Separation
304.6
9466.1
-6.5
182.8
3.4.1.2 LM-3B Main Parameters Curves of Typical Orbit
The flight acceleration, velocity, Mach numbers and altitude vs. time are shown in Figure
3-12a and Figure 3-12b.

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Figure 3-12a LM-3B Load Factor & Velocity vs. Flight Time

Figure 3-12b LM-3B Flight Mach & Altitude vs. Flight Time
3.4.2 Launch Capability for LM-3B Launch Vehicle
3.4.2.1 Standard GTO Mission
The GTO launch capability, which varies with the fairing encapsulation mode, is as follows:
Encapsulation-on-Pad: 5,100 kg (With 4000F fairing)
Encapsulation-in-BS3: 4,000 kg (For dual launch with 3700Z fairing only)

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The launch capability of LM-3B for the standard and recommended GTO mission as the
baseline in this manual is 5,100 kg.
3.4.2.2 Non-standard GTO Missions
The LM-3B GTO launch capability with the third stage operating command shutdown and
satellite encapsulation at the pad varies with apogee altitudes in a range of inclinations, as
shown in Figure 3-13.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
100,000
14
3,110
2,880
2,680
2,570
16
3,520
3,250
3,030
2,900
18
3,890
3,590
3,370
3,230
20
4,240
3,920
3,670
3,520
22
4,560
4,210
3,940
3,780
24
4,805
4,440
4,150
3,980
26
4,990
4,600
4,300
4,130
28.5
5,100
4,700
4,400
4,210
Figure 3-13 LM-3B GTO Performance (Encapsulation-on-Pad)

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The LM-3B GTO launch capability with the third stage operating command shut-down and
satellite Encapsulation-in-BS3 varies with apogee altitudes for a range of inclinations, as
shown in Figure 3-14.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
100,000
14
2,010
1,780
1,580
1,470
16
2,420
2,150
1,930
1,800
18
2,790
2,490
2,270
2,130
20
3,140
2,820
2,570
2,420
22
3,460
3,110
2,840
2,680
24
3,705
3,340
3,050
2,880
26
3,890
3,500
3,200
3,030
28.5
4,000
3,600
3,300
3,110
Figure 3-14 LM-3B GTO Performance (Encapsulation-in-BS3)
3.4.2.3 Low-Earth Orbit (LEO) Mission
The launch capability of the LM-3B for LEO mission into a 200 km circular orbit with an
inclination of 28.5 deg is 11,200 kg.

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3.4.2.4 Sun-Synchronous Orbit (SSO) Mission
The LM-3B launch vehicle has the capability of injecting a satellite directly into a
Sun-Synchronous Orbit. The launch performance of LM-3B for a sun synchronous mission
is shown in Figure 3-15.

Figure 3-15 LM-3B Sun synchronous Capability
3.4.2.5 Deep Space Exploration Mission
The LM-3B has the capability of injecting a satellite into an Earth Escape orbit for deep
space exploration and interplanetary missions. The launch capability of LM-3B for
Earth-Escape mission is shown in Figure 3-16, where C3 is the square of the velocity at
infinite in km2/s2.

Figure 3-16 LM-3B Earth Escape Mission Capability

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3.5 LM-3BE Launch Vehicle
3.5.1 Typical Orbit Parameters
3.5.1.1 Flight Sequence and Characteristic Parameters
The typical flight sequence of the LM-3BE launch vehicle for a standard GTO mission is
shown in Table 3-7 and Figure 3-17. The characteristic parameters of typical trajectory are
shown in Table 3-8.
Table 3-7 LM-3BE Typical Flight Sequence
Events
Time (s)
Liftoff
0.0
Pitch Over
11.0
Boosters Shutdown
140.1
Boosters Separation
141.6
S-1 Shutdown
157.5
S-1/S-2 Separation
159.0
Fairing Jettisoning
235.2
S-2 Main Engine Shutdown
339.4
S-2 Vernier Engine Shutdown
344.4
S-2/S-3 Separation/S-3 First Start
345.4
S-3 First Shutdown
628.6
Coast Phase Beginning
632.1
Coast Phase Ending/S-3 Second Start
1276.1
S-3 Second Shutdown/Velocity Adjustment Beginning
1470.5
Velocity Adjustment Ending
1490.5
SC/LV Separation
1570.5

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Figure 3-17 LM-3BE Flight Profile

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Table 3-8 LM-3BE Characteristic Parameters of Typical Trajectory
Events
Flight Altitude
(km)
Ground Range
(km)
Projection
Latitude (deg)
Projection
Longitude (deg)
Liftoff
1.8
0.0
28.2
102.0
Booster Shutdown
55.7
79.2
28.0
102.8
Boosters Separation
57.0
82.5
28.0
102.9
Stage-1 Shutdown
70.7
121.2
28.0
103.2
S-1/S-2 Separation
72.0
125.3
28.0
103.3
Fairing Jettisoning
125.2
356.9
27.6
105.6
S-2 Main Engine
Shutdown
167.5
794.8
26.8
110.0
S-2 Vernier Engine
Shutdown
169.2
820.7
26.8
110.2
Stage-2/Stage-3
Separation
169.5
825.9
26.8
110.3
Stage-3 First
Shutdown
178.8
2552.6
22.4
126.7
Coast Phase
Beginning
178.7
2578.0
22.3
126.9
Stage-3 Second
Start
181.6
7248.1
3.4
165.6
Stage-3 Second
Shutdown
231.4
8852.2
-3.4
178.3
Velocity Adjustment
Ending
246.0
9041.3
-4.2
179.9
SC/LV Separation
331.2
9786.8
-7.1
-174.1
3.5.1.2 LM-3BE Main Parameters Curves of Typical Orbit
The flight acceleration, velocity, Mach numbers and altitude vs. time are shown in Figure
3-18a and Figure 3-18b.

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Figure 3-18a LM-3BE Load Factor & Velocity vs. Flight Time

Figure 3-18b LM-3BE Altitude & Mach vs. Flight Time
3.5.2 Launch Capability for GTO Missions
3.5.2.1 Standard GTO
The GTO launch capability of LM-3BE, which varies with fairing encapsulation mode, is as
follows:
Encapsulation-on-Pad: 5,500 kg (Encapsulation-on Pad with 4000F fairing)
Encapsulation-in-BS3: 5,300 kg (Encapsulation-in-BS3 with 4200Z fairing)
The LM-3BE launch capability for the standard and recommended GTO mission as the
baseline in this manual is 5,500 kg.

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3.5.2.2 Non-standard GTO Missions
The LM-3BE GTO launch capability with third stage operating command shut-down and
satellite encapsulation on the pad varies with apogee altitudes for a range of inclinations, as
shown in Figure 3-19.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
90,000
100,000
14
3,470
3,220
3,010
2,880
2,840
16
3,870
3,590
3,360
3,220
3,180
18
4,260
3,950
3,700
3,550
3,500
20
4,620
4,280
4,020
3,850
3,800
22
4,940
4,580
4,290
4,110
4,060
24
5,200
4,820
4,510
4,330
4,270
26
5,390
4,990
4,670
4,480
4,420
28.5
5,500
5,080
4,760
4,560
4,500
Figure 3-19 LM-3BE GTO Performance (Encapsulation-on-Pad)

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The LM-3BE GTO launch capability with third stage operating command shut-down and
satellite encapsulation in BS3 varies with apogee altitudes for a range of inclinations, as
shown in Figure 3-20.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
90,000
100,000
14
3,270
3,020
2,810
2,680
2,640
16
3,670
3,390
3,160
3,020
2,980
18
4,060
3,750
3,500
3,350
3,300
20
4,420
4,080
3,820
3,650
3,600
22
4,740
4,380
4,090
3,910
3,860
24
5,000
4,620
4,310
4,130
4,070
26
5,190
4,790
4,470
4,280
4,220
28.5
5,300
4,880
4,560
4,360
4,300
Figure 3-20 LM-3BE GTO Performance (Encapsulation-in-BS3)
3.5.2.3 Low-Earth Orbit (LEO) Mission
The launch capability of LM-3BE for LEO mission into a 200 km circular orbit with an
inclination of 28.5 deg is 11,500 kg.
3.5.2.4 Sun-Synchronous Orbit (SSO) Mission

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The LM-3BE launch vehicle has the capability of injecting a satellite directly into
Sun-Synchronous Orbit. The launch performance of LM-3BE for sun synchronous missions
is shown in Figure 3-21.

Figure 3-21 LM-3B Enhanced SSO Capability
3.5.2.5 Deep Space Exploration Mission
The LM-3BE launch vehicle has the capability of injecting a satellite into Earth Escape orbit
for deep space exploration and interplanetary missions. The launch capability of LM-3BE
launch vehicle for Earth-Escape mission is shown in Figure 3-22, where C3 is the square of
the velocity at infinite in km2/s2.

Figure 3-22 LM-3BE Earth Escape Mission Capability

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3.6 LM-3C Launch Vehicle
3.6.1 Typical Orbit Parameters
3.6.1.1 Flight Sequence and Characteristic Parameters
The typical flight sequence of LM-3C for standard GTO mission is shown in Table 3-9 and
Figure 3-23. The characteristic parameters of typical trajectory are shown in Table 3-10.

Table 3-9 LM-3C Typical Flight Sequence
Events
Time (s)
Liftoff
0.0
Pitch Over
10.0
Boosters Shutdown
127.5
Boosters Separation
129.0
S-1 Shutdown
145.2
S-1/S-2 Separation
146.7
Fairing Jettisoning
258.7
S-2 Main Engine Shutdown
328.0
S-2 Vernier Engine Shutdown
333.0
S-2/S-3 Separation/S-3 First Start
334.0
S-3 First Shutdown
650.6
Coast Phase Beginning
654.1
Coast Phase Ending/S-3 Second Start
1323.2
S-3 Second Shutdown/Velocity Adjustment Beginning
1474.9
Velocity Adjustment Ending
1494.9
SC/LV Separation
1574.9

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Figure 3-23 LM-3C Flight Profile

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Table 3-10 LM-3C Characteristic Parameters of Typical Trajectory
Events
Flight
Altitude (km)
Ground Range
(km)
Projection
Latitude (deg)
Projection
Longitude (deg)
Liftoff
1.8
0.0
28.2
102.0
Booster Shutdown
48.7
50.6
28.2
102.5
Boosters Separation
50.0
52.9
28.2
102.6
Stage-1 Shutdown
64.7
82.0
28.1
102.9
S-1/S-2 Separation
66.1
85.1
28.1
102.9
Fairing Jettisoning
147.5
374.7
27.7
105.8
S-2 Main Engine
Shutdown
181.9
640.6
27.3
108.4
S-2 Vernier Engine
Shutdown
184.3
663.4
27.2
108.7
Stage-2/Stage-3
Separation
184.8
667.9
27.2
108.7
Stage-3 First
Shutdown
208.7
2465.0
22.8
125.8
Coast Phase
Beginning
208.6
2490.0
22.7
126.1
Stage-3 Second
Start
194.8
7295.2
3.2
165.9
Stage-3 Second
Shutdown
215.8
8541.6
-2.5
175.6
Velocity Adjustment
Ending
226.4
8730.8
-3.3
177.0
SC/LV Separation
295.1
9478.8
-6.6
182.9

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3.6.1.2 LM-3C Main Parameters Curves of Typical Orbit
The flight acceleration, velocity, Mach numbers and altitude vs. time are shown in Figure
3-24a and 3-24b.

Figure 3-24a LM-3C Load Factor & Velocity vs. Flight Time

Figure 3-24b LM-3C Altitude & Mach vs. Flight Time

3.6.2 Launch Capability for GTO Missions
3.6.2.1 Standard GTO

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The LM-3C launch capability for GTO, which varies with the fairing encapsulation mode, is
as follows:
Encapsulation-on-Pad: 3,800 kg (With 4000F fairing)
Encapsulation-in-BS3: 3,600 kg (Single launch with 3700Z fairing)
The LM-3C launch capability for the standard and recommended GTO mission as the
baseline in this manual is 3,800 kg.
3.6.2.2 Non-standard GTO Missions
The LM-3C launch capability for GTO mission with the third stage operating command
shut-down and the satellite encapsulation on the pad varies with apogee altitudes for a
range of inclinations, as shown in Figure 3-25.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
100,000
14
2,215
2,004
1,835
1,702
16
2,532
2,304
2,121
1,971
18
2,834
2,585
2,385
2,228
20
3,150
2,876
2,658
2,480
22
3,420
3,120
2,882
2,697
24
3,600
3,287
3,039
2,845
26
3,721
3,402
3,141
2,944
28.5
3,800
3,471
3,210
3,001
Figure 3-25 LM-3C GTO Performance (Encapsulation-on-Pad)

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The LM-3C launch capability for GTO mission with the third stage operating command
shut-down and the satellite encapsulation in BS3 varies with apogee altitudes for a range of
inclinations, as shown in Figure 3-26.

Apogee Altitude (km)
Inclination
(deg)
35,958
50,000
70,000
100,000
14
2,015
1,804
1,635
1,502
16
2,332
2,104
1,921
1,771
18
2,634
2,385
2,185
2,028
20
2,950
2,676
2,458
2,280
22
3,220
2,920
2,682
2,497
24
3,400
3,087
2,839
2,645
26
3,521
3,202
2,941
2,744
28.5
3,600
3,271
3,010
2,801
Figure 3-26 LM-3C GTO Performance (Encapsulation-in-BS3)
3.6.2.3 Low-Earth Orbit (LEO) Mission
The launch capability of the LM-3C for LEO mission into a 200 km circular orbit with an
inclination of 28.5 deg is 9,100 kg.
3.6.2.4 Sun-Synchronous Orbit (SSO) Mission
The LM-3C launch vehicle has the capability of injecting a satellite directly into

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Sun-Synchronous Orbit. The launch performance of LM-3C for sun synchronous missions is
shown in Figure 3-27.

Figure 3-27 LM-3C SSO Capability
3.6.2.5 Deep Space Exploration Mission
The LM-3C launch vehicle has the capability of injecting a payload into Earth Escape orbit
for deep space exploration and interplanetary missions. The launch capability of LM-3C for
Earth-Escape mission is shown in Figure 3-28, where C3 is the square of the velocity at
infinite in km2/s2.

Figure 3-28 LM-3C Earth Escape Mission Capability

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CHAPTER 4 PAYLOAD FAIRING
4.1 Introduction
The fairing is designed to limit the aerodynamic thermal load transmitted to the satellite and
attenuate the acoustic loads from the main engines during launch.

Figure 4-1 Payload Faring Encapsulation on Launch Pad

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The aerodynamic thermal flux upon fairing jettison is not greater than 1,135 W/m2. This is
the standard mission profile and all satellites are required to meet the thermal flux
requirements.
There are five fairings available for the LM-3A Series launch vehicles, which are designated
as 3350, 3700Z, 4000F, 4200F and 4200Z.
The two encapsulation methods are as follows:
Encapsulation-on-Pad: When the satellite is encapsulated on the pad the satellite and
fairing are transported to launch pad separately, then the fairing is encapsulated following
the mate of satellite to the launch vehicle.
Encapsulation-in-BS3: When the satellite is encapsulated in BS3 the satellite is mated to
the Payload Adapter (PLA) and encapsulated in the fairing in BS3. The encapsulated
satellite is shipped to the launch pad in the fairing and the complete assembly is mated to
the launch vehicle.
The application of these fairings is shown in Table 4-1 below.
Table 4-1 Fairing Type
Fairing Type
Diameter (mm)
LV Version
Encapsulation
3350
3,350
LM-3A
On Pad only
3700Z
3,700
LM-3B/3C
In BS3 only
4000F
4,000
LM-3B/3BE/3C
On Pad only
4200F
4,200
LM-3BE
On Pad only
4200Z
4,200
LM-3BE
In BS3 only
4.2 3350 Fairing
The 3350 fairing has an external diameter of 3,350 mm, a total height of 8,887 mm. This
fairing is only available for the LM-3A launch vehicle and designed to be encapsulated on
the pad. The fairing can be used with the 937B, 1194 and 1194A Payload Adapter (PLA)
interfaces and its configuration and static envelope are shown in Figure 4-2a and Figure
4-2b.

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Figure 4-2a LM-3A Fairing

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Figure 4-2b 3350 Fairing Static Envelope

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Figure 4-3 3350 Fairing Encapsulation-on-Pad
4.3 3700Z Fairing
The 3700Z fairing has an external diameter of 3,700 mm, a total height of 10,796 mm, and is
used for dual launch with the LM-3B launch vehicle with satellite encapsulated in the BS3.
This fairing can be used with the 1194 and 1194A Payload Adapter (PLA) interfaces. The
configuration and static envelope are shown in Figure 4-4a and 4-4b.

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Figure 4-4a 3700Z Fairing

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Figure 4-4b 3700Z Fairing Static Envelope

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4.4 4000F Fairing
The 4000F fairing has an external diameter of 4,000 mm a total height of 9,561 mm, and is
used with the LM-3B and LM-3C launch vehicle for satellite encapsulation on the launch pad.
This fairing can be used with the 937B, 1194 and 1194A payload adapters. The configuration
and static envelope are shown in Figure 4-5a and Figure 4-5b.

Figure 4-5a 4000F Fairing

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Figure 4-5b 4000F Fairing Static Envelope

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Figure 4-6 4000F Faring Encapsulation-on-Pad

Figure 4-7 4000F Fairing Encapsulation-on-Pad

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4.5 4200F Fairing
The 4200F fairing has an external diameter of 4,200 mm, a total height of 9,777 mm, and is
used with the LM-3BE launch vehicles for satellite encapsulated on the launch pad. This
fairing can be used with the 973B, 1194 and 1194A Payload Adapter (PLA) interfaces. The
configuration and static envelope are shown in Figure 4-8a and 4-8b.

Figure 4-8a 4200F/4200Z Fairing

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Figure 4-8b 4200F Fairing Static Envelope

4.6 4200Z Fairing
The 4200Z fairing has an external diameter of 4,200 mm, a total height of 9,381 mm, and is
used with the LM-3BE launch vehicles for satellite encapsulated in BS3. This fairing can be
used with the 973B, 1194 and 1194A Payload Adapter (PLA) interfaces. The configuration
and static envelope are shown in Figure 4-8a and Figure 4-8c.

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Figure 4-8c 4200Z Fairing Static Envelope

4.7 Static Envelope
The static envelope of the fairing that is usable with the customers’ satellite defines the
maximum physical dimensions of the satellite to ensure clearance within the fairing. The
static envelope varies with the fairing and adapter selected, but within the cylindrical section
of the fairing the static envelope for each group of fairings is 3,000 mm diameter for the 3350

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fairing, 3,350 mm diameter for the 3700Z fairings, 3,650 mm diameter for the 4000F fairings,
and 3,850 mm diameter for the 4200F and 4200Z fairings.
The fairing and adapter static envelopes are shown in Figure 4.2b to Figure 4-8c.
Sections of the satellite appendages that extend into the exclusion zones of the static
envelope will require technical coordination with CALT.
4.8 Dynamic Envelope
The dynamic envelope is determined during the mission analysis based on the SC/LV
Coupled Loads Analysis (CLA).
4.9 Thermal Protection
The fairings are designed to ensure that the temperature of the inner surface of the fairing
structure does not exceed 80°C during the powered flight phase.
4.10 Fairing Separation
The fairing separation system includes a lateral and longitudinal release mechanism with a
rotating separation mechanism. The release mechanism for the 3350 fairing uses explosive
bolts, and the separation mechanism consists of separation springs with rotating hinges.
The release mechanism for the 3700Z, 4000F, 4200F and 4200Z fairing use the notched
bolts, explosive cord and explosive bolts, and the separation mechanism consists of
separation springs with rotating hinges. See Figure 4-9a and Figure 4-9b.

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Figure 4-9a Structure & Separation Device of LM-3A Fairing

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Figure 4-9b Structure & Separation Device of LM-3B/3BE/3C Fairing

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4.10.1 Lateral Release Mechanism
For all types of fairing, the base ring of the fairing is connected to the forward skirt of the
third stage with explosive bolts. The lateral explosive bolts are simultaneously detonated to
release the lateral connection for fairing separation.
4.10.2 Longitudinal Release Mechanism
For the 3350 fairing, the longitudinal separation system comprises twelve explosive bolts,
which are simultaneously released with the lateral explosive bolts at fairing separation.
For 3700Z, 4000F, 4200F and 4200Z fairings, the longitudinal release mechanism uses
notched bolts, explosive cord, expanding hose, and two explosive bolts. All these parts
jointly take their actions to perform the release functions.
4.10.3 Fairing Separation Mechanism
The fairing separation mechanism comprises hinges and springs, with each half of the
fairing being supported by two hinges. There are six separation springs on each half of the
fairing, and after the lateral and longitudinal release mechanisms have been activated, each
half of the fairing turns around the hinge under the spring force to accomplish the fairing
jettisoning.
4.11 RF Windows and Access Doors
Radio frequency (RF) transparent windows can be incorporated into the fairing biconical
section and cylindrical section to allow the satellite command and telemetry signals to be
available to the satellite via RF transmission. The RF windows are tailored to the customer’s
specific requirements as defined in the ICD. The insertion loss and transparency rate for the
RF transparent windows are shown in Table 4-2.

Table 4-2 Insertion loss and transparency rate for the RF transparent windows
Frequency (GHz)
0.4
4
8
10
13
15
17
Insertion loss (dB)
0.25
0.47
0.52
1.63
1.4
2.73
4.11
Transparency Rate
0.94
0.89
0.88
0.68
0.72
0.53
0.38
Access doors can be provided in the cylindrical section to permit limited access to the
satellite after the fairing encapsulation and these are tailored to the customer’s requirements.

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The un-shadowed areas in Figure 4-10a, Figure 4-10b and Figure 4-10c are available for RF
windows and access doors. The size, location and number of RF window / access door will
be fixed at eight months prior to launch (launch - 8 months).
4.12 Customer Logo
The customer can have a specially designed logo painted or sticked to the cylindrical
section of the fairing and appropriately positioned. The overall design and location will be
determined through the Interface Control Document (ICD) and Mission Reviews. The
artwork for the logo is to be supplied by the customer not later than eight months prior to
launch (Launch - 8 months).

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Figure 4-10a Access Door & RF Window Allowable of 3350 Fairing

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Figure 4-10b Access Door & RF Window Allowable of 4000F/4200F/4200Z Fairing

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Figure 4-10c Access Door & RF Window Allowable of 3700Z Fairing

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CHAPTER 5 MECHANICAL AND ELECTRICAL
INTERFACES
5.1 Introduction
This Chapter defines the mechanical and electrical interfaces between the satellite and the
LM-3A Series launch vehicles. The mechanical interface is to mate the satellite with the
launch vehicle via the Payload Adapter (PLA). The electrical interface electrically connects
the satellite with launch vehicle and satellite ground support equipment.
5.2 Satellite Reference Axes
The satellite coordinate system shall be aligned with the coordinate system of the LM-3A
Series launch vehicle as shown in Figure 5-1.

Figure 5-1 LM-3A Series Launch Vehicle Coordinate System

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5.3 Mechanical Interface
5.3.1 Introduction
The satellite is mounted on the launch vehicle through the Payload Adapter (PLA) which is
mated with the VEB by bolts. The top ring of the PLA is the primary mechanical interface
with the satellite, which is attached to the PLA via a clampband. The PLA also includes the
separation springs, separation sensors and the umbilical connectors.
5.3.2 Payload Adapter
The LM-3A Series launch vehicle uses the international standard adapter types 937B, 1194
and 1194A to interface with the satellite. These are flight proven adapters and all include the
separation system and umbilical brackets. Other interfaces can also be accommodated
through technical coordination with CALT.
5.3.2.1 937B Interface (Encapsulation-on-Pad)
The 937B mechanical interface is the upper part of 937B interface adapter, which is a 900
mm high truncated cone. (Refer to Figure 5-2). The top ring, which interfaces with the
satellite, is made of high-strength aluminum alloy.
5.3.2.2 1194 Interface (Encapsulation-on-Pad)
The 1194 interface adapter is a 650 mm high truncated cone with top ring diameter of 1,215
mm (Refer to Figure 5-3). The top ring, which interfaces with the satellite, is made of
high-strength aluminum alloy.

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Figure 5-2 937B Interface (Encapsulation-on-Pad)

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Figure 5-3 1194 Interface (Encapsulation-on-Pad)

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5.3.2.3 1194A Interface (Encapsulation-on-Pad)
The 1194A adapter for Encapsulation-on-Pad is a 650 mm high truncated cone (see Figure
5-4 and Figure 5-5).

Figure 5-4 Payload Adapter with 1194A Interface

5.3.2.4 1194A interface (Encapsulation-in-BS3)
The 1194A adapter for encapsulation in BS3 is a 450 mm high truncated cone (see Figure
5-6).

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Figure 5-5 1194A Interface (Encapsulation-on-Pad)

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Figure 5-6 1194A Interface (Encapsulation-in-BS3)

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5.3.3 Separation System
5.3.3.1 Satellite to Adapter Separation System
The separation system uses a clampband to firmly install the satellite to the Payload Adapter
(PLA) and reliably release the satellite from the launch vehicle upon separation. The
separation mechanism mounted on the adapter provides the relative separation velocity
between the satellite and launch vehicle via separation springs. The separation system is
shown in Figure 5-7a, Figure 5-7-b, Figure 5-7c, Figure 5-7d and Figure 5-7e.
5.3.3.2 Separation Mechanism
The separation mechanism consists of a clampband, the non-contamination explosive bolts
or cutters, V-shoes, and retention springs.
The clampband consists of two halves made of high-strength steel. There are two models of
clampband available, the C100 model which is 100 mm wide and uses two explosive bolts
(see Figure5-7a), and C60 model which is 60 mm wide and uses two explosive cutters (see
Figures 5-7b and Figure 5-7c).

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Figure 5-7a C100 Separation Mechanism

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Figure 5-7b C60 Separation Mechanism

Figure 5-7c C60 Separation Mechanism

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Figure 5-7d SC Separation Springs

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Figure 5-7e 1194/1194A Clampband
The non-contamination explosive bolts and explosive cutters are fully redundant with each
bolt or cutter having two igniters. The maximum allowable tension force for cutters is 45 kN,
and for the explosive bolt the margin is much greater.
The V-shoes, which are made of high strength aluminum, are used for clamping the
interface ring of the satellite to the top ring of the Payload Adapter (PLA).
The lateral-restraining springs are used for restraining the outward movement of the
clampband and the longitudinal-restraining springs restrict the movement of the clampband
toward the satellite after clampband release. The two halves of the clampband are therefore
restrained to the PLA below the separation plane to avoid re-contact.
The separation spring assembly is mounted on the PLA providing a separation velocity of
0.5 m/s or greater, and can also provide a lateral axis spin rate of less than 3.0 deg/s, the
spin rate being dependent on the customer’s requirement.
5.3.3.3 Installation Requirements
During the installation of the clampband system, 10 strain gauges are installed on each half

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of the clampband so that the strain and pretension at each measuring point can be
monitored in real time. A special purpose tool is used to apply the pretension, which is
normally in the range of 24~30 kN. The pretension can, however, be adjusted to meet the
specific requirements of the satellite and the results of the coupled load analysis.
For both C100 and C60 clampband, a minimum clearance of ≥ 85 mm for C100 and ≥ 60
mm for C60 is required between the lowest point on the satellite and the separation plane
after installation. If that cannot be achieved, there should be a minimum clearance of ≥ 20
mm between the lateral-restraining springs and the lowest point on the satellite.
5.4 Electrical Interfaces
5.4.1 Summary
CLTC and CALT provide the RF link and the umbilical link between the satellite and the
satellite Electrical Ground Support Equipment (EGSE) after the satellite is mated with
launch vehicle. These links allow the customer to conduct pre-launch testing and monitor
the satellite during the launch count down.
The schematic of the umbilical cable and RF links between the satellite and the EGSE is
shown in Figure 5-8 and Figure 5-9.
The umbilical cable is designed and manufactured by CALT to meet the requirements of the
customer based on their satellite platform. The 350 m cable from FER to the Umbilical Tower,
and the hardware labeled EB26/EB36, BOX3 and Power-supply 1&2 are common to all
missions. The umbilical cable integral to the launch vehicle plus the ground facility cable
from WXTC to ED10, ED12, ED13, ED14, BOX1 and BOX2, will be designed specifically for
customer.

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Figure 5-8 SC Umbilical Cable Configuration

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Figure 5-9 SC RF Links & Umbilical Cables
5.4.2 In-Flight-Disconnectors
Two sets of In-Flight-Disconnectors (IFDs) are required and are symmetrically mounted
outside the top ring of the Payload Adapter (PLA). The exact location and clocking will be
coordinated between customer and CALT and defined in the ICD.
It is normal practice for customer to select the IFDs. In such cases, it is advisable that the
customer supply the launch vehicle portion of IFDs including installation tools to CALT. IFDs
manufactured by DEUTSCH Engineered Connecting Devices are recommended (see Table
5-1), while the Chinese made YF8-64 are also available if selected.
Table 5-1 IFDs
LV Side
SC Side
Code
Type
Code
Type
P01
D8179E37-OPN
J01
D8174E37-OSN
P02
D8179E37-OPY
J02
D8174E37-OSY
Notes: The IFDs will be totally disconnected when the two halves have separated from each
other by 13.5 mm. The customer can also select other products by Deutsch as required, e.g.
the DBAS7061, which is manufactured by cold-press processing.
5.4.3 In-Flight-Disconnectors Characteristics
The customer shall specify all the characteristics of the IFDs, including the pin assignments,

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usage, maximum voltage, maximum current, one-way maximum resistance, type and
pin-assignment of connectors to be mated with the ground equipment, etc. CALT will design
the umbilical cable according to the above requirements, which will also be defined in the
ICD.
5.4.4 Umbilical Cable Onboard Launch Vehicle
(a) Configuration
The umbilical cables onboard the launch vehicle are cables from the IFDs (P01, P02) to
WXTC, which will fly with launch vehicle (See Table 5-2).
Table 5-2 Umbilical Cable
Code
Description
P01, P02
IFD
EC1, EC2
Mechanical technological interfaces between
PLA and LV
EY1
Interface between umbilical cable and LV TM
system, through which the SC/LV IFD
(J01/P01, J02/P02) separation signal is sent
to LV TM system
WXTC
Umbilical cable connector (type: JF5-231)
G1
Grounding points to connect the shielding of
the wires to the shell of the LV
(b) Separation Signal
The separation signal is generated through break-wires on the IFDs P01 & P02. The
satellite will receive the separation signals once the break-wires are disconnected upon
separation.
In addition, there are another two break-wires on the IFDs J01 & J02 that will generate the
separation signal for the launch vehicle once the break-wires are disconnected upon
separation. This separation signal will be sent to the launch vehicle telemetry system
through the EY1 interface. The break-wires allow a maximum current of 100 mA and a
maximum voltage of 30 V.
Figure 5-10 shows an example the break-wire circuitry.

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Figure 5-10 An Example of Separation Signals & Break-wires
There are also two micro-switches on the PLA to give the mechanical separation signal,
which is also sent to the launch vehicle telemetry system.
5.4.5 Ground Umbilical Cables
(a) Configuration
The ground umbilical cables are cables from the WXTC to COTE in the FER which are
shown in Figure 5-9, and the functions are shown in Table 5-3.

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Table 5-3 Ground Umbilical Cable Functions
Code
Description
WXTC
WXTC is the umbilical cable connector (LV-Ground) whose female half (socket) is
installed at the VEB, while the male half (pin) is attached to the top end of ground
cable. The disconnection of WXTC is electrically controlled. (The disconnection is
powered by BOX3. In the mean time, forced disconnection is also used as a
back-up separation method.) Generally, WXTC disconnects at about 10 minutes
prior to launch. If the launch was terminated after the disconnection, WXTC could
be reconnected within 30 minutes.
The SC should switch over to internal power supply and cut off ground power
supply at 5 minutes prior to WXTC disconnection. Therefore, during disconnection
only a low current monitoring signal (such as 30 V, ≤100 mA) is permitted to pass
through the WXTC.
BOX1
BOX1 is a box adapter for the umbilical cable that is located inside the SC Cable
Measurement Room on floor 8.5 of the umbilical tower. (If needed, BOX1 can
provide more interfaces for the connection with SC ground equipment.)
BOX2
BOX2 is another box adapter for the umbilical cable that is located inside the SC
FER on ground. The SC ground support equipment are also located inside the
FER.
(b) Ground Interface
The customer determines the requirement for the ground interface, but normally there are
two ground interfaces for the satellite test console (J11 & J12). The detailed requirements for
these ground interfaces shall be jointly agreed by customer and CALT (See Table 5-4).
CALT typically prepares short interface cables, so the customer is required to provide the
mating half of the connectors (J11 & J12) for the cables that interface with the satellite
ground equipment.
Table 5-4 Ground Interfaces for Satellite Test
Location
Code
Specification
Quantity
LV side interfaces
J11
J12
To be defined by SC side
2
2
If the customer cannot provide the connectors, there are two options available. CALT can
provide the interface cables with each wire labeled with a number in accordance with the pin
assignment of the connector for installation into the customer’s connectors when available.
The second option is for the customer to manufacture this cable and provide it to CALT, in
which case CALT will provide the mating set of connectors for BOX2 in Figure 5-8. The
length of this cable is approximately 5 m.

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5.4.6 Shielding and Grounding
5.4.6.1 Shielding
In order to protect against lightning and stray voltages, the following measures have been
implemented:
(a) The ground umbilical cable has two shielding layers. The outer shielding is for protection
against lightning and the inner shielding is for EMI protection.
(b) The ground umbilical cables from WXTC to BOX2 have a grounding point every 20
meters for the outer shielding layer to ensure that any charge buildup from a lightning strike
or other inductance can be discharged immediately.
(c) After the umbilical connectors are mated, the cable shielding is automatically connected
to the launch vehicle shell structure.
(d) The inner shield has a single ground. The inner shields of the umbilical cables onboard
launch vehicle are connected to BOX2.
(e) The inner and outer shields are insulated from each other within the cables.
5.4.6.2 Continuity of Satellite Grounding
The satellite shall have a ground reference point close to the separation plane. The
resistance between all other metal parts of satellite (shell, structures, etc.) and the reference
ground shall be less than 10 m for a current of 10 mA. In order to maintain the continuity of
the grounding, the bottom of satellite structure to be mated with the Payload Adapter (PLA)
shall not be treated with any protective process that would affect its ability to meet the
resistance requirements between the satellite and PLA.
5.5 Satellite Constraints
The satellite shall not transmit any remote signal that could be dangerous or interfere with
the launch vehicle flight until the satellite has been completely separated from the launch
vehicle. The satellite shall not start its automated post separation program until it receives
the separation signal from the launch vehicle.
5.6 RF Links

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5.6.1 RF Relay Path
The launch site can provide an RF link between the Satellite Test Equipment (STE) and the
SC, whenever the SC is in BS or at the launch pad. The RF link equipment is set up
between BS2 and the launch towers. It provides uplink and downlink RF channels that can
connect the STE with the SC and transmit the test and control signals when the SC is mated
to the launch vehicle at the launch pad. The RF system uses optical fiber to transmit and
receive the signals.
5.6.2 Characteristics and Interface of RF Link
(a) Frequency
The optical fiber of the RF system allows the transmission of 350 MHz to 15 GHz signal,
while the microwave processing unit allows the transmission of signals in UHF, L, S, C, Ku
and X bands.
(b) Signal Level
The signal level interface between RF system and STE is shown in Table 5-5.
Table 5-5 Signal level interface between RF system and STE
SC Antenna
System Terminal
Band
EIRP (dBm)
PFD (dBW/m2)
Input (dBm)
Output (dBm)
Uplink
-20 ~ 10
/
/
-50 ~ -80
S
Downlink
/
-30 ~ -10
0 ~ 60
/
Uplink
-20 ~ 10
/
/
-50 ~ -80
C
Downlink
/
-30 ~ -10
0 ~ 60
/
UHF
Uplink
-20 ~ 10
/
/
-50 ~ -80
X
Downlink
/
-30 ~ -10
0 ~ 60
/
L
Downlink
/
-30 ~ -10
0 ~ 60
/
Uplink
/
-20 ~ 10
/
-50 ~ -80
Ku
Downlink
-30 ~ -10
/
30
/
5.7 Post Encapsulation Interfaces
5.7.1 Payload Usable Volume
The payload usable volume is the area within the fairing that is available to accommodate

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the satellite when mounted on the Payload Adapter (PLA) of the launch vehicle. This
represents the total maximum envelope available for the satellite and includes all
manufacturing tolerances, thermal protection, satellite appendages, etc, and also includes
the dynamic tolerances based on results of the standard coupled loads analysis.
If the satellite has appendages that protrude to outside the usable envelope, they shall be
reviewed during the ICD and Mission Reviews. If the protusions are found
after these satellite appendages have been completed, a special review shall be held to
resolve the issue.
5.7.2 Satellite Access
The satellite can be accessed while in the fairing via special access doors on the fairing as
agreed in the ICD. The satellite can be accessed up until (Launch – 2 hours) after which
time the fairing is sealed.
5.8 Interface Verification
5.8.1 Satellite AIT
During the satellite AIT program and prior to final acceptance for delivery to the launch site,
a full mechanical and electrical fit check can be performed. This interface verification, or
SC/LV fitcheck, is mandatory for all satellite platforms that have not previously been
launched with the LM-3A Series launch vehicles.
The primary objective of the test is to verify that the mechanical interface is correct, the
electrical interfaces are compatible and all the connectors and arming plugs can be
accessed. This fitcheck can then be followed by a SC/LV separation shock test to verify if
the satellite could be separated from the Payload Adapter (PLA) properly within the required
shock level.
The SC/LV fitcheck and separation shock test is typically performed at the satellite
manufacturers’ facility using the flight adapter and separation system provided by CALT.
5.8.2 Launch Site Interface Verification
The electrical and RF interfaces between the satellite and each site for satellite processing,
including those where the satellite is mated to the launch vehicle, shall be validated prior to
the satellite moving to a new facility and after its arrival at a new facility. This verification is to
be performed whenever the umbilical connecters are disconnected and reconnected.

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5.8.3 Satellite EGSE Interface Verification
There are typically two sets of EGSE. One set is permanently located in BS2 that is used as
the primary system for controlling the satellite. The second set moves with the satellite and
functions as a remote test interface with the satellite. The interfaces of this EGSE shall be
verified prior to its arrival at the launch site and following each move at the launch site.
The interfaces with the launch site facility are defined in Chapter 7.

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CHAPTER 6 ENVIRONMENTAL CONDITIONS
6.1 Summary
This Chapter provides the natural environment at Xichang Satellite Launch Center (XSLC),
the thermal environment during satellite processing, the thermal and mechanical
environment (vibration, shock & noise) experienced during the launch vehicle flight and the
electromagnetic environment during ground processing and the flight.
6.2 Environment at XSLC
The environmental data for the XSLC has been monitored over a significant period of time.
Table 6-1 and Table 6-2 provide the average monthly temperature and average monthly
ground wind speed respectively.

Table 6-1 Average Monthly Temperature in XSLC
Month
Highest (°C)
Lowest (°C)
Average (°C)
January
7.9
4.5
5.9
February
10.4
5.0
8.0
March
14.5
9.7
11.7
April
17.5
13.1
15.0
May
20.2
15.6
17.7
June
21.1
17.7
19.1
July
21.3
19.3
20.0
August
21.3
18.5
19.8
September
19.3
16.2
17.2
October
16.4
13.2
14.1
November
12.3
8.4
10.0
December
8.9
4.6
6.5

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Table 6-2 Average Monthly Ground Wind Speed in XSLC
Month
Average Speed (m/s) Days (>13 m/s)
January
2.2
0.5
February
2.3
1.1
March
2.3
2.5
April
2.0
1.6
May
1.5
0.6
June
1.0
0.4
July
1.1
0.2
August
1.2
0.1
September
0.9
0.2
October
1.1
0.1
November
1.4
0.0
December
1.7
0.2
The relative humidity at launch site:
 Maximum: 90% in rainy season
 Minimum:
42% in dry season
6.3 Satellite Processing Environment
Following arrival at XSLC, the satellite is integrated, tested and fueled in SC Processing
Building BS2 and BS3, and then, it is transported to the launch pad for combined operations.
The processing environments for the satellite therefore fall into three phases: processing in
BS2 and BS3, transportation to launch pad and final preparation on launch tower. The
following environment is provided and maintained throughout the whole process:
 Temperature: 15°C ~ 25°C
 Relative humidity: 35% ~ 55%
 Cleanliness: 100,000 class
 Acoustic level: ≤ 90 dB
For the detailed information about satellite processing environment, please refer to the
XSLC User’s Manual (Issue 2009).
6.3.1 Environment in SC Processing Building BS2 and BS3
During the integration and checkout of satellite in BS2 and BS3, the environment inside the
building is maintained by air-conditioning system.

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6.3.2 Environment during Transportation to Launch Pad
6.3.2.1 For Fairing Encapsulation-on-Pad
The satellite is transferred to the launch pad in an environmentally controlled and sealed
container, and the transfer takes approximately 30 minutes. The satellite transfer container
is a sealed cylinder with a height of 8,070 mm and diameter of 4,120 mm with its wall made
of aluminum sandwich for thermal insulation (See Chapter 8).
Before transportation, the container is pressurized with dry nitrogen, followed by a thermal
stabilization period to allow the satellite and container to stay at a steady temperature
between 15°C and 25°C. The container is then sealed and moved out of BS3 for transport to
the pad.
When the container arrives at the launch pad, it is hoisted to the 8th floor of the Service
Tower, where an environmentally controlled clean room has been established. The
procedures for this operation are detailed in Chapter 8.
After arrival at the launch pad clean room, the container goes through another stabilization
period for the satellite thermal conditions to become adapted to those of the clean room.
When the satellite and container have reached a thermally stabilized status, the container is
opened and the satellite is moved out to mate to the launch vehicle.
6.3.2.2 For Fairing Encapsulation-in-BS3
The same environment conditions as Paragraph 6.3.2.1 are provided and maintained except
that the fairing replaces the SC container.
When the satellite encapsulated with fairing arrives at the launch pad, it is hoisted to the 8th
floor of the Service Tower. The fairing is mated to the launch vehicle. The procedures for this
operation are detailed in Chapter 8.
6.3.3 Air-conditioning inside Fairing
The fairing air-conditioning system is connected following the mating of the satellite
encapsulated in the fairing to the launch vehicle (Encapsulation-in-BS3) or final
encapsulation of the satellite at the launch pad (Encapsulation-on-Pad). The air-conditioning
system becomes operational immediately following its connection to the fairing and the
configuration is shown in Figure 6-1.
The air-conditioning system maintains the environmental conditions within the fairing as

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shown in Table 6-3.
Table 6-3 Air-conditioning parameters inside Fairing
Temperature
13°C ~ 23°C
Relative Humidity
35% ~ 55%
Cleanliness
100,000 class
Air Speed inside Fairing
≤ 2 m/s
Noise inside Fairing
≤ 90 dB
Air Flow Rate
1500 ~ 3000 m3/hour (adjustable)
The air-conditioning is shut off 45 minutes before launch and can be re-established within 40
minutes if the launch is aborted. The air-conditioning inlets are shown in Chapter 4.

Figure 6-1 Pre-launch Air-condition System of LM-3A Series LVs

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6.4 Electromagnetic Environment
The satellite has to withstand the electromagnetic field generated by both the launch site RF
facility and launch vehicle RF systems while mounted on the launch vehicle. These
transmissions have been minimized to a very low level in the vicinity of the launch complex.
This section details the frequencies and levels of the potentially interfering radiation.
6.4.1 Radio Equipment onboard LM-3A Series Launch Vehicles and Ground
Test Equipment
Characteristics of on-board radio equipment and ground test equipment are shown in Table
6-4.
Table 6-4 Characteristics of on-board radio equipment and ground test equipment

Equipment
Frequency
(MHz)
Power
(W)
Sensitivity
Polarization
Antenna
position
Telemetry
Transmitter 3
2200~2300
10
/
linear
VEB
GNSS
receiver
1500~1700
/
≤ -130 dBW
RHCP
VEB
Telemetry
Transmitter 2
2200~2300
5
/
linear
Responder 1
Rx 5860~5910
Tx 4210~4250
2
≤ -120 dBW
linear
Stage-2
Intertank
Responder 2
Rx and Tx
5580~5620
300 (peak)
0.8~1.0 µs
800 Hz
<300 mW
(average)
≤ -90 dBW
linear
Stage-3
Rear shell
Beacon
2730~2770
2
/
linear
Stage-3
Rear shell
LAUNCH VEHICLE
Telemetry
command
Receiver
550~750
/
≤ -128 dBW
/
Stage-2
Intertank
Tester for
Responder 1
5870~5910
0.5
/
/
Tester for
Responder 2
5570~5620
100 W
(peak)
/
/
GROUND
Telemetry
Command
Transmitter
550~750
1 W
/
/
Tracking &
safety
system
ground test
room at
launch pad
The locations of the onboard radio equipment are shown in Figure 6-2.

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Figure 6-2 Locations of On-board Radio Equipment
6.4.2 RF Equipment and Radiation Strength at XSLC
 Working frequency:
5,577 ~ 5,617 MHz
 Antenna diameter: 4.2 m
 Impulse power: <1,500 kW
 Impulse width: 0.0008 ms
 Minimum pulse duration: 1.25 ms
 Average power: <1.2 kW
6.4.3 Launch Vehicle Electromagnetic Radiation and Susceptibility
The electromagnetic radiation levels generated by the launch vehicle and the launch site
equipment will not exceed those shown in Figure 6.3. The launch vehicle electromagnetic
susceptibility is shown in Figure 6.4. The energy levels of launch vehicle electromagnetic
radiation and susceptibility are measured at 1 m above VEB. The electromagnetic radiation

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of satellite shall not exceed the values shown in Figure 6-4.

Frequency (MHz)
Field Strength (dBV/m)
0.01 - 0.05
80
0.05 - 3
90
3 - 300
70
300 - 550
80
550 - 750
103
750 - 2200
80
2200 - 2300
138
2300 - 2730
80
2730 - 2770
107
2770 - 4200
80
4200 - 4250
107
4250 - 5580
80
5580 - 5620
99
5620 - 6000
80
6000 - 6500
80
6500 - 13500
80
13500 - 15000
80
15000 -
80
Figure 6-3 Electromagnetic Radiation from LV and XSLC

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Frequency (MHz)
Field Strength (dBV/m)
0.01 - 550
134
550 - 750
15
750 - 1000
134
1000 - 1500
140
1500 - 1700
10
1700 - 5580
140
5580 - 5910
35
5910 - 100000
140
Figure 6-4 Launch Vehicle Electromagnetic Radiation Susceptibility
6.4.4 EMC Analysis
CALT and the customer shall jointly conduct an EMC analysis for the satellite, launch vehicle
and launch site to verify their compatibility. The customer shall use the data provided in this
User’s Manual and shall provide the following information, as required in the ICD, to CALT
for analysis:

Satellite RF system configuration, characteristics, operating and radiating time
periods, antenna position and its transmission direction, etc.

Values and curves of the electric field of intentional and parasitic radiation generated
by the satellite RF system at satellite separation plane

The values and curves of the electromagnetic susceptibility of the satellite.

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6.5 Mechanical Environments
6.5.1 Steady State Acceleration
The launch vehicle flight generates external forces on the satellite due to the engine thrust
and aerodynamic forces. The typical maximum longitudinal steady state acceleration during
the launch vehicle powered flight is shown in Table 6-5. It can be seen that the maximum
longitudinal static acceleration occurs during the first stage flight. The maximum static
acceleration will vary slightly with the satellite mass.
Table 6-5 Typical Maximum Static Longitudinal Acceleration

LM-3A
LM-3B
LM-3C
During First Stage Flight
+5.0 g
N/A
N/A
During First Stage Booster Flight
N/A
+5.3 g
+5.3 g
During First Stage Core Flight
N/A
+3.6 g
+3.6 g
During Second Stage Flight
+2.9 g
+2.8 g
+2.8 g
During Third Stage first powered flight
+1.6 g
+1.2 g
+1.2 g
During Third Stage second powered flight
+2.7 g
+2.5 g
+2.5 g
Note: “+” means the direction of the acceleration coincides with the launch vehicle +X axis.
6.5.2 Vibration Environments
6.5.2.1 Sinusoidal Vibration
Low level sinusoidal vibration is present during the flight, but the significant vibration events
occur during engine ignition and engine shut-off, transonic flight and the stage separations.
The sinusoidal vibration (zero-peak value) at the satellite separation plane is shown in Table
6-6 below.
Table 6-6 Sinusoidal Vibration Levels
Direction
Frequency Range
Amplitude & Acceleration
5 - 8 Hz
3.11 mm
Longitudinal
8 - 100 Hz
0.8 g
5 - 8 Hz
2.33 mm
Lateral
8 - 100 Hz
0.6 g

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6.5.2.2 Random Vibration
The maximum random vibration levels are primarily generated by the noise during the
launch vehicle lift-off and transonic flight periods. The acoustic spectrum defined in
Paragraph 6.5.3 covers excitations produced by random vibration.
6.5.3 Acoustic Noise
The acoustic noise spectrum includes the engine noise and aerodynamic noise during flight,
with the maximum acoustic noise being experienced by the satellite during the lift-off and the
transonic flight phase. The values in Table 6-7 below are the maximum noise levels inside
the fairing from the launch and transonic flight phase.
Table 6-7 Acoustic Noise in the Fairing
Central Frequency of Octave Bandwidth (Hz)
Acoustic Pressure Level (dB)
31.5
124
63
129
125
134
250
138
500
133
1000
129
2000
128
4000
127
8000
122
Total Acoustic Pressure Level
141.5
Note: 0 dB corresponds to 2×10-5

Pa.

6.5.4 Shock Environment
The maximum shock experienced by satellite occurs at SC/LV separation. The shock
response spectrum (Q=10) at the separation plane is shown in Figure 6-5 for the C100
clampband and Figure 6-6 for C60 clampband.
Additionally, CGWIC could also support the customer to procure low shock level separation
system (clampband and the pyros) developed by European suppliers, if needed.

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Frequency (Hz)
Shock Response Spectrum (Q=10)
100~1000
10.5 dB/octave
1000~4000
4000 g
Figure 6-5 Shock Response Spectrum at SC/LV Separation Plane - C100 Clampband

Figure 6-6 Shock Response Spectrum at SC/LV Separation Plane - C60 Clampband

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6.6 Thermal Environment
This section covers the thermal environment during the launch vehicle flight up to separation,
which is described separately for the two flight phases, i.e. before fairing jettison and after
fairing jettison.
6.6.1 Thermal Environment prior to Fairing Jettison
During the ascent, the net thermal flux density radiated by the fairing does not exceed 500
W/m2 at any point.
This does not include any thermal input from the satellite due to power dissipation by the
satellite electronics.
6.6.2 Thermal Environment Post Fairing Jettison
The free molecular heating flux upon fairing jettison is not greater than 1,135 W/m2. This flux
is calculated as the free molecular flow acting on a plane surface perpendicular to the
velocity vector of the launch vehicle. The typical free molecular thermal flux after fairing
jettisoning is shown in Figure 6-7. The thermal effects caused by Sun radiation, Earth
infrared radiation and Earth albedo are also considered in the thermal analysis and will be
determined through the ICD and Mission Review inputs.

Figure 6-7 Free Molecular Heating Flux of LM-3A/3B/3BE/3C

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6.6.3 Launch Vehicle Generated Heat Flux
The second stage retro-rockets burn for approximately 1.5 s and will generate a heat flux of
< 300 W/m2 at the satellite separation plane.
The heat flux with the third-stage engines in operation will not exceed 350 W/m2 at the
satellite separation plane.
6.7 Pressure Environment
During the ascent phase of launch vehicle flight, the fairing is vented in order to balance the
pressure inside and outside the fairing and ensure there is no significant pressure difference.
The maximum depressurization rate inside fairing during launch vehicle flight will not exceed
6.9 kPa/s.
The typical design range for fairing internal pressure is shown in Figure 6-8.

Figure 6-8 Fairing Internal Pressure
6.8 Contamination and Cleanliness Control

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6.8.1 Contamination Control
During the satellite processing at XSLC, the maximum organic non-volatile contamination on
the satellite shall not exceed 2 mg/m2/week from opening of the satellite container in BS up
to launch.
The materials used in the clean room facilities and the parts of the launch vehicle in the
fairing have been chosen to minimize outgassing and ensure that the total organic
non-volatile contamination on the satellite does not exceed 4 mg/m2 from the satellite initial
encapsulation up to its separation from the launch vehicle.
6.8.2 Cleanliness Control
The cleanliness of satellite processing facilities is maintained to class 100,000. Any
hardware has to be subjected to a cleaning process followed by an inspection before it
enters the clean rooms.
All the materials used on the launch vehicle in the vicinity of the satellite are designed not to
accumulate dust and selected to be non-shedding during ground processing.
Prior to encapsulation of the satellite, the fairing, SC Container and Payload Adapter (PLA)
are cleaned and inspected.

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CHAPTER 7 LAUNCH SITE
7.1 General Description
This Chapter provides an overview of the facilities and services available at the Xichang
Satellite Launch Center (XSLC) for LM-3A Series launch vehicles. The detailed information
relating to the facilities and services is described in the XSLC User’s Manual (Issue 2009).

Figure 7-1 Xichang Satellite Launch Center
XSLC is mainly used to launch broadcast satellites, communication satellites and
meteorological satellites. XSLC is located in Xichang region, Sichuan Province,
southwestern China with the launch center headquarters being located in Xichang City, 65
km from the launch site. Figure 7-1 and Figure 7-2 shows the location of XSLC.
Xichang has a subtropical climate and its annual average temperature is 14.2ºC. The
ground wind in the area is usually very gentle in all the four seasons.
XSLC provides full domestic and international telephone, fax and internet services for the
user via a cable and satellite communication network.
XSLC includes the Headquarters (located in Xichang City), Launch Center (technical center
and two launch complexes), Communication Center, Mission Command and Control Center
(MCCC), tracking stations and other logistic support systems.

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Figure 7-2 Local Map of XSLC
7.1.1 Xichang Airport
Located in the northern suburb of Xichang City, Xichang Airport is 15 km away from XSLC
headquarters and 50 km away from the launch site. The 3,600 m long and 50 m wide
runway runs from south to north. The large parking apron can hold planes of various kinds,
and is fully capable of accommodating jumbo cargo aircraft such as the Boeing 747 and

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Antonov-124 (See Figure 7-3 and Figure 7-4).

Figure 7-3 Antonov-124 Landing at Xichang Airport

Figure 7-4 Unloading Operation at Xichang Airport

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7.1.2 Transfer to XSLC
The Cheng-Kun Railway and the Sichuan-Yunnan National Highway pass by XSLC. The
distance between Chengdu (Capital of Sichuan Province) and XSLC is 535 km by railway.
There is a dedicated railway branch and a highway branch leading to the Technical Center
and the launch pads of XSLC. In addition, there is a major highway, the Lu-Huang
Expressway, which links Manshuiwan (the town closest to XSLC) and Xichang City with a
dedicated highway from the Lu-Huang Expressway to XSLC. This is the primary
transportation route for the satellite from the airport to launch site, with the Sichuan-Yunnan
National Highway and the XSLC old dedicated road as a back-up route (See Figure 7-5).

Figure 7-5 Satellite Transfer from Xichang Airport to XSLC
7.2 Xichang Satellite Launch Center
XSLC includes the LV Processing Building (BL), SC Processing Buildings (BS2 and BS3),
Launch Complex No. 2 (LC-2) and Launch Complex No. 3 (LC-3). The ground support
systems are comprised of the Mission Command and Control Center (MCCC), Tracking
Stations and Communication Center.
7.3 Technical Center
The Technical Center (Figure 7-6) consists of the LV Processing Building (BL), SC
Processing Buildings (BS2 and BS3), Solid Propellant Motor (SPM) Testing and Processing

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Buildings (BM), X-ray Building (BMX), Propellant Storage Rooms (BM1 and BM2), Power
Station, etc. The LV and SC will be processed, tested, checked, assembled and stored in the
technical center. Refer to Figure 7-6 for the arrangement of the Technical Center.

Figure 7-6 Technical Center of XSLC

7.4 LV Processing (Test) Building (BL)
The LV Processing Building (BL) is composed of the Transit Building (BL1) and Testing
Building (BL2).
7.4.1 Transit Building (BL1)
BL1 is mainly used for loading and unloading the LV and other large ground equipment prior
to transiting into the test buildings. BL1 is 54 m long, 30 m wide, 13.9 m high and the railway
branch passes through BL1 to facilitate unloading. BL1 is equipped with a movable
overhead crane that has two hooks with a capability of 50 tons and 10 tons respectively. The
crane’s maximum lifting height is 9.5 m (See Figure 7-7).

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Figure 7-7 Transit Building (BL1)
7.4.2 Testing Building (BL2)
BL2 (Figure 7-8) is mainly used for the test operations, assembly and storage of the launch
vehicle. This building is 90 m long, 27 m wide and 15.58 m high, with the capability of
processing one launch vehicle and storing another one at the same time. BL2 is equipped
with a two-hook overhead movable crane with a lifting capability of 20 tons and 5 tons
respectively. The lifting height is 12 m. There are test rooms and offices inside the hall.

Figure 7-8 Testing Building (BL2)

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7.5 SC Processing Buildings
The SC Processing Buildings include the Test and Fueling Building (BS) (Figure 7-9), Solid
Propellant Motor (SPM) Testing and Processing Building (BM), X-ray Building (BMX),
Propellant Storage Rooms (BM1 and BM2).
The SC Test and Fueling Building (BS) include a Non-Hazardous Operations Building (BS2)
and a Hazardous Operations Building (BS3). All of the SC pre-transportation testing,
assembly, fuelling and SC/PLA operations will be performed in BS2 and BS3. Refer to
Figure 7-10 and Figure 7-11 for the layout of BS2 and BS3, and its area and environment
conditions are shown in Table 7-1.

Figure 7-9 SC Processing Building (BS)

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Figure 7-10 Layout of BS First Floor

Figure 7-11 Layout of BS Second Floor

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Table 7-1 Room Area and Environment in BS2
Dimensions
Door
Environment
Room
Function
L×W×H
(m mm)
Area
(m2)
W×H
(mm)
Temperature
(C)
Relative
Humidity (%)
Cleanness
level
101
Transit Hall
121818
216
5.413
18~28
5010
100,000
102
Air-lock Room
65.6413
33.8
5.412.5
18~28
5010
100,000
103
SC Test Hall
421818
756
5.412.5
(17~23)±2
35~55
100,000
105
Aisle

1.02.1

106
Compressor Room*

1.02.1

107
Clean Rooms
66.9
41.4
1.52.1
222
35~55
100,000
108
Air-conditioning Equipment Room
188.4
151.2
1.52.1

109
Clean Rooms
186.9
124.2
1.52.1
222
35~55
100,000
111
Air-conditioning Equipment Room
963
54
1.52.1
20~25
35~55

112
Storage Room
6.963
41.4
1.52.1
20~25
35~55
10,000
113
Office
363
18
1.52.1
20~25
35~55

114
Storage Room
6.963
41.4
1.52.1
15~25
35~55
10,000
115
Office
663
36
1.52.1
18~28
≤70

116
Clinic
663
36
1.52.1
18~28
≤70

117
Test Room
186.93.0
124.2
1.52.1
18~28
≤70

118
Boiler Room

15.7
1.02.1

119
Lady’s Rest Room

20
1.52.1

121
Change Room & Wind Shower Room
66.93.0
41.4
1.52.1

122
Entrance Hall

72
1.52.1

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Dimensions
Door
Environment
Room
Function
L×W×H
(m mm)
Area
(m2)
W×H
(mm)
Temperature
(C)
Relative
Humidity (%)
Cleanness
level
123
Hall Way
9.33.75
52.5
1.52.1

124
Duty Room

1.02.1

125
Office
10.56.93
72.5
1.52.1
18~28
≤70

128
Office
15.96.93
110
1.52.1
18~28
≤70

129
Security Equipment
663.0
36
1.52.1
18~28
≤70

130
Communication Terminal Room*
663.0
36
1.52.1

131
Battery Refrigerator
6.93.93.0
27
1.52.1
18~28
60

132
Communications Rack Room*
64.253.0
25.5
1.52.1

133
Test Equipment Room
186.93.0
124.2
1.52.1
18~28
30~60

134
Test Equipment Room
186.93.0
124.2
1.52.1
18~28
30~60

135
Asile
6.9613
41.4
5.012.5
18~28

136
Leakage Test
129.37
111.6
3.86
20~25
30~60

137
Leakage Control Room
63.62
21.7
1.52.1
18~28
70

138
Vacant Room
63.9
23.4
1.52.1

141
Oxidizer Fueling Equipment Storage Room
8.163.5
48.6
2.82.7
18~23
60

142
SC Fueling Control Room
8.163.5
48.6
1.52.1
18~28
60

143
Fuel Fueling Equipment Room
8.163.5
48.6
2.82.7
18~28
60

144
Fueling/Assembly Hall
181818
324
5.413
(17~23) ± 2
35~55
100,000
145
Power Distribution Room*

1.52.1

146
Air-conditioning Equipment Room*

1.0x2.1

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Dimensions
Door
Environment
Room
Function
L×W×H
(m mm)
Area
(m2)
W×H
(mm)
Temperature
(C)
Relative
Humidity (%)
Cleanness
level
201
Meeting Room
10.5x6.9
72.5

202
Meeting Room
15.9x6.9

203
Pipe line Room

204
Air-conditioning Equipment Room*

205
Air-conditioning Equipment Room*

1.52.1

206
Duty Room*

1.52.1

207
Air-conditioning Equipment Room*

1.52.1

208
Air-conditioning Equipment Room*

Note:* Referring rooms reserved by XSLC for operatio

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7.5.1 Non-Hazardous Operation Building (BS2)
7.5.1.1 General Description
The Non-Hazardous Operation Building (BS2) (Figure 7-12 and Figure 7-13) consists of the
following parts:
Transit Hall (101), Air-lock Room (102), Satellite Test Hall (High Bay, 103), System Test
Equipment (STE) rooms (133, 134), Clean Rooms (107, 109), Battery Refrigerator (131),
Leakage Test Rooms (136, 137), etc.

Figure 7-12 Non-Hazardous Operation Building (BS2)

Figure 7-13 Non-Hazardous Operation Building (BS2)

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7.5.1.2 Transit Hall (101)
Lifting Capability of the crane provided in the Transit Hall is as follows:
Main Hook: 16 tons
Subsidiary Hook: 3.2 tons
Lifting Height: 15 m
7.5.1.3 Satellite Testing Room (High Bay 103)
This is used for the satellite electric testing, integrating the solar-panels, antenna, etc. The
SC weighing and dry-dynamic-balancing are also performed in high bay 103. The High Bay
is equipped with a two hook crane that has the following lifting capability:
Main hook: 16 tons
Subsidiary hook: 3.2 tons
Lifting height: 15 m
Maximum capacity of dynamic balance instrument: 7,700 kg
A mounting fixture for the antenna is attached to the inner wall and is accessible via a ladder
and a platform to facilitate the installation of the antenna. The High Bay also has a viewing
area fitted with large glass windows for watching the testing procedure from outside. A
hydra-set is also available for the SC lifting during assembly and test.
7.5.2 Hazardous Operation Building (BS3)
The Hazardous Operation Building (BS3) (Figure 7-14) is a clean room that can support the
satellite assembly and test under hazardous conditions, such as monopropellant or
bi-propellant fuel and oxidizer fueling, the integration of the satellite solid rocket motors and
final pyrotechnic installation. The facility also supports spin balancing and weighing of the
fully loaded satellite.
7.5.2.1 General Description
The Hazardous Operation Building (BS3) comprises the following parts: SC fueling and
assembly hall (144); Oxidizer fueling-equipment room (141); Fuel fueling-equipment room
(143); Fueling operation room (142). Refer to Figure 7-14 and Table 7-1.
7.5.2.2 SC Fueling and Assembly Hall (144)
This room is used for the fueling of hydrazine or bi-propellant, the integration of the satellite
and SPM, wet satellite dynamic balancing, leakage check and SC/LV combined operations.

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An explosion-proof movable crane is equipped in this hall. The crane’s specifications are as
follows:
Main hook: 16 tons
Subsidiary hook: 5 tons
Lifting height: 15 m
The power supply, power distribution and the illumination devices are all explosion-proof.
The walls between the fueling operation room and the assembly room, leakage test room,
air-conditioning equipment room are all reinforced concrete walls for safety and protection.
The door between the fueling and assembly hall and the high-bay 103 in BS2 can withstand
the overpressure incase of an incident. A hydra-set is available for satellite assembly and
lifting.
Inside Hall 144, there are eye-wash, poisonous gas alarms and shower devices for
emergency.

Figure 7-14 Hazardous Operation Building (BS3)

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7.5.2.3 Test Equipment Room (133 & 134)
Room 133 is for system-level test equipment and room 134 is for storage of supporting test
equipment.
An RF system is provided so that the SC test equipment (STE) in BS2 can be used to
monitor the satellite whether it is in BS3, at the LC-2 or LC-3. Both uplink and downlink RF
channels are provided.
7.5.3 Solid Propellant Motor Checkout and Processing Building (BM)
7.5.3.1 General Description
The Solid Propellant Motor (SPM) Checkout and Processing Building (BM) is used for the
storage of the SPM and pyrotechnics, SPM assembly, pyrotechnics test, preparation for
X-ray inspection of the SPM, etc.
The BM consists of the following areas: Checkout and Processing Hall, SPM Storage Room,
Pyrotechnics Storage and Checkout Room, Offices, Locker Room and air-conditioning
Room. Refer to Figure 7-15. The room area and environment are listed in Table 7-2.

Figure 7-15 Solid Propellant Motor Processing Buildings (BM)

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Table 7-2 Room Area and Environment in BM
Measurement
Environment
Room
Usage
L×W×H
(m mm)
Area
(m2)
Door
W×H
(m)
T (C) Humidity
(%)
101 Duty Room
5.133.5
15.3
0.82.0

102 Break Room
3.533.5
10.5
0.82.0

103 Office
6.05.13.5
30.6
1.42.0

104 Spare Room
5.133.5
15.3
0.82.0

105 Spare Room
5.133.5
15.3
1.42.0

106 Pyro Check
5.133.5
15.3
0.82.0
215
<55
107 Pyro Storage
5.133.5
15.3
0.82.0
215
<55
108 Air-Conditioning
10.663.5
63.6
1.42.2

109 Preparation for SPM test and
X-ray inspection
1299.5
108
3.24.5
215
<55
110 SPM Storage
63.93.5
23.4
3.24.5
215
<55
7.5.3.2 Preparation Room for SPM Test and X-Ray Detection (109)
This hall is equipped with an explosion-proof movable crane that has a lifting capacity of 5 t
and a lifting height of 7 m.
There is a railway spur (1,435 mm of rail gap) in the hall leading to the SPM X-ray inspection
hall (BMX) and the cold soak chamber.
7.5.4 SPM X-ray Detection Building (BMX)
7.5.4.1 General Description
The BMX is used for X-ray inspection and cold-soak of solid motors. The BMX consists of
the following areas: cold soak room, X-ray operation hall, control room, detecting equipment
room, modulating cabinet room, film operation, processing and interpretation rooms,
chemical and instrument room, offices, locker room and air-conditioning equipment room.
Refer to Figure 7-16. The area and environment of each room are listed in Table 7-3.

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Figure 7-16 Solid Propellant Motor X-ray Building (BMX)

Table 7-3 Room Area and Environment in BMX
Dimensions
Door
Environment
Room
Function
L×W×H
(m mm)
Area
(m2)
W×H
(mm)
Temperature
(C)
Humidity
(%)
101 X-ray Detection
Hall
12.51015
125
3.24.5
(23~24)±2
35~55
103 X-ray Control Room
53.63.7
18
0.82.0
(23~24)±2
40~60
105 Modulating Cabinet
Room
53.33.7
16.5
1.52.4
(23~24)±2
40~60
106 Film Processing
Room
65.13.7
30.6
1.02.0
18~22
<70
107 Film Preparation
Room
3.65.13.7
18.36
0.82.0

108 Chemical
/Instrument Room
5.13.33.7
16.83
0.82.0

109 Film Evaluation
Room
5.13.33.7
16.83
0.82.0

7.5.4.2 X-ray Detection Room (101)
This hall is for x-ray inspection of the SPM and uses a Linatron 3000A linear accelerator.
The nominal electron beam energy is 6, 9 and 11 million electronic volts (mev). The
continuous duty-rated output and nominal energy at one meter on the central axis of the
X-ray is 3,000 rad/min.
The X-ray protection in the hall is defined according to the calculation based on the

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specifications of the Linatron 3000A. The main concrete wall is 2.5 m thick.
The doors between the hall and the control room and the large protection door are equipped
with safety locking devices. The hall is provided with dosimeter and warning device,
high-voltage emergency cut-off button for X-ray equipment, X-ray beam indicator and
additional protection systems to assure the safety of the operators.
The hall is equipped with an explosion-proof movable overhead crane with lifting height of 8
m and a telescopic arm that supports the head of the X-ray machine. A railway (1,435 mm in
width) is laid in the hall and leads to the cold-soak room and the SPM checkout and
processing hall (BM).
7.5.5 Storehouse of Hazardous Substances
Storehouses for hazardous substances are used for the storage of flammable and explosive
articles. BM1 and BM2 are for the storage of satellite propellants. There are additional
facilities for the test and storage of LV pyrotechnics.
7.5.6 Power Supply, Grounding, Lightning Protection, Fire-Detection &
Alarm
7.5.6.1 Power Supply System
All SC processing hall and rooms, such as 103, 144, 133 and 134 etc., are equipped with
two types of UPS working at 60Hz and 50Hz respectively.
(a) 60 Hz UPS
Voltage: 208 / 120V ± 1%
Frequency:
60 ± 0.5Hz
Power:
64 kVA
(b) 50 Hz UPS
Voltage: 380 / 220V ± 1%
Frequency:
50 ± 0.5Hz
Power:
130 kVA
Four kinds of power distributors are available in all SC processing halls and rooms. Each of
them has a panel with both Chinese and English descriptions indicating its frequency,
voltage, rated current, etc.
All of the sockets inside Hall 144 and other hazardous operation areas are explosion-proof.

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7.5.6.2 Lightning Protection and Grounding
In the technical areas, there are three kinds of grounding, namely technological grounding,
protection grounding and lightning grounding. All grounding resistance is lower than 1Ω.
Copper grounding bars are installed to eliminate static at the entrance of the fueling and
assembly hall, in the oxidizer fueling equipment room and the propellant fueling equipment
room.
The SPM checkout room (109), SPM storage room (110), pyrotechnics storage and
checkout rooms (106, 107) are also equipped with a copper grounding bar at the entrance to
eliminate static. In BMX and the terminal room, there are also copper grounding bars to
eliminate static electricity.
The SPM checkout and processing building is equipped with a grounding system for
lightning protection. There are two separate lightning rods outside SPM.
7.5.6.3 Fire Detection and Alarm System
The SPM checkout room (109), SPM storage room (110), pyrotechnics storage/checkout
rooms (106, 107), and air-conditioning equipment room (108) are all equipped with ionic
smoke detectors. The office (103) is equipped with an automatic fire alarm system. When
the detector detects smoke, the automatic fire alarm system will give an audible warning to
notify the safety personnel of a problem so they can take the necessary measures. The
X-ray operation hall, control room, equipment room, modulating cabinet room, film
preparation and processing room, air conditioning room are all equipped with smoke
sensors. The control room is equipped with a fire alarm system, and in case of a fire, the
alarm system will give an audible warning to notify the safety personnel of a problem so they
can take the necessary measures.
7.5.7 Launch Control Center (LCC)
The Launch Control Center (LCC) is the command and information center for the test of the
launch system during the launch campaign, from which the integrated test, functional test
and ignition and launch control are performed via the LV remote launch control ground
support system and command & monitoring equipment, and the satellite remote test station.
The Launch Control Center (LCC) is an explosion-proof blockhouse structure, and is a 3,900
m2 three-floor building.
7.6 Launch Complexes

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7.6.1 General Description
The launch site is 2.2 km (via a dedicated path) away from the Technical Center. Facilities in
the launch area consist of Launch Complex No.2 (LC-2) and Launch Complex No.3 (LC-3).
Refer to Figure 7-17.

Figure 7-17 Launch Complexes of XSLC

Figure 7-18 Two Launch Pads in XSLC (Left: LC-2; Right: LC-3)

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7.6.2 No.3 Launch Complex (LC-3)
LC-3 (Figure 7-18 and Figure 7-19) is for the LM-3A launch vehicle, and includes the launch
pad, a fixed service tower, fueling system, gas supply system, water supply system, power
supply system, lightning-proof tower, etc.

Figure 7-19 No. 3 Launch Complex (LC-3)
7.6.3 No.2 Launch Complex (LC-2)
LC-2 (Figure 7-20) is about 300 m northward from LC-3.
LC-2 includes the launch pad, a mobile service tower, a fixed umbilical tower, fueling system,
gas supply system, water supply system, power supply system, lightning-proof tower, etc.

Figure 7-20 No. 2 Launch Complex (LC-2) The Service Tower is composed of tower crane, running gear, platforms, elevators, power

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supply and distributor, fueling pipeline for storable propellant, fire detectors and
extinguishers, etc.
This tower is 90.6 m high with two cranes mounted on the top of the tower. The effective
lifting height is 84 m and the lifting capability is 20 tons (main hook) and 12.5 tons
(subsidiary hook). There are two elevators (capacity 1.5 tons) for the lifting of the personnel
and test equipment. The tower has platforms for the checkout and test operations of the
launch vehicle and the satellite.
The upper part of the tower is an environment-controlled clean area with a cleanness class
of 100,000. The temperature within this clean room, which is the satellite operation area,
can be controlled in the range of 15 ~ 25 °C. SC/LV mating, SC test, fairing encapsulation
and close-out activities will be performed in this area. A telescopic/rotating overhead crane
is provided for these operations. This crane can rotate in a range of 110 degrees and its
lifting capability is 12.5 tons.
In the Service Tower, Room 812 is exclusively prepared for SC operations and 60 Hz and 50
Hz UPS (single phase 120 V, 5 kW) are provided. The grounding resistance is less than 1 Ω.
The room area is 8 m2.
Besides a hydrant system for fire suppression in the Service Tower, it is also equipped with a
powder suppression system and 1,211 fire extinguishers.
The Umbilical Tower supports the electrical conduit, gas pipelines and liquid pipelines
between SC and LV in addition to the connectors. The Umbilical Tower uses a swinging-arm
system, platforms and cryogenic fueling pipelines through which the cryogenic propellant
fueling is performed.
The ground power supply cables will be connected to the satellite and the launch vehicle via
the Umbilical Tower. The ground air conditioning pipelines will also be connected to the
fairing via the Umbilical Tower to provide clean air into the fairing. The cleanliness of the
conditioned fairing air is of class 100,000, the temperature is 15~25°C and the humidity is
35~55%.
Umbilical Tower Room 722 is exclusively for the use of the SC team. Its area is 8 m2 and it is
equipped with a 60 Hz and 50 Hz UPS (single phase 120 V/220 V, 15 A). The grounding
resistance is lower than 1 Ω.

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7.6.4 Power Supply for Launch Pad
There are two kinds of power supplies at the launch pad:
380V / 220V (50 Hz for both) from transformer station.
120V (60 Hz) from power generators.
7.6.5 SC Fore-positioned Equipment Room (FER)
7.6.5.1 General Description
The SC FER is located in an explosion-proof building that is 150 m away from the launch
pad. The total usable area of the building is 1,000 m2, which includes LV subsystems test
rooms, SC operation rooms, air-conditioning system and emergency escape system, etc.
Figure 7-21 shows the layout.
7.6.5.2 Satellite Working Room (104,105)
There are two working rooms in LC-2 for the satellite team, see Figure 7-21. The area of
each room is 48.6 m2. The inside temperature is 18~28°C and the relative humidity is less
then 70%. The grounding resistance is less than 1 Ω. 380V / 220V, 50 Hz UPS are provided
in each room. An additional 120V/208V, 60 Hz UPS is provided in Room 104.

Figure 7-21 SC Test Rooms

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7.7 Mission Command & Control Center (MCCC)
7.7.1 General Description
The MCCC (Figure 7-22, Figure 7-23 and Figure 7-24) is located in Xichang City and
includes the Command & Control Hall, Safety Control Center, SC team room, subsystem
operation rooms and offices. The MCCC is the command, safety control, communications
and data processing center for XSLC. The MCCC has data links with the Xi’an Satellite
Control Center (XSCC) and Beijing Aerospace Control Center for the exchange of data.
Figure 7-24 shows the layout of MCCC.

Figure 7-22 XSLC Mission Command & Control Center (MCCC)

Figure 7-23 XSLC Mission Command & Control Center (MCCC)

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7.7.2 Functions of MCCC
a)
Execute instructions as the master command center;
b)
Instruct and monitor the performance and status of all units of XSLC;
c)
Perform range safety control;
d)
Record and display LV liftoff status in real-time;
e)
Transmit display information and data to Beijing Aerospace Control Center (BACC) and
Xi’an Satellite Control Center (XSCC);
f)
Transmit guidance information to every tracking station along the range;
g)
Provide SC orbit parameters to customer;
h)
Perform post-launch data processing.

Figure 7-24 Mission Command & Control Center (MCCC)

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7.7.3 Configuration of MCCC
a)
Command and control system
b)
Real-time and post-mission computer processing system
c)
Monitor and display for safety control, including computers, D/A and A/D converters, TV
display, X-Y recorders and multi-pen recorders
d)
Remote command system
e)
Communications system
f)
Timing system
g)
Data transmission system
h)
Film developing and printing equipment
7.8 Tracking, Telemetry and Control System (TT&C)
7.8.1 General Description
The TT&C system of XSLC and TT&C system of XSCC comprise a TT&C network for all
launch missions.
The TT&C system (Figure 7-25) at XSLC includes optical, radar, telemetry and remote
control equipment. It is responsible for measuring and processing the launch vehicle flight
data, and also the data for the range safety control. Data received and recorded by the
TT&C systems are used for post-mission processing and analysis.
7.8.2 Main Functions of TT&C
a)
Record the initial launch vehicle flight data in real time.
b)
Measure the trajectory of the launch vehicle in real time.
c)
Receive, record, transmit and process the telemetry data of the launch vehicle and the
satellite.
d)
Make flight range safety decisions.
e)
Compute the parameters of orbit injection and SC/LV separation status.
7.8.3 Composition of TT&C
The TT&C system of XSLC mainly consists of:
a)
Xichang Tracking Station
b)
Yibin Tracking Station
c)
Guiyang Tracking Station

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Figure 7-25 Antennas of TT&C System
The TT&C system of Xi’an Satellite Control Center (XSCC) mainly consists of:
a)
Weinan tracking station
b)
Xiamen tracking station
c)
Ocean going ships. Refer to Figure 7-26.
7.8.4 Tracking Sequence of TT&C System
After LV lift-off, the optical, telemetry and radar equipment around the launch site
immediately track the LV. The received data will be sent to the MCCC. These data will be
initially processed, and sent to the relevant stations. The station computers receive these
data and conduct the coordinate conversion and use the obtained data to guide the TT&C
system to acquire and track the flight target.
After the tracking station acquires the flight target, the measured data are sent to the
computers at the stations and MCCC at XSLC for data processing. The processed data are
used for the flight safety control. The results of computation are sent to XSCC and BACC in
real time via the data transmission system.
In case of a failure during the first and second stage flight phases, the range safety officer
will make a decision based on the range safety criteria.
The orbit injection of the SC is tracked at the XSCC facilities. The tracking data are sent to
MCCC at XSLC for processing and monitoring. The tracking operations for LM-3A, LM-3B,
LM-3BE and LM-3C launch vehicles are shown in Figure 7-26.

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Figure 7-26 TT&C Operations of LM-3A/3B/3BE/3C GTO Launch Missions
7.9 Transportation and Handling
XSLC has various types of transportation vehicle and loading/unloading equipment as well
as encapsulation handling and transportation equipment applicable for the launch vehicles
such as LM-3A, LM-3B, LM-3BE and LM-3C, etc. The technical specification of vehicles and
equipment are shown in Table 7-4.
Table 7-4 Specification of vehicle and loading/unloading equipment
No.
Vehicle
Quantity
Function
Main specification
1
3.5 tons
concave wagon 1
SC encapsulation
container
transportation
Loading capacity: 3.5 tons
Concave length: 3.58 m
Ground clearance: 0.92 m
2
13.5 tons
concave wagon 1
SC encapsulation
container or fairing
transportation
Loading capacity: 13.5 tons
Concave length: 4.1 m
Ground clearance: 1.05 m
Concave circular base: Φ4.1 m
3
12 tons SC
container trailer
1
SC container
transportation
Dead weight: 6.5 tons
Loading capacity: 12 tons
Effective width: 3,600 mm
Effective length: 12,960 mm
No-load height: 1,650 mm
4
30 tons SC
container trailer
1
SC container
transportation
Dead weight: 15 tons
Loading capacity: 30 tons
Effective width: 4,100 mm
Effective length: 15,000 mm
No-load height: 600 mm

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No.
Vehicle
Quantity
Function
Main specification
5
Air-ride flatbed
1
SC container
transportation
Loading capacity: 5 tons
Ground clearance: 1.4 m
6
“Benz” trailer
1
SC container
transportation
Loading capacity: 30 tons
Length: 12.2 m
Weight: 2.83 m
Ground clearance: 1.15 m
7
“Benz” tractor
3
SC container
transportation
Maximum pull: 50 tons
Vertical loading capacity of
connector: 14 tons
8
“Jiefang” truck

Bulk cargo
transportation
Loading capacity: 5 tons
Carriage length: 4.2 m
Width: 2.4 m
Ground clearance: 1.35 m
9
“Jiefang” trailer
3
Container
transportation
Loading capacity: 8 tons
Length: 7.95 m
Weight: 2.25 m
Ground clearance: 1.38 m
10
“STAIR” truck
1
Container
transportation
Loading capacity: 22 tons
Ground clearance: 1.45 m
11
“Huanghe”
trailers
2
Container
transportation
Loading capacity: 19 tons
Length: 7.3 m
Weight: 2.27 m
Ground clearance: 1.45 m
12
15 tons forklift
1
Loading/Unloading
Fork length: 1.8 m
(length with extender: 3.5 m)
Max lifting height: 4 m
Lifting capacity: 15 tons
Forks distance: 1.22 m~2.21 m
Max speed: 27 km/h
13
5 tons forklift
2
Loading/Unloading
Fork length: 1.2 m
Max lifting height: 3 m
Lifting capacity: 5 tons
Forks distance: 0.15 m~1.3 m
Max speed: 22 km/h
14
3 tons electric
forklift
1
Loading/Unloading
Fork length: 1 m
Max lifting height: 3.5 m
Lifting capacity: 3 tons
Forks distance: 0.15 m~1 m
15
40t mobile
crane
1
Loading/Unloading
Lifting capacity: 40 tons
Max working radius: 30 m
Arm: 35 m
Sling weight: 0.7 tons
Lift angle: 20 deg

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No.
Vehicle
Quantity
Function
Main specification
16
16 tons mobile
crane
2
Loading/Unloading
Lifting capacity: 16 tons
Max working radius: 22 m
Arm: 24.5 m
Sling weight: 0.8 tons
Lift angle: 20 deg
17
8 m mobile
elevator
2
Operation
assistant
Loading capacity: 230 kg (2
people)
Max lifting height: 8 m
Platform rotation angle: 360 deg
Max working slope: 25 deg
Ground clearance: 67 mm
18
Mobile elevator
2
Overhead
operation
Loading capacity: 200 kg
Max lifting height: 13.8 m
Platform rotation angle: 360 deg
Max elevation: 42 deg
19
Explosion-proof
mobile elevator
2
Overhead
operation
Max height: 10 m
Max operation height: 11.7 m
Capacity: 300 kg
Platform dimension: 120 x 60 cm2
Explosion-proof Class: ExdellBT4
Dead weight: 780 kg
20
CDZ32 fire
engine
1
Fire fighting/
Installation
Max working height: 32 m
Operation range: 18 m
Rotation angle: 360 deg
Swing angle between main arm
and forearm: 0~180 deg
Working platform: 360 kg (4
people)
Range of water gun: 50 m

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7.10 Propellants and Gases
SC propellants and chemicals from the satellite manufacturer are usually transported to the
launch site one month prior to the arrival of SC and stored in BM1 and BM2. The satellite
contractor packs the SC propellants and chemicals in sea containers and ships them to
Huangpu Harbor, Guangzhou in Guangdong Province. CLTC will assist in customs
clearance and arrange a special train from Huangpu to Xichang. Fueling of SC propellants is
operated by the SC team.
The on-site gas supply system can provide various gases for the launch campaign, such as
compressed air, helium and nitrogen. The gases specifications are shown in Table 7-5.
Table 7-5 Specification of Gases
Specification
Nitrogen
Air
Helium
Purity
≥ 98%
/
≥ 99.99%
Dew point
≤ -55℃
≤ -55℃
Water content < 10 ppm
Dust
≤ 50 μm
≤ 50 μm
Solid particle < 10 μm
Oil content
≤ 3 x 10-7
(Volume ratio)
≤ 3 x 10-7
(Volume ratio)
Oxygen content <10 ppm

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CHAPTER 8 LAUNCH OPERATIONS
8.1 Overview of Launch Site Operations
The launch operations at XSLC include:

SC Checkout and Processing

LV Checkout and Processing

SC and LV Combined Operations
The typical work flow and requirements of the launch operations are provided in this Chapter.
Although all launch missions are similar, each has its own special characteristics, which
means the launch operations at XSLC for each mission are unique. The main area where
differences occur is the combined operations where the satellite and launch vehicle teams
interface directly. Thus combined operations can only be successfully performed if the
operational procedures are fully coordinated and approved by both sides.
8.2 SC Operations at the Launch Site
The SC operations are conducted by the SC team with XSLC providing the necessary
facilities and support as required.
8.3 Launch Vehicle Launch Operations
The launch vehicle is transported from the CALT facilities in Beijing to XSLC by train, and
undergoes various checkouts and processing in the Technical Center and on the launch pad
of XSLC. The typical launch vehicle work flow (Encapsulation-on-Pad) for launch operations
at the launch site is listed in Table 8-1.

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Table 8-1 Typical Launch Vehicle Work Flow (Encapsulation-on-Pad) at XSLC

Activity
Working
day(s)
Accumulative
day(s)
LV unloading from train & moving to BL1
1
1
LV status recovery for system tests
1
2
Unit tests of subsystem
3
5
Technical Center
LV status preparation for transport
2
7
LV transfer, erection & integration
1
8
Preparation for system tests
1
9
Tests of subsystems
SC/LV combined operations in Technical Center in
parallel
4
13
First and second overall checkouts
SC encapsulated with container
2
15
Transfer SC & fairing to launch pad
SC/LV integration & fairing encapsulation
2
17
SC Test Period
1
18
Third overall checkout with SC involved
1
19
Fourth overall checkouts
1
20
Review on checkout results
1
21
Functional check before fueling
Gas replacement of propellant tanks
1
22
N2O4/UDMH fueling
1
23
Launch Pad
LOX/LH2 fueling and launch
1
24
Total
24 days
Note: After the SC is transferred to the launch pad, some of the SC and LV operations
can be performed in parallel as detailed in the combined operation procedures.
If encapsulation is performed in BS3, the SC and fairing are integrated as a complete unit.
The integrated SC and fairing are transferred to the launch pad, and mated with the launch
vehicle. The SC/LV mating is performed after the second overall LV checkout and before the
third overall checkout. In this case, the working procedures for the launch vehicle are
substantially the same as encapsulation on the pad.

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8.4 SC/LV Combined Operations
8.4.1 General Description
The SC/LV combined operations are conducted at the Technical Center and the launch pad.
LM-3A Series launch vehicles offer two options for the SC to LV integration:
Encapsulation-on-Pad and Encapsulation-in-BS3.
Encapsulation-on-Pad (Refer to Section 8.4.2):
The Payload Adapter (PLA) and fueled SC are mated inside BS3. The SC/PLA stack is put
into the SC Container, which is sealed and then shipped to launch pad. The SC Container
and fairing are hoisted to the 8th floor of the Service Tower of Launch Complex No.2 (LC-2).
A clean room is established there on the tower following the arrival of SC Container and
fairing. The SC/LV integration and fairing encapsulation are performed on the 8th floor of the
Service Tower.
Encapsulation-in-BS3 (Refer to 8.4.3):
The Payload Adapter (PLA) and fueled SC are mated in BS3. The SC/PLA stack is then
encapsulated with the fairing. The integrated SC/Fairing is shipped to launch pad, and
mated with launch vehicle.
The remote SC ground equipment will be moved to the launch complex and the functional
checkout will be completed before the SC/LV combined operations commence.
The launch rehearsal and launch countdown procedures are part of the SC/LV combined
operations, so the SC team is required to participate in and perform the SC operations and
checkouts, as detailed in the procedures.
8.4.2 SC/LV Combined Operations for Encapsulation-on-Pad
8.4.2.1 SC Container
SC Container is a dedicated container used for transferring the SC from BS3 to the launch
pad (Figure 8-1 and Figure 8-2).
The SC Container is 4,120 mm in diameter, 8,070 mm in height and the usable envelope of
SC Container is 3,650 mm in diameter and 6,650 mm in height for the 1194/1194A PLA. See
Figure 8-1. The SC Container is composed of a base pad and five cylindrical sections. The
five sections can be assembled together one by one. Two guide poles are provided outside

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the SC Container. The core of the wall is made of aluminum, and thermal materials cover
the outer surface and inner surface of the wall. This SC Container is only used when the
fairing Encapsulation-on-Pad method is adopted.

Figure 8-1 SC Container

Figure 8-2 SC Container Transfer

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If the satellite dimensions in launch configuration exceed the allowable static envelope of the
SC Container, the interference must be reviewed and accepted by CALT in advance of the
launch campaign.
In addition, a bigger SC Container with 3,850 mm allowable envelope diameter is also
available.
The characteristics of the SC Container are as follows:
a)
SC/PLA stack is fastened inside the SC Container;
b)
The SC Container is filled and pressurized with dry nitrogen gas (N2) so that the SC is
protected by a positive pressure;
c)
The maximum vertical and horizontal overload factor is 0.5 g;
d)
The satellite container is fitted with thermal insulation;
e)
Temperature, humidity, noise and acceleration, etc. inside the SC Container can be
monitored and recorded during the transportation.
8.4.2.2 SC/LV Integration
The PLA and the SC are mated in BS3 after the SC is fueled and weighed. The SC team
carries out all the SC operations. CALT is responsible for mating the SC with the PLA and
installing SC/LV clampband. The integration procedures are as follows:
a)
The Payload Adapter (PLA), with the clampband in the pre-installation position, is
installed on a technological stand by the CALT operations team;
b)
SC is lifted and mated with PLA by the SC operations team;
c)
Clampband is locked and fastened by the CALT operations team;
d)
SC/LV separation connectors are mated by the SC operations team;
e)
SC/PLA stack is put into the SC container by the SC operations team;
f)
SC Container is sealed and filled with dry N2 by the CALT operations team;
g)
SC Container is ready for transfer to Launch Center. Environmental sensors are already
installed on the inner side of the container, which can measure and record the inner
environmental parameters in real-time during transfer to the Launch Center.
The SC/LV integration process in BS3 is shown in Figure 8-3.

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Figure 8-3 SC/LV integration process

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8.4.2.3 SC Transfer
XSLC is responsible for using the special vehicle to transfer the SC container to the launch
pad. See Figure 8-4 and Figure 8-5. XSLC lifts the SC container, which already contains the
SC/PLA stack, onto the special vehicle and fastens the SC container to the vehicle. XSLC
drives the special vehicle from BS3 to the Service Tower at the launch pad (Figure 8-6).
When in position under the crane, XSLC releases the SC Container from the transfer vehicle
and lifts the SC Container up to the 8th floor of the Service Tower. See Figure 8-7 for the
arrangement of the SC, Container and the Fairing on the Service Tower. All the doors on the
8th and upper floors are closed to establish the clean room conditions for the SC/LV mating
operations; XSLC opens the SC container when the environmental conditions, including
temperature, humidity and cleanness, achieve the SC requirements.

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Figure 8-4 Transfer SC Container to the Launch Pad

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Figure 8-5 SC Container Transfer to the Launch Pad

Figure 8-6 SC Container Lift

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Figure 8-7 Layout on the 8th floor of the Service Tower
8.4.2.4 SC/LV Mating
The SC/LV mating operations at the launch pad include integrating the SC/PLA stack with
the launch vehicle third stage, and encapsulating the satellite in the fairing. CALT is
responsible for the launch vehicle third stage/PLA mating and fairing encapsulation in the
clean room. The SC team is responsible for the SC lifting to the third stage interface. The
combined operations procedures are as follows:
a)
XSLC team opens the SC Container when the environmental conditions inside the
clean room, including temperature, humidity and cleanness, achieve the SC
requirements;
b)
SC team begins to install the lifting slings on the SC after the container is open;
c)
XSLC team unbolts the Payload Adapter (PLA) from the base pad of the SC Container;
d)
SC team hoists and moves the SC/PLA stack to a position above the launch vehicle
third stage by using the crane in the clean room;

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e)
CALT mates the SC/PLA stack with launch vehicle third stage;
f)
SC team removes the SC lifting slings;
g)
XSLC team removes the SC Container from the Service Tower;
h)
SC team carries out the SC checkouts prior to the fairing encapsulation;
i)
CALT team carries out fairing encapsulation.
Combined operations of fairing encapsulation are shown in Figure 8-8 and Figure 8-9.

Figure 8-8 Faring Encapsulation-on-Pad

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Figure 8-9 Combined operations of fairing Encapsulation-on-Pad

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The technical status of SC should be ready for flight before the fairing encapsulation. CALT
is responsible for fairing encapsulation after both SC and launch vehicle teams confirm that
the SC and the launch vehicle are ready for fairing encapsulation.
After encapsulation of the fairing, the ground umbilical cable for the SC is connected to the
onboard cable for the SC, and the fairing air-conditioning duct is connected to the fairing and
the air-conditioning is turned on. The air-conditioning system maintains the environmental
conditions inside the fairing so the clean room controls on the Service Tower 8th floor are
relaxed. The SC team can only perform the limited closeout operations and checkouts for
the SC through the access doors on the fairing. All functional testing is performed via the
umbilical cable and the RF links. The combined operations procedures define the
requirements for launch vehicle accessibility and RF silence time that must be respected by
the SC team when performing operations on the SC.
8.4.3 SC/LV Combined Operations for Encapsulation-in-BS3
8.4.3.1 SC/LV Integration
For fairing encapsulation in BS3, the combined operation procedures are as follows:
a)
SC team makes the final closeouts on the SC ready for encapsulation in the fairing;
b)
CALT team bolts the Payload Adapter (PLA) on a technological stand;
c)
SC team lifts the SC up and moves the SC to above the PLA;
d)
CALT team mates the SC with the PLA, and then installs the clampband with the
support from the SC team;
e)
SC team checks the connection of the umbilical cable;
f)
CALT team unbolts the PLA from the working stand;
g)
SC team lifts up and moves the SC/PLA stack to the Transit Connector;
h)
CALT team bolts the SC/PLA stack on the Transit Connector;
i)
SC team removes the lifting slings from the SC/PLA stack;
j)
CALT team encapsulates the payload fairing on the SC/PLA stack.
The SC/LV integration process in BS3 is shown in Figure 8-10.

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Figure 8-10 SC/LV integration process in BS3

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8.4.3.2 SC Transfer to Launch Center
XSLC is responsible for the transport of the SC/PLA/PLF assembly from BS3 to the launch
pad using a special vehicle (Figure 8-11). XSLC drives the special vehicle from BS3 to LC-2
in the Launch Center and releases the SC/PLA/PLF stack from the transfer vehicle. The
stack is then hoisted up to the 8th floor of the Service Tower.

Figure 8-11 Transit of SC/PLA/PLF from BS3 to the Launch Pad

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8.4.3.3 SC/LV Integration
The mating of SC/PLA/PLF stack with launch vehicle is carried out on the 8th floor of the
Service Tower at LC-2 of Launch Center. The procedures of combined operation of
SC/PLA/PLF stack mating with launch vehicle are as follows:
a)
XSLC team installs the slings on SC/PLA/PLF stack, lifts the stack to the 8th floor, and
moves the stack over launch vehicle;
b)
CALT team performs the SC/PLA/PLF stack mating with launch vehicle with the support
from XSLC team. The mating is through bolts at the interface plane;
c)
XSLC team removes the slings away after Fairing stack/LV integration.
d)
Satellite ground umbilical cables are connected to the SC onboard umbilical cable;
e)
The air-conditioning duct is connected to the payload fairing and the air-conditioning for
fairing is turned on.
The SC/LV mating process is shown in Figure 8-12.

Figure 8-12 SC/LV mating process
8.4.4 SC Preparation and Checkouts
CALT and XSLC are responsible for checking and verifying the umbilical cables and RF links,
and SC team may witness the operation if requested.
The SC team can only perform the limited closeout operations and checkouts for the SC
through the access doors on the fairing. All functional testing is performed via the umbilical
cable and the RF links. The combined operations procedures define the requirements for
launch vehicle accessibility and RF silence time that must be respected by the SC team

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when performing operations on the SC.
8.5 Launch Limitation
Launch Criteria of “GO/NO GO"
a)
The average ground wind velocity in the launch area is lower than 10 m/s;
b)
The winds aloft limitation: qα ≤ 2500 N/m2•rad (qα reflects the aerodynamic loads acting
on unit cross section area of the launch vehicle, where q is the dynamic pressure, and α
is angle of attack of launch vehicle);
c)
No thunder and lightning in the area of 10 km around the launch site;
d)
The electrical field strength of the atmosphere is less than 10 kV/m.
e)
The status of satellite, launch vehicle and ground equipment are normal and ready for
launch;
f)
All personnel have evacuated the launch complex to the safe area.
8.6 Pre-launch Countdown Procedure
The typical pre-launch countdown procedure on the launch day is listed in the following
Table 8-2.

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Table 8-2 Typical Pre-Launch Countdown Procedure on the Launch Day

Time
Events
1
-7 hr LV Stage-3 LOX Fueling
2
-5 hr LV Stage-3 LH2 Fueling
3
-2 hr LV Control System power-on & functional checkout
4
-80 min LV Telemetry System power-on & functional checkout
5
-45 min Stopping & disconnecting of Fairing air-conditioning
6
-40 min Accurate aiming; Flight data presetting and checking; Gas pipe
connectors for Stage-1, Stage-2 & boosters drop-off
7
-22 min Stage-3 Engines pre-cooling
8 -15~-10 min SC umbilical disconnection
9
-4 min Telemetry and Tracking Systems power switch-over;
Fueling pipes disconnection of LV Stage-3
10
-2 min Gas pipes disconnection of LV Stage-3
11
-90 sec Control System power switch-over
12
-60 sec Umbilical disconnections of Control System & TT&C System
Swinging off the supporting arms of Umbilical Tower
13
0 sec Ignition
8.7 Post-Launch Activities
CLTC will provide the parameters of the injected orbit to the customer approximately thirty
minutes after separation. CALT will provide the launch vehicle flight evaluation report to the
customer two months after launch.

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CHAPTER 9 RANGE SAFETY CONTROL
This Chapter provides an overview of the range safety controls. The safety management for
personnel and facilities during launch site operations, including operations associated with
the hazardous materials are detailed in the XSLC User’s Manual (Issue 2009).
9.1 Responsibility and Requirements of Range Safety
XSLC designates a range safety officer, whose responsibilities are:
a)
To prepare Launch Vehicle Flight Safety Control Plan in conjunction with the LV
designer in accordance with the safety system concepts previously developed. This
includes the safety boundary line derived from the flight trajectory of the LV.
b)
To know the distribution of population and major infrastructures in the down range.
c)
To guarantee that the measuring equipment provides sufficient flight information for
safety control, i.e. a clear display of any flight anomaly or assurance of the normal flight
status inside the predetermined safe range parameters.
d)
To terminate the flight according to the Launch Vehicle Flight Safety Control Plan in
cases where the launch vehicle has demonstrated an uncorrectable anomaly such that
the flight mission needs to be terminated due to the risk of a ground impact outside the
designated safe areas.
9.2 Range Safety Issues for a LV Anomaly
9.2.1 Safety Control Strategy
From the 17th second (for LM-3A/3BE/3C) or 15th second (for LM-3B) following the LV lift off
(T0) to the moment of reaching the theoretical impact area on the open ocean, the launch
vehicle must be under the control of the range safety officer so a destruct command can be
sent if a flight anomaly occurs. This period of time is called Safety Control Range.
In the period between T0 and T0+15 sec (for LM-3B), or T0 and T0+17 sec (for
LM-3A/3BE/3C), the LV will not be destroyed by ground command even though an anomaly
occurs, unless LV explodes itself. This is to ensure that the launch vehicle flies over 400 m
(range protection radius) away from the launch pad for the protection of the launch facilities.

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Safety Control Ranges of LM-3A Series launch vehicles are defined as follows:
a)
Safety Control Range of LM-3A: The period from T0+17 sec to T2/3 (the time of
second/third stage separation);
b)
Safety Control Range of LM-3B: The period from T0+15 sec to T2/3 (the time of
second/third stage separation);
c)
Safety Control Range of LM-3BE: The period from T0+17 sec to T2/3 (the time of
second/third stage separation);
d)
Safety Control Range of LM-3C: The period from T0+17 sec to T2/3 (the time of
second/third stage separation);
9.2.2 Safety Control Procedure
The destruction of the launch vehicle will be implemented by a combination of the Command
Destruction System (CDS) and Autonomous Destruction System (ADS).
Command Destruction System
The ground tracking and telemetry system will acquire and interpret the flight information
independently. If a flight anomaly reaches the destruction criteria limits, the range safety
officer will select the impact area and send the destruction command. The ground control
computer can also automatically send a command to the on-board remote control system to
destroy the launch vehicle.
Autonomous Destruction System
The on-board ADS is able to make the decision to self destruct based on the deviations of
flight attitude and range safety criteria. If the launch vehicle attitude exceeds the safety limits,
the ADS will send a destruction signal to the on-board explosive devices. After a delay of 15
seconds, the launch vehicle will self destruct. The range safety officer can use the 15
seconds delay to select an acceptable impact location and send the destruction command,
which overrides the automated system. If the range safety officer could not find a suitable
area within 15 seconds, the ADS will automatically destroy the LV.
The objective of choosing an impact location is to ensure that launch vehicle debris does
little or no damage when it reaches the ground.
The flowchart of range safety control is shown in Figure 9-1.

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Figure 9-1 Flowchart of Range Safety Control System
9.3 Composition of the Range Safety Control System
The range safety control system includes an on-board segment and a ground segment. The
on-board safety segment functions with the onboard tracking system, and jointly is called the
Tracking and Safety System. The on-board safety control system consists of the ADS, CDS,
explosive devices, tracking system and telemetry system. The ground safety control system
consists of the ground remote control station, tracking station, telemetry station and
communication system. See Figure 9-2.
The flight data that the safety control system requires includes the flight velocity, flight
coordinates, working status of LV subsystems, and the safety command receiving status.
The working status of the onboard safety control system is provided by the Telemetry and
Tracking system. The safety command to destroy the LV is provided by the ground remote
control station (CDS) or the onboard safety control system (ADS).

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Figure 9-2 Range Safety Control System at XSLC
9.4 Safety Criteria
The range safety criteria are regulations that have been developed to allow the range safety
officer to make an informed decision regarding a launch vehicle if the designated flight
trajectory deviates from the designed flight path. This may result in the destruction of the
launch vehicle when the designated flight zone is exceeded. The criteria are developed
based on an overall consideration of the combination of the launch trajectory, regions to be
avoided as an impact zone, tracking equipment, mission objectives, etc.

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9.4.1 Approval Procedure of Range Safety Criteria
The range safety criteria vary with each launch mission, and are updated before each
launch. The criteria are drafted by XSLC, reviewed by CALT and CLTC and then
implemented by the range safety officer at each launch.
9.4.2 Common Criteria of Range Safety Control
a)
If the launch vehicle flies in the reverse direction, the safety officer will select a suitable
time to destroy the launch vehicle considering the impact area.
b)
If the launch vehicle flies vertically to the sky rather than pitching over to the
predetermined trajectory, it will be destroyed at a certain altitude.
c)
If the launch vehicle has demonstrated specific anomalies, e.g. rolling over, fire on the
vehicle, it will be destroyed at a certain time.
d)
If the launch vehicle flies beyond the predefined destruction limits (including the attitude
limits beyond which the launch vehicle becomes seriously unstable), it will be destroyed
at a certain altitude considering the impact area.
9.4.3 Special Criteria
a)
If the distance between the abnormal launch vehicle and launch pad is less than 400 m,
the launch vehicle will not be destroyed in order to protect the launch site.
b)
If the launch vehicle leaves the normal trajectory and flies towards the Technical Center
during the initial 17 to 30 seconds (LM-3A, LM-3BE, LM-3C) or 15-30 seconds (LM-3B)
and Z ≥ 400 m (in the launch coordinate system), the launch vehicle will be destroyed
immediately to protect the Technical Zone. In this case Z is the distance between the
abnormal launch vehicle trajectory and the normal trajectory plane.
c)
If launch vehicle is flying beyond the predefined safety limit towards the old MCCC
during the period of 30-60 seconds, it will be destroyed immediately to protect the old
MCCC.
9.5 Emergency Measures
Before the launch takes place, specific facilities and areas within the launch complex will be
evacuated of all non-essential personnel in accordance with the predetermined evacuation
plan.
XSLC has the following capabilities for implementing emergency procedures:

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a)
A designated emergency commander
b)
First aid team
c)
Firemen and ambulances
d)
Standby motor vehicles
e)
Helicopters
Rescue equipment and food, water, oxygen for one-day usage are available in the Launch
Center and the Technical Center.
The customer can check all the emergency equipment and verify the safety measures
implemented for range safety before the launch. Any comments or suggestions regarding
the range safety can be discussed during the launch campaign and launch site review.

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CHAPTER 10 SATELLITE DESIGN REQUIREMENTS
10.1 Introduction
This Chapter introduces the satellite design requirements from the launch vehicle
compatibility perspective, which include load conditions, coupled load analysis, safety
requirements, satellite mass and CoG, specifications of satellite qualification and
acceptance tests.
10.2 Load Conditions for Satellite Design
10.2.1 Frequency Requirement
To avoid the SC dynamic coupling with the launch vehicle, the primary frequency of SC
structure should meet the following requirement (under the condition that SC/LV separation
plane is considered as a rigid body):
The first order lateral frequency of the main mode:

> 10 Hz
The first order longitudinal frequency of the main mode: > 30 Hz
10.2.2 Loads Applied for SC Structure Design
The maximum lateral load occurs when the launch vehicle flies with a transonic speed or
under Maximum Dynamic Pressure (MDP).
The maximum longitudinal static load occurs at the first stage engines shutdown for LM-3A.
The maximum longitudinal static load occurs at the boosters’ engines shutdown for LM-3B,
LM-3BE and LM-3C.
The maximum longitudinal dynamic load occurs immediately following the first and second
stage separation.
Therefore, the following limit loads corresponding to different flight conditions are
recommended for SC design consideration. See Table 10-1 and Table 10-2.

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Table 10-1 LM-3A Limit Loads
Flight Condition
Transonic Phase
& MDP
Stage-1 Engines
Shutdown
1st /2nd Stage
Separation
Static
+ 2.0
+ 5.0
+ 1.4
Dynamic
+ 0.8
- 0.8
+ 0.8
- 5.5
+ 3.0
- 3.8
Longitudinal
Load Factor
(g)
Combined
+ 2.8
+ 5.8
- 0.5
+ 4.4
- 2.4
Lateral Load
Factor (g)
/
2.0
1.0
1.0
Table 10-2 LM-3B/3BE/3C Limit Loads
Flight Condition
Transonic Phase
& MDP
Stage-1 Engines
Shutdown
1st /2nd Stage
Separation
Static
+ 2.2
+ 5.3
+ 1.0
Dynamic
+ 0.8
- 0.8
+ 0.8
- 3.6
+ 2.7
- 3.6
Longitudinal
Load Factor
(g)
Combined
+ 3.0
+ 6.1
+ 3.7
- 2.6
Lateral Load
Factor (g)
/
1.5*
1.0
1.0

Notes for Table 10-1 and Table 10-2:
a) * means that 1.5g is effective only under the following conditions:
i) The SC frequency meets the requirement in Paragraph 10.2.1;
ii) Mass of SC: ≤ 5,100 kg for LM-3B; ≤ 5,500 kg for LM-3BE; ≤ 3,800 kg for LM-3C.
iii) The distance between SC CoG to the SC/LV separation plane: ≤1.6 m.
For specific SC, the “1.5 g” may be larger. The users should consult with CALT to
determine the actual load conditions according to the specific SC conditions.
b) Usage of the above loads in the Tables:
Loads for SC design
=
Limit loads
x
Safety Factor**
** The safety factor is determined by the SC designer. A safety factor ≥1.25 is recommended
by CALT.
c) The lateral load means the load acting in any direction perpendicular to the longitudinal
axis.
d) For a given flight condition, both lateral and longitudinal loads should be considered
simultaneously.
e) The plus sign "+" in longitudinal load factor means compression.
10.3 Coupled Load Analysis

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The SC manufacturer should provide its SC mathematical model to CALT for the Coupled
Loads Analysis (CLA). CALT will predict the SC maximum dynamic response by coupled
load analysis. The SC manufacturer should confirm that the SC can survive the predicted
environment and has an adequate safety margin. The safety factor recommended by CALT
is ≥ 1.25.
10.4 Safety Requirements
XSLC has implemented a safety program to protect the launch site personnel, the launch
site facilities and the visitors to the launch site for the launch services. The program has
several facets all based on a typical launch processing schedule for a commercial satellite
using conventional propellants and gases. The major aspects of the program are as follows:
a)
Launch site safety control;
b)
Safety approval based on the satellite safety submissions;
c)
Training of the customer and satellite manufacturer’s team in safety procedures
implemented at XSLC;
d)
Safety requirements during hazardous operations;
e)
Response to accidents and medical support.
XSLC has overall responsibility for preparing and implementing the safety procedures and is
responsible for ensuring that the safety procedures are followed. The customer and satellite
manufacturer will have the responsibility for ensuring that their respective teams have all
participated in the safety training before being allowed to work within XSLC and are
responsible for ensuring that their respective teams observe all the safety regulations at
XSLC.
All customer operations at XSLC require the approval of the XSLC safety authorities, in
particular the hazardous operations, all transportation at XSLC and combined operations.
The customer is required to demonstrate that the satellite, the support equipment and all
operations at XSLC comply with the XSLC safety regulations. This is achieved in a
maximum of three stages by the submission of reports detailing all the hazardous
subsystems, hazardous materials, hazardous procedures used or implemented at XSLC.
The customer submits these reports to CGWIC for review and approval by the relevant
authorities and a typical submission schedule, with the requirements at different stages
according to the baseline specified in Table 11-2 of this User’s Manual.

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10.5 Satellite Mass and CoG Requirements
LM-3A Series launch vehicles can accommodate a wide range of satellite mass and CoG
positions. See Table 10-3.
Table 10-3 Satellite Mass and CoG Requirement
Tolerance of CoG
LV
Mass (kg)
Axis (mm)
Transverse (mm)
LM-3A
2,600
±20
±10
LM-3B
5,100
±20
±10
LM-3BE
5,500
±20
±10
LM-3C
3,800
±20
±10
10.6 Qualification and Acceptance Test Specifications
10.6.1 Static Loads Test (Qualification)
The main SC structure should pass the static load qualification tests without damage. The
test load level should not be lower than SC design load required in Paragraph 10.2.2.
10.6.2 Sinusoidal Vibration Test
During sinusoidal vibration tests, the SC should be rigidly mounted on the shaker. The table
below specifies the vibration amplitude and acceleration level (zero-peak) at the SC/LV
interface for the SC qualification and acceptance tests. (See Figure 10-1)

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Test Load

Frequency (Hz)
Acceptance
Qualification
5-8
3.11 mm
4.66 mm
Longitudinal
8-100
0.8 g
1.2 g
5-8
2.33 mm
3.50 mm
Lateral
8-100
0.6 g
0.9 g
Scan rate (Oct/min)
/
4
2
Notes:
a) Frequency tolerance is allowed to be 2%
b) Amplitude tolerance is allowed to be 10%
c) Acceleration notching is permitted after consultation with CALT and
concurred by all parties. In no case should the notched accelerations
be lower than the results of coupled loads analysis on the interface
plane. CALT recommends that the safety factor be ≥1.25.
Figure 10-1 SC Sinusoidal Vibration Test Specification
10.6.3 Acoustic Test
The qualification and acceptance levels of the acoustic test are given in the following table
(also see Figure 10-2).

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Central Frequency of
Octave Bandwidth
(Hz)
Acceptance
Test Level
(dB)
Qualification
Test Level
(dB)
Deviation of
Test Level
(dB)
31.5
124
128
63
129
133
125
134
138
250
138
142
500
133
137
1000
129
133
2000
128
132
-2 / +4
4000
127
131
-5 / +4
8000
122
126
-5 / +5
Total Acoustic
Pressure Level
141.5
145.5
-1 / +3
Notes:
0 dB referenced to 2×10-5

Pa.
Duration of acceptance test: one minute.
Duration of qualification test: two minutes.
Figure 10-2 SC Acoustic Test Specification
10.6.4 Shock Test
The shock test level is specified in this Paragraph (See Figure 10-3 and 10-4). Such test
shall be performed once for acceptance, and twice for qualification. A tolerance of ±6.0 dB in
test specification is allowed, however, the test levels applied shall ensure that the shock

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response spectral analysis is at least 1/6 octave above the maximum and the response
acceleration values at central frequencies shall be at least 30% greater than or equal to the
nominal values given in Figures 10-3 and 10-4.

Frequency (Hz)
Shock Response Spectrum (Q=10)
100~1000
10.5 dB/octave
1000~4000
4000 g
Figure 10-3 Shock Response Spectrum at SC/LV Separation
Plane - C100 Clampband

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Figure 10-4 Shock Response Spectrum at SC/LV Separation
Plane - C60 Clampband
Additionally, CGWIC could also support the customer to procure low shock level separation
system (clampband and the pyros) developed by European suppliers, if needed.
Alternatively, the shock test can be performed through a SC/LV separation test by use of
flight SC, Payload Adapter (PLA), and separation system, all being flight hardware. This test
shall be performed once for acceptance, and twice for qualification.
10.6.5 Proto-flight Test
If the satellite is to be launched by the LM-3A Series launch vehicles for the first time but the
same type of satellite has been launched by other launch vehicles and has been
environmentally tested to levels that are equal to or greater than the qualification test level
described in Paragraph 10.6.1 to Paragraph 10.6.4, CALT will accept the following
Proto-flight test conditions:
a)
Vibration test should be performed in accordance with the scan rate of the acceptance
test and at the qualification test level as specified in Paragraph 10.6.2.
b)
Acoustic test should be performed in accordance with the test duration of the
acceptance test and at the qualification test level as specified in Paragraph 10.6.3.
c)
Shock test should be performed once according to Paragraph 10.6.4.
The test conditions for proto-flight testing should be determined by the satellite manufacturer
and CALT, and should be higher than the acceptance test levels but lower than the
qualification test levels.

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CHAPTER 11 MISSION INTEGRATION AND
MANAGEMENT
11.1 Introduction
In order to ensure the launch mission proceeds in a smooth way, CGWIC has implemented
a standard approach through a single interface.
This Chapter outlines the management process applied to contract implementation, the
customer responsibilities, the interface requirements and documentation required.
11.2 Management Overview
11.2.1 Management Organization
The Launch Services Contract (LSC, or the Contract) is the top level contractual agreement
with the Statement of Work (SOW) defining the technical requirements for the program.
The program organization for the customer’s program formed by the personnel from CGWIC,
CALT and CLTC is shown in Figure 11-1.

Figure 11-1 Program Organization
Generally, the Program Manager of CGWIC and the customer are to supervise and
coordinate the contract implementation through consultation. The Program Managers will
CALT
Program
Manager
CALT
Director
General
CGWIC
Vice President
CLTC
Vice President
CGWIC
Program
Manager
CLTC
Program
Manager
Launch
Vehicle Team
Launch Site
Team

CUSTOMER

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call upon other individuals from their respective organizations, as well as their contractors
and consultants, to participate, as necessary and appropriate, in such consultations. The
CGWIC Program Manager will process all the documentation, organize the meetings and
keep the customer advised of the schedule status.
The liaison concerning all commercial and contractual aspects of the program will be directly
established between the CGWIC Program Manager and the customer Program Manager.
Technical matters which are the responsibilities of CALT or CLTC will be passed immediately
to their nominated Program Managers.
The CGWIC Program Manager reports to the Vice President of CGWIC and is responsible
for the launch services. He or she is authorized to coordinate with the customer in
implementing the contract. The Program Manager is also responsible for all commercial
affairs concerning the launch services. He or she also directs the program management
office of CGWIC according to the instructions of the Vice President of CGWIC.
The Program Manager of CALT directly reports to the Director General of the launch vehicle
and interfaces with CGWIC to coordinate with the customer in all matters relating to the
launch vehicle and is responsible for planning launch vehicle preparation and supervising
the implementation of the instructions from the Director General of the launch vehicle.
The CLTC Program Manager, who reports to the Vice President of CLTC, interfaces with
CGWIC to coordinate with the customer in all matters concerning launch preparation
activities. He or she also directs the CLTC work team.
11.2.2 Approach for Successful Implementation of the Program
In establishing the organization for the program, CGWIC emphasizes the following:
a)
Assignment of highly competent, experienced personnel with demonstrated capability
for each responsibility;
b)
Strong engineering direction and performance verification through a dedicated team of
systems engineers;
c)
Close
engineering
and
management
relationships
between
partners
and
subcontractors, in order to minimize risk versus program schedule.
11.2.3 Program Management Reviews
The program status will be reviewed by the customer and CGWIC through documentation

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and scheduled face-to-face meetings. The meetings and reviews will be organized along
with the Technical Interchange Meetings (TIM) and can be held as required. The
conclusions of the reviews and any actions items will be agreed by CGWIC (CALT and
CLTC) and the customer. Notes or minutes documenting the agreements and actions will be
prepared and jointly signed by both parties at the end of each review. The following program
management reviews will be conducted:
a)
Contractual and Mission Integration Kick-Off Meeting
b)
ICD reviews
c)
Preliminary and Final Mission Analysis Reviews (PMAR & FMAR)
d)
Technical Interchange Meetings (if needed)
e)
Quality Reviews
f)
Site Survey
g)
LV Pre-shipment Review
h)
Launch Site meetings
i)
Launch Facilities Acceptance Review
j)
Combined Operations Review
k)
Launch Readiness Review
11.2.4 SC/LV Interface Control
The SC/LV interface is controlled by the Interface Control Document (ICD) based on the
template provided by CGWIC and details all of the customer interfaces, the customer’s
technical, documentation and the schedule requirements.
11.2.5 Launch Services Schedule
The launch services schedule should be in line with the SC mission process, the typical
program schedule is shown in Figure 11-2 and Figure 11-3.

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SC Manufacture
PMAR
TIM#1
Site Survey
PDR
Vibration,
Acoustic,
Fit-check &
Separation Shock Test
CDR
Preliminary
Design
FMA Cycle
Critical
Design
SC AIT
SC Kick-Off
Launch
Campaign
Launch
LSC
EDC
LV Manufacture
LV AIT
LV Kick-Off
PMA Cycle
FMAR
* for SC manufacturers that first work with LM-3A Series launch vehicles
LV Activities
SC Activities
1st Site Survey*

Figure 11-2 Typical Launch Services Program Schedule

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The typical schedule is based on a 24 month procurement schedule although the launch
vehicle can be delivered on a shorter schedule by special arrangement.
There is a typical schedule for the provision of the customer interface data based on a 24
month schedule, which is shown in Figure 11-3. The schedule for the specific launch is
developed in conjunction with the customer and satellite manufacturer based on the satellite
schedule.

Figure 11-3 The Typical Schedule
The typical schedule can be shorter if requested and supported by the customer.
11.2.6 Launch Campaign Schedule
The typical Launch Campaign period for LM-3A Series launch vehicle in XSLC is 24 days.
The activities assigned to each day are described in Chapter 8, Launching Operations, of
this User’s Manual.
11.2.7 Launch License and Permits
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and approvals for the launch services and the launch of the customer’s satellite have been
obtained and are in place.
11.2.8 Satellite Technology Safeguard
CGWIC commits to meet customer’s requirement to ensure the satellite technical security
during the whole process of contract implementation.
CGWIC is responsible for obtaining customs clearance without inspection and assisting in
arrangement of transportation (on a charter rented by the customer) between the China
airport of entry and Xichang airport for the satellite and its support equipment.
When the SC arrives at XLSC, the satellite processing facility is handed over to and will be
exclusively controlled by the customer during the whole launch campaign.
11.3 Launch Services Program
The launch services will include a launch vehicle dedicated for the customer, launch
operations, meetings & reviews during the whole program cycle and other related services.
The launch vehicle will be completed with Payload Fairing (PLF), Payload Adapter (PLA),
separation system and umbilical interface plus the provision of a customer designed logo on
the fairing.
There are a series of standard reviews during the LSC implementation and those the
customer may participate in are detailed in this section.
11.4 Interface Control Document (ICD)
The ICD is based on the standard ICD developed by CGWIC and is revised to indicate the
specific requirements for the mission as supplied by the customer. The inputs required from
the customer are defined in Paragraph 11.9 of this Chapter.
The ICD defines all the interfaces between the satellite and the launch vehicle, the
documentation requirements, the design verification requirements for the satellite, the
launch campaign requirements, and the mission parameters.
The ICD is developed jointly by the customer and CGWIC and formally signed by the parties.
Although the ICD can be updated as required, there are two planned revisions, the first after
the Preliminary Mission Analysis Review (PMAR) and the second after the Final Mission

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Analysis Review (FMAR). All revisions are jointly agreed and signed by the customer and
CGWIC.
The ICD is the master document for the launch services program and takes precedence
over all other technical documents.
11.5 Mission Analysis
11.5.1 Introduction
The launch services are based on a standard mission profile as defined in the User Manual,
however, each mission is tailored to the specific customer’s requirements. A brief and
preliminary assessment will be made to ensure that the customer’s injection requirements
are compatible with the launch vehicle capability.
The mission analysis process has two major stages, the completion of the preliminary
mission analysis following the SC PDR and the final mission analysis after the SC CDR that
finalizes the flight configuration.
The two stages culminate in two mission analysis reviews, the Preliminary Mission Analysis
Review (PMAR) and the Final Mission Analysis Review (FMAR), but this does not preclude
additional Mission Reviews, should the customer request them as optional services
according to the LSC.
11.5.2 Preliminary Mission Analysis
The preliminary mission analysis is based on the initial injection requirements according to
the Interface Requirement Document (IRD, see Paragraph 11.12) and ICD, and will assess
the following:
a)
Define the compliance of the satellite to the launch vehicle interfaces
b)
Initially assess the satellite environment during the launch
c)
Review and assess the satellite design compliance
d)
Review and assess the satellite test program
e)
Document and agree all deviations from the User’s Manual requirements, including
those identified in the ICD
f)
Document all open issues relating to the mission and agree a closure plan that resolves
all these issues prior to the FMAR.

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The results of the preliminary mission analysis are used to define the mission profile, the
launch site requirements, the changes required to the satellite test program and any
interface issues that impact the satellite design. The ICD will be updated and agreed based
on the agreements made during the PMAR.
The key areas included in the preliminary mission analysis and reviewed at the PMAR are
as follows.
11.5.2.1 Trajectory Analysis
The preliminary launch profile to injection and the injection accuracy will be derived from the
standard mission injection profile and assess the following topics:
a)
Verification of launch trajectory for this specific mission
b)
Flight sequence up to SC/LV separation
c)
Predicted vehicle performance
d)
Satellite mass margin against contracted lift-off mass
e)
Propellant reserves
f)
Predicted injection accuracy
g)
Review of satellite attitude requirements for the flight
11.5.2.2 Separation Analysis
The preliminary separation analysis and launch vehicle reorientation will also be presented
at the PMAR and assess the following topics:
a)
Verification of the satellite attitude at SC/LV separation
b)
Review of the SC/LV separation velocity
c)
Review of the collision avoidance maneuver
d)
Verification of the separation dynamics, including fuel slosh
11.5.2.3 Coupled Loads Analysis (CLA)
This is the preliminary CLA using the preliminary satellite dynamic model as provided by the
customer. The preliminary CLA will address the following:
a)
Modal analysis of the satellite and launch vehicle

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b)
The preliminary satellite dynamic loads under the worst case of launch vehicle
environmental loads
c)
The preliminary loads and accelerations responses of the satellite at the satellite model
nodes over the launch mission
d)
Notching requirements requested by the customer
The results of the CLA and its review at the PMAR will allow the customer to preliminarily
review the notching requirements and verify that the satellite qualification loads encompass
the predicted loads with adequate margin.
11.5.2.4 EMC and RF Compatibility Analysis
The EMC and RF compatibility analysis is required to verify the compatibility among the
satellite, the launch vehicle and the launch site during processing; to verify the compatibility
between the launch vehicle and satellite during the flight. The ICD is used as the baseline
for this analysis and allows CALT and XSLC to verify that the values provided in the ICD are
correct.
11.5.2.5 Electrical Interface Analysis
The Electrical Interface Analysis is conducted to verify the compatibility between SC/LV
Electrical Ground Support Equipment (EGSE) interfaces and SC/LV on-board interfaces,
and to design the specific umbilical cables, such as connectors’ pin allocation, shielding
specification and umbilical signal definition, according to the inputs from the satellite
manufacturer.
11.5.2.6 Mechanical Interface Analysis
The Mechanical Interface Analysis is performed to verify the compatibility of the mechanical
interface between the satellite and launch vehicle relating to the fairing, adapter, envelope,
access door and RF window in the fairing. The ICD will provide the baseline input, however,
the customer and satellite manufacturer will be required to validate the satellite dimensions
based on the actual hardware.
11.5.2.7 Thermal Analysis
The thermal analysis uses the thermal model provided by the customer as an input to the
launch vehicle thermal model to predict SC temperatures during the mission. The thermal
analysis relates to the launch mission from lift-off to separation providing a time phased
thermal profile of the thermal model nodes selected by the customer. The analysis also

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covers the ground operations so that the thermal environment of the satellite from leaving
BS to lift-off can be verified and the fairing cooling requirements can be defined.
11.5.2.8 Venting Analysis
The venting analysis is conducted to verify whether the requirement defined for the satellite
depressurization in the fairing can be satisfied during the flight, especially in the transonic
period.
11.5.3 Final Mission Analysis
The Final Mission Analysis is based on the final mission plan and injection predictions so it
becomes the formal mission baseline. The objectives of the Final Mission Analysis Review
(FMAR) are to verify the mission baseline so that the final flight software can be prepared, to
verify that the mission meets the customer requirements and to review the satellite test
program for compliance with the ICD requirements. The completion of the FMAR freezes the
mission plan and results in an update of the ICD to reflect the new agreements.
The final mission analysis will assess the following:
a)
Confirm the compliance of the satellite to the launch vehicle interfaces
b)
Confirm the satellite environment during the launch
c)
Review and assess the satellite design compliance
d)
Review and assess the satellite test program
e)
Document and agree all deviations from the User’s Manual requirements, including
those identified in the ICD
f)
Document all open issues relating to the mission from the PMAR and confirm that they
are closed
The key areas included in the final mission analysis and reviewed at the FMAR are as
follows.
11.5.3.1 Trajectory Analysis
The final launch profile to injection and the injection accuracy will be reviewed to confirm the
following:
a)
Verification of launch trajectory for this specific mission
b)
Flight sequence up to SC/LV separation

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c)
Predicted vehicle performance
d)
Satellite mass margin against the contracted lift-off mass
e)
Predicted orbital parameters of the satellite at separation
f)
Predicted injection accuracy
g)
Review of satellite attitude requirements for the flight
11.5.3.2 Separation Analysis
The final separation analysis and launch vehicle reorientation is updated and the following
topics confirmed:
a)
Verification of the satellite attitude at SC/LV separation
b)
Review of SC/LV separation velocity
c)
Review of the collision avoidance maneuver
d)
Verification of the separation dynamics, including fuel slosh
11.5.3.3 Coupled Loads Analysis (CLA)
This is the final CLA using the satellite dynamic model as updated and validated by the
customer. The final CLA will address the following:
a)
Modal analysis of the satellite and launch vehicle
b)
The final satellite dynamic loads under the worst case of launch vehicle environmental
loads
c)
The final loads and accelerations responses of the satellite at the satellite model nodes
over the launch mission
d)
Notching requirements requested by the customer
The results of the CLA and its review at the FMAR will allow CALT to review the final
notching requirements, verify that the satellite qualification loads encompass the predicted
loads with adequate margin and confirm that the satellite test program is adequate.
11.5.3.4 Electromagnetic Compatibility (EMC) Analysis
The final EMC and RF compatibility analysis verifies the complete compatibility among the
satellite, the launch vehicle and the launch site during processing as well as the
compatibility between the launch vehicle and satellite during the flight. This includes the final
launch configuration and the final satellite testing at the launch pad. The ICD is used as the

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baseline for this analysis and allows the CALT and XSLC to verify that the values provided in
the ICD are correct.
11.5.3.5 Electrical Interface Analysis
The Final Electrical Interface Analysis is conducted to verify the compatibility between
SC/LV Electrical Ground Support Equipment (EGSE) interfaces and SC/LV on-board
interfaces, and to verify the specified umbilical cable design, such as connectors’ pin
allocation, shielding specification and umbilical signal definition, according to the inputs from
the customer.
11.5.3.6 Mechanical Interface Analysis
The Mechanical Interface Analysis is performed to verify the compatibility of the mechanical
interface between the satellite and launch vehicle relating to the fairing, adapter, envelop,
access door and RF window in the fairing, etc. The ICD will provide the baseline input,
however, the customer will be required to validate the satellite dimensions based on the
actual hardware.
11.5.3.7 Thermal Analysis
The thermal analysis uses the thermal model provided by the customer as an input to the
launch vehicle thermal model. The thermal analysis relates to the launch mission from lift-off
to separation providing a time phased thermal profile of the thermal nodes selected by the
customer. The analysis also covers the ground operations so that the thermal environment
of the satellite from leaving BS to lift-off can be verified and the fairing cooling requirements
can be defined.
11.6 Verification of Satellite Design
In support of the mission analysis, the customer will be required to demonstrate that the
satellite design is able to survive the launch vehicle environment during the mission. The
customer shall be required to deliver the following test reports for CALT review and
approval.
11.6.1 Satellite Environmental Test Plan
The customer shall provide a comprehensive satellite test plan that clearly shows how they
will comply with the environmental requirements defined in this User’s Manual. The plan
shall detail the satellite manufacturer’s overall test philosophy and how this is translated into

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a qualification and acceptance test program. The plan shall provide an overview of the
environmental testing to be performed to clearly demonstrate that the satellite can meet the
ground processing and flight loads. The plan shall also include the test objectives, the
acceptance criteria, the satellite configuration for the tests with its applicability to the launch
configuration, the test methodology including the monitoring requirements, and the test
schedule showing that all testing will be completed such that the final CLA can be performed
before the FMAR. It should be noted that this is a test plan and the test specifications and
procedures are not required unless there is an issue with the test plan.
11.6.2 Satellite Environmental Test Program
A SC/PLA fit check shall be performed to verify the interface and may be combined with a
separation shock test.
CALT reserves the right to monitor the satellite environmental testing directly should there
be concerns with the margins to the dynamic environment or the notching selected.
11.6.3 Satellite Environmental Test Report
The customer shall provide a comprehensive environmental test report following the satellite
dynamic testing so that compliance with the environmental requirements of this User’s
Manual can be verified. The test report shall include the theoretical analysis required to
validate the dynamic testing, the results of the static load tests, and the dynamic test results.
The test report shall provide a summary of the test program with any anomalies clearly
identified and completed with the corrective actions taken or rationale for acceptance. The
test report shall provide evidence of adequate margin to the dynamic launch environment.
For structural elements not directly tested, an analysis shall be provided clearly showing the
assumptions, boundary conditions, results and margins.
11.7 Meetings and Reviews
This section details the formal reviews/meetings to be held between the customer and
CGWIC based on the typical launch vehicle processing schedule up to and including the
post mission review. All meetings shall be chaired by the CGWIC Program Manager, who
shall also coordinate the meeting dates, locations and the agenda with the customer. In
addition, CGWIC will issue the invitations to the participants and maintain the minutes. The
typical schedule of meetings is included in the typical launch service schedule shown above.
The meetings outlined in this section are the formal meetings required to support the

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provision of the launch services for a customer that has contracted for the launch of a single
satellite. This does not preclude special meetings being held to address specific issues as
requested by the customer or CGWIC. The baseline meeting schedule is shown in Table
11-1.
Unless otherwise stated, all dates are in months relative to the first day of the launch period
until a launch date is agreed, from which time on they are relative to the launch date.
Table 11-1 Meeting Schedule
Meeting
Estimated Date
Kick Off Meeting (KOM)
after SC PDR
Preliminary Mission Analysis Review (PMAR) 4 months after KOM
Technical Interchange Meeting (TIM)
after SC CDR
Final Mission Analysis Review (FMAR)
4 months after TIM
Launch Site Survey
6 months before launch
Launch Vehicle Pre-shipment Review (PSR)
one month before the start of
launch campaign
Combined Operations Procedure Review
During the launch campaign
Launch Readiness Review (LRR)
2 days before launch
11.7.1 Kick-Off Meeting (KOM)
This is the formal start of the launch services program where CGWIC and customer provide
an overview of the management organization, the technical status, the program schedule
including the meeting schedule and document delivery schedule. KOM will cover the
following issues:
a)
Program management and schedule
b)
Satellite program, launch requirement and SC interface requirements
c)
Launch vehicle performance and mission technical specification
d)
Launch site operation and safety
e)
Outline of ICD for the program
The first issue of the ICD will be jointly drafted by CGWIC/CALT/CLTC and customer after

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this KOM.
11.7.2 Phase 1/2/3 Safety Submission
The safety submission is required to show that the satellite and customer operations at the
launch site meet all the safety requirements of XSLC and CALT. The safety submission shall
contain a description of all the hazardous systems, testing and materials the customer will
utilize while processing the satellite at XSLC.
11.7.3 Preliminary Mission Analysis Review (PMAR)
This is a review of the initial mission analysis based on the ICD inputs with the objective of
verifying that the customer injection requirements can be met. This meeting will also result in
the first issue of the ICD for the program.
11.7.4 Technical Interchange Meeting (TIM)
Generally, customer will introduce the technical status in detail after the satellite Critical
Design Review (CDR) so that the Final Mission Analysis can be conducted. A TIM can be
held whenever needed during the program when agreed by CGWIC and the customer;
however, TIMs do not replace the scheduled technical meetings and are only held when a
specific issue requires discussion between scheduled reviews.
11.7.5 Final Mission Analysis Review (FMAR)
The objectives of the Final Mission Analysis Review (FMAR) are to verify the mission
baseline so that the flight software can be prepared, to verify if the mission meets the
customer requirements and to review the satellite test plans for compliance with the ICD
requirements. The completion of the FMAR freezes the mission plan and results in an
update of the ICD to reflect the new technical status.
11.7.6 Launch Site Survey
The review is held in XSLC around six months before launch. The satellite team will be
invited to this survey to verify that the facilities and equipment of XSLC on-site, such as the
MCCC, BS, clinic facility and launch pad, are ready to support the satellite. The review is
also to verify that the launch site facilities meet the launch requirements defined in the ICD.

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11.7.7 Launch Vehicle Pre-shipment Review (PSR)
This review is held in Beijing at least one month before the launch campaign. The purpose
of this review is to confirm that the launch vehicle meets the specific requirements in the
process of design, manufacture and testing. The launch vehicle delivery date to XSLC will
be discussed in the meeting. CALT provides a detailed report to the customer introducing
the technical configuration and quality assurance of the launch vehicle. The review is
focused on qualification status of the launch vehicle and the various interfaces and
compatibility with the satellite.
11.7.8 Launch Site Operation Meetings
Daily meetings will be held at the launch site at a mutually agreed time. The routine topics
are reporting the status of satellite, launch vehicle and launch site, applying for support from
launch site and coordinating the activities of all parties.
11.7.9 Combined Operations Procedure Review
This review will be held at XSLC following the submission of the Combined Operation
Procedure prepared by XSLC.
The combined operations procedures start with the mate of the satellite and the Payload
Adapter (PLA). The procedures will detail the interfaces between the launch vehicle team
and the customer’s team for the operation including the constraints, safety requirements,
handling requirements, personnel requirements and their responsibilities. These procedures
are to ensure that the events are coordinated and the teams do not run into facility conflicts.
The Combined Operations Procedure will be finalized by incorporating the comments put
forward in the review.
11.7.10 Launch Readiness Review (LRR)
This review is held at XSLC after the launch rehearsal and prior to launch vehicle fueling.
The review will cover the status of satellite, launch vehicle, launch facilities and TT&C
network. The launch campaign will commence the preparation of the launch vehicle fuelling
after this review.
11.8 Launch Reports after SC/LV Separation
CGWIC/CLTC/CALT will provide a series of launch reports, some in real time and the final
one after analysis of the launch vehicle performance.

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11.8.1 Orbit Injection Parameters Report
CLTC will provide the initial injection parameters, including the satellite attitude and orbit
parameters to the customer within 30 minutes after SC/LV separation.
11.8.2 Orbit Tracking Report
CGWIC also requires the customer’s tracking network to confirm acquisition of the satellite
and the orbital parameters as soon as possible after SC/LV separation.
11.8.3 Launch Mission Evaluation Report
CALT will prepare a report on the mission from lift-off to separation and deliver it to the
customer within 60 days after a successful launch or a brief introduction 15 days after a
launch failure or a launch anomaly. The report will address all the flight events, the launch
vehicle performance, the injection orbit and accuracy, the separation attitude and roll rates,
and the comparison with the predictions.
11.9 The Customer Commitment
The customer shall assign a dedicated Program Manager no later than one (1) month after
the Effective Date of the Contract (EDC). The customer shall be responsible for coordination
of technical and programmatic activities related to the execution of the LSC.
The customer shall submit and supply all the documentation and hardware specified
hereunder.
11.9.1 Documentation
The following documents shall be provided to CGWIC by customer in accordance with the
milestone schedule agreed upon by both parties:
a)
Satellite Interface Requirement Document
b)
Satellite Dynamic Model
c)
Satellite Thermal Model
d)
CAD Model
e)
Mechanical Interface Drawing
f)
Satellite Mechanical Environment Test Plan

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g)
Satellite Mechanical Environment Test Results
h)
Safety Submissions
i)
Satellite Mass Properties
j)
Satellite Launch Operation Plan
k)
Hazardous Satellite Test Procedures at XSLC
l)
Orbit Tracking Report
m) Shock Compatibility Analysis (in case the shock test is not included in the Satellite
Mechanical Environment Test Plan)
11.9.2 Hardware for Integration and Launch
The customer shall provide the following hardware in support of the mission:
a)
Satellite
b)
Ground Support Equipment and Test Equipment
c)
Propellant
The customer shall also provide the following umbilical connector halves to CGWIC: Two
pairs of connectors halves for flight (1 nominal + 1 spare)
11.10 Launch Campaign
11.10.1 Introduction
The launch campaign starts with the arrival of the customer’s advance team in XSLC, which
is followed by the arrival of the satellite at Xichang airport. The customer’s advance team
would arrive ahead of the satellite to do the launch site acceptance review so that the team
could verify that all the facilities are ready and the transportation is in place for the transfer of
the satellite to the launch site.
Prior to starting the launch campaign, all the requisite documentation shall be formally
issued and the customer will have completed the acceptance review of the launch site and
certified that the launch site is ready to receive the satellite.
The launch campaign has three major phases: satellite preparation, combined operations,
and the final preparation and launch. The satellite preparation phase includes all operations
up to mating to the Payload Adapter (PLA) and is managed by the customer with XSLC

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support as required. The combined operations start with the mating of the satellite to the
PLA and include all operations up to the mating of the encapsulated satellite with the launch
vehicle. These operations are managed by XSLC with the customer’s support for interface
verification. The final preparation and launch are also managed by XSLC with customer’s
support.
The typical launch vehicle working period (Encapsulation-on-Pad) for the launch campaign
is 24 days. The typical launch campaign flowchart (assuming the launch campaign period
for satellite is 35 days) is listed in Table 11-2.
11.10.2 Launch Campaign Preparation
In support of the customers preparation for the launch campaign, CGWIC, CALT and XSLC
will issue a series of documents covering the activities and facilities. The key documents are
as follows:
a)
XSLC User’s Manual
b)
Combined Operations Plan
c)
Combined Operations Procedures
d)
Countdown Procedure
e)
Go/No Go Criteria
f)
Launch Rehearsal Procedures
CGWIC will organize the launch site acceptance review at XSLC so that the customer can
review the facilities and certify that the launch site is ready for the satellite processing before
the launch campaign starts.
11.10.3 Launch Campaign Documentation
11.10.3.1 Introduction
This section outlines the documents to be prepared by the customer and CLTC/CALT that
are required to plan and implement the launch campaign.
11.10.3.2 Input to ICD
The ICD not only defines the satellite to launch vehicle interfaces and injection requirements
but also details the satellite support requirements at the launch site. This initial input will
cover the basic requirements that will be updated following the Mission Analysis Reviews

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Table 11-2 Typical Launch Campaign Flowchart (Encapsulation-on-Pad)
No.
Work Items
-35
-31 -30
-29
-23
-22
-21
-20
-19
-16
-15
-14
-12
-11
-10
-8
-7
-6
-4
-3 -2
-1
0
1
2
1
SC Advance Team Arrival
▌

2
Launch Facilities Check & Acceptance Review
▼-------
▼

3
SC arrival
▌

4
SC Operations in BS2
▼----------------------▼

5
LV arrival

▌

6
LV status recovery & unit test in BL

▼----------------------▼

7
SC Pre Filling Review

▌

8
SC Fueling in BS3

▼--------------▼

9
LV Transfer, Erection & Integration on Pad

▌

10 LV Subsystem Test, #1 & #2 Overall Checkout

▼------------------------------▼
11 SC Flight Readiness Review

▌

12 SC/LV Combined Operation in BS3 (SC mating with
Payload Adapter)

▼-------------▼
13 SC transfer

▌
14 SC/LV Combined Operation on Pad (Encapsulation &
SC Test)

▼------------▼
15 LV #3 & #4 Overall Checkout, Overall Checkout Result
Review

▼-----▼
16 LV Launch Readiness Review

▌
17 LV Functional check before fueling

▌
18 LV Fueling (Conventional & Cryogenic)

▼-----▼
19 Launch

20 Post-launch Operations

▼----▼

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11.10.3.3 XSLC User’s Manual
The XSLC User’s Manual (Issue 2009) is a standard document describing the facilities at
XSLC available to the customer.
11.10.3.4 Satellite Operations Plan
The Satellite Operations Plan is prepared by the customer and defines all the activities
required to handle, test and fuel the satellite from arrival at Xichang airport through to mating
to the Payload Adapter (PLA). The plan will also include all the XSLC interface requirements,
the support required for the standalone satellite operations and communications
requirements. Although the plan is primarily to define the satellite processing requirements,
it will also be required to detail the post launch activities that the customer will implement to
remove all their equipment from XSLC.
11.10.3.5 Combined Operations Plan
XSLC will prepare a combined operations plan based on the typical launch vehicle
processing schedule that will cover all activities from satellite mating to the Payload Adapter
(PLA) to the launch. The objective of the plan is to provide an overview of all the activities in
sequence between the launch vehicle and satellite with specific emphasis on the following:
a)
Operations that involve both satellite and launch vehicle
b)
All operations that are standalone and require the satellite or launch vehicle to be
inactive
c)
All hazardous operations
d)
All operational constraints from the satellite and launch vehicle
e)
The responsibilities of each party for each operation
f)
Reference to an operational procedure for each operation.
The initial release of the plan may not have the actual procedure references but will include
all the identified operations. The final version will be released following the final Combined
Operational Procedure Review and refer to all the released procedures.
11.10.3.6 Satellite Operational Procedures
These are the procedures that will be used during the launch campaign by the customer in
processing the satellite. The hazardous procedures shall detail every step required including
the constraints, safety requirements, handling requirements and personnel requirements.

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The procedures to be used during combined operations shall also detail every step required
including the constraints, safety requirements, handling requirements, personnel
requirements and any actions required by the launch vehicle in support of the operation.
The operational procedures will also include those procedures that the customer will use to
validate the XSLC interfaces for communications and monitoring the satellite.
11.10.3.7 Combined Operations Procedures
XSLC/CALT will issue the combined operations procedures starting with the mating of
satellite to Payload Adapter (PLA). These procedures will detail the interfaces between the
launch vehicle team and the customer team for the operation including the constraints,
safety requirements, handling requirements, personnel requirements and the responsibilities.
These procedures are to ensure that the events are coordinated and the teams do not run
into facility conflicts.
11.10.4 Launch Campaign Management
Although CGWIC will coordinate all the formal meetings for the launch campaign, the
customer will interface directly with the XSLC team for daily activities during the launch
campaign. The customer shall nominate a single point of contact for all interface activities
with XSLC, who will also nominate a single point of contact for the launch campaign. The
overall control of the launch campaign will be provided by XSLC as the representative of the
range.
The responsibilities of the launch campaign key personnel are as follows:
a)
Customer Representatives
Program Manager – In the case of a customer-procured launch, responsible for the
satellite launch campaign, primary interface with CGWIC/XSLC and provides formal
satellite readiness status for launch.
b)
Satellite Manufacturer Representatives
Satellite Program Manager – The satellite Program Manager has overall responsibility
for the satellite program, which includes the launch campaign. He is the sole point of
contact with the customer and CGWIC/XSLC for all formal communications.
Satellite LV Manager – the Satellite LV Manager is the primary working contact with
CALT and XSLC for the launch campaign operations.

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c)
CGWIC Representative
CGWIC Program Manager – the CGWIC Program Manager is the primary interface
with the customer Program Manager, CALT and XSLC
d)
CLTC and XSLC Representatives
CLTC and XSLC Program Manager
e)
CALT Representatives
CALT Program Manager
The launch countdown organization is based on the standard CALT and XSLC launch
program with the customer interface for providing the GO/NO GO criteria.
11.10.5 Launch Campaign Working Meetings
11.10.5.1 Introduction
This section covers the meetings required in advance of the launch campaign to plan the
activities as well as those held during the campaign to review progress and verify that the
key milestones have been achieved. The formal launch program reviews/meetings are
covered above and this section addresses those meetings held specifically to plan and
manage the launch campaign.
11.10.5.2 XSLC Facility Acceptance Meeting
Prior to the launch campaign start, the customer’s advance team will arrive at XSLC earlier
than the satellite so as to re-verify that the facilities at XSLC meet the launch requirements
and the launch site is ready for the satellite processing.
11.10.5.3 Satellite Transport Meeting
This meeting is to coordinate the transportation of the satellite from the Xichang airport to
XSLC.
The meeting will be held in conjunction with the launch site review to plan the off-loading
and transfer to the SC processing facility (BS) at XSLC.
11.10.5.4 Daily Progress and Planning Meetings
Daily progress meetings will be held between the key personnel from the customer, the
satellite manufacturer, CGWIC, CALT and XSLC. The routine topics are reporting the status

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of satellite, launch vehicle and launch site, requesting support from the launch site and
coordinating the activities of all sides.
The objective of these meetings is to verify that the satellite processing is on schedule,
coordinate support for the next few days of satellite processing, arrange services for the
satellite team and report on overall launch vehicle status. Weekly planning meetings can be
arranged if necessary.
11.10.5.5 Combined Operations Meetings
There are two combined operation meetings to verify that the satellite, the launch vehicle
and the XSLC facilities are ready for combined operations. The first Combined Operations
Meeting is held before the satellite propellant filling activities and the second is held after the
satellite is fuelled.
These meetings will review any special requirements for the satellite transportation and
verify that all the required communications links are in place and have been validated.
11.10.5.6 Launch Rehearsal Procedure Review Meeting
XSLC will arrange a dedicated meeting among the parties discussing and finalizing the
launch rehearsal procedure.
11.10.5.7 GO/NO GO Criteria Review Meeting
XSLC will arrange a dedicated meeting among the parties discussing and finalizing the
GO/NO GO criteria.
11.10.5.8 Countdown Procedure Review Meeting
XSLC will arrange a dedicated meeting among the parties discussing and finalizing the
countdown procedure.
11.11 Safety Requirements
11.11.1 Safety Training
Safety training for the SC processing facilities and general launch site safety will be provided
before the launch campaign with hands-on training in the facilities when the advance team
arrives but before the satellite arrives. The remaining team members will be trained when
they arrive to support the launch campaign.

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Training for the hazardous operations such as fuelling will be provided as required to the
customer team members nominated to participate in these operations.
11.11.2 Safety Requirements for Hazardous Operations
The customer is responsible for all the satellite and ground equipment hazardous operations;
however, these require prior approval from XSLC before implementation. XSLC safety
representatives will monitor and coordinate all hazardous operations to ensure that the
safety requirements are followed and all safety requirements are complied with.
Hazardous operations cannot commence until authorized by the XSLC safety representative.
Any incident must be reported immediately so that the XSLC safety representative can
activate the emergency response teams.
If the same satellite platform has been launched from XSLC in a previous mission and there
are no significant changes in the hazardous systems or components, a shorter safety
approval cycle can be implemented.
11.12 Launch Services Documentation
This section addresses the documentation to be provided by either the customer or the LV
side during the implementation of the launch services contract. The delivery dates are based
on typical program duration of 24 months but may be modified to adapt the customers
schedule requirements and satellite schedule.
Unless otherwise stated, all dates are in months relative to the first day of the launch period
until a launch date is agreed from which time on they refer to the launch date and are listed
in Table 11-3.

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Table 11-3 Launch Services Documentation Submission Date

Documents
Provider
Due Date
1
Launch Vehicle’s Introductory Technical Documents
Launch March User’s Manual
XSLC User’s Manual
Long March Safety Requirement Documents
Required Format of Spacecraft Dynamic Model
Required Format of Spacecraft Thermal Model
CGWIC
1 month
after EDC
2
Interface Requirement Document (IRD) for Using
LM-3A Series LV
The customer shall prepare the IRD, which generally
includes the following information:
General mission technical requirements
Launch Safety and Security Requirements
Special Requirements for LV and Launch site
The IRD is used for the start of the program. Further
detailed technical data can be defined during the
implementation of the contract.
Customer
2 months
after EDC
3
Spacecraft Dynamic Model (Preliminary and Final)
The customer shall provide hard copies or soft copies
(such as a CD) according to Required Format of
Spacecraft Dynamic Model. CALT will perform the
dynamic Coupled Load Analysis (CLA) with the model.
The customer shall specify the output requirements in
written form. The spacecraft dynamic model could be
submitted once or twice according to progress of the
program.
Customer
2 months
after No.1
4
Dynamic Coupled Load Analysis (Preliminary and
Final)
CALT will integrate the SC model, launch vehicle model
and flight characteristics together to calculate loads on
SC/LV interface at selected critical points. The customer
may get the dynamic parameters inside the spacecraft
using the analysis results. Analysis would be carried out
once or twice depending on the progress of the program.
CGWIC
3 months
after No.3

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Documents
Provider
Due Date
5
Spacecraft Thermal Model (Preliminary and Final)
The customer shall provide hard copies or soft copies
(such as a CD) of the spacecraft thermal model
according to Required Format of Spacecraft Thermal
Model. CALT will use the model for thermal environment
analysis. The analysis output requirement should be
specified in written form.
Customer
2 months
after No.1
6
Thermal Analysis (Preliminary and Final)
This analysis determines the spacecraft thermal
environment from the arrival of the spacecraft to its
separation from the launch vehicle.
CGWIC
3 months
after No.5
7
Spacecraft Interface Requirements and Spacecraft
Configuration Drawings (Preliminary and Final)
cover but are not limited to the documents listed
below:
a)
Orbital data, Mass properties, launch constraints &
separation conditions;
b)
Detailed information of the spacecraft mechanical
interfaces, electrical interfaces and RF
characteristics;
c)
Requirements and constraints for combined
operations.
The customer shall provide the spacecraft configuration
drawings to CGWIC/CALT. For all minimal or potential
protrusion out of the fairing envelope, agreement has to
be reached with CALT on its acceptability one year
before launch.
Customer
3 months
after EDC
8
Safety Phase 1 – Design
This submission is for the design of the spacecraft and
ground support equipment with details of all hazardous
systems or components. It will provide details of all safety
measures implemented, warning devices included and a
list clearly showing the risks for each hazardous system
and component.
Customer
Before the
PMAR

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Documents
Provider
Due Date
9
Mission Analysis Reports (Preliminary and Final)
f)
Trajectory Analysis
To optimize the launch mission;
To determine the launch sequence, flight trajectory and
performance margin.
g)
Separation Analysis (Statistical Analysis & Far Field
Analysis)
To analyze the separation dynamics at SC/LV
separation, including fuel slosh;
To verify the spacecraft velocity and attitude at SC/LV
separation;
To review if the LV collision avoidance maneuvers can
satisfy the spacecraft requirements.
h)
Interface Compatibility Analysis (EMC Analysis,
Electrical Interface Analysis and Mechanical
Interface Analysis)
To ensure that the SC/LV interfaces are compatible.
i)
Venting Analysis (Final Analysis Only)
To verify if the spacecraft design can satisfied and the
depressurization requirement in the fairing during the
launcher trajectory, especially during transonic period.
CGWIC
3 months
after No.7
10
Safety Phase 2 – AIT and Qualification Status
This submission provides the qualification status,
manufacturing status and acceptance verification of all
hazardous systems and components for the spacecraft
and ground support equipment. This submission will also
include the initial spacecraft operations procedures.
Customer
12 months
before
launch

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Documents
Provider
Due Date
11
Spacecraft Environmental Test Plan
The customer shall provide a comprehensive spacecraft
test plan that clearly shows how they will confirm
compliance with the environmental requirements defined
in this User’s Manual. The plan shall detail the spacecraft
manufacturer’s overall test philosophy and how this is
translated into a qualification and acceptance test
program. The plan shall provide an overview of the
environmental testing to be performed to clearly
demonstrate that the spacecraft can meet the ground
processing and flight loads. The plan shall also include
the test objectives, the acceptance criteria, the
spacecraft configuration for the tests with its applicability
to the launch configuration, the test methodology
including the monitoring requirements, and the test
schedule showing that all testing will be completed such
that the final CLA can be performed before the FMAR. It
should be noted that this is a test plan so test
specifications and procedures are not required unless
there is an issue with the test plan.
Customer
One month
before the
test
12
Spacecraft Environmental Test Report
The reports provided by the customer should include the
detailed environmental test and some related analysis
conclusions. The adaptability and the margins of the
spacecraft should also be included. The document will be
jointly reviewed by both sides to ensure the compatibility
of SC and LV.
Customer
15 days
after the
test
13
Safety Phase 3 – Final Acceptance and Hazardous
Operations
The final submission includes the results of the AIT
program for the hazardous systems plus the final details
of all hazardous operations to be conducted at the launch
site. This includes the spacecraft, ground support
equipment and deliveries of hazardous materials.
Customer
6 months
before
launch

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Documents
Provider
Due Date
14
Spacecraft Operation Plan
This plan shall describe the spacecraft operations in the
launch site, the launch team composition and
responsibilities. This plan is prepared by the customer
and defines all the activities required to handle, test and
fuel the spacecraft from arrival at Xichang airport through
to mating to the Payload Adapter (PLA). The plan will
also include all the XSLC interface requirements, the
support required for the standalone spacecraft
operations and communications requirements. Although
the plan is primarily to define the spacecraft processing
requirements, it will also be required to detail the post
launch activities the customer will implement to remove
all their equipment from XSLC.
Both sides will jointly review this document. Part of the
document will be incorporated into ICD and written into
SC/LV Combined Operation Procedure.
Customer
8 months
before
launch
15
Spacecraft Operation Procedure
These are the procedures that will be used during the
launch campaign by the customer in processing the
spacecraft. The hazardous procedures shall detail every
step required including the constraints, safety
requirements, handling requirements and personnel
requirements. The procedures to be used during
combined operations shall also detail every step required
including the constraints, safety requirements, handling
requirements, personnel requirements and any actions
required by the launch vehicle in support of the
operation.
The operational procedures will also include those
procedures that the customer will use to validate the
XSLC interfaces for communications and monitoring the
spacecraft.
Customer
2 months
before
launch

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Documents
Provider
Due Date
16
SC/LV Combined Operations Plan
XSLC will prepare a combined operations plan based on
the typical launch vehicle processing schedule that will
cover all activities from spacecraft mating to the payload
adapter to the encapsulated spacecraft mating with the
launcher. The objective is to provide an overview of all
the activities in sequence between the launch vehicle
and spacecraft with specific emphasis on the following:
a) SC/LV joint operations
b) All SC or LV stand-alone operations that require the
LV or SC to be inactive
c) All hazardous operations
d) All operational constraints from the SC/LV
e) The responsibilities of SC or LV for each operation
f) Reference to an operational procedure for each
operation.
CGWIC
8 months
before
launch
17
SC/LV Combined Operations Procedure
XSLC will issue the combined operations procedures
starting with the spacecraft mating to the payload
adapter. These procedures will detail the interfaces
between the launch vehicle team and the customer team
for the operation including the constraints, safety
requirements, handling requirements, personnel
requirements and the responsibilities. These procedures
are to ensure that the events are coordinated and the
teams do not run into facility conflicts.
The launch vehicle side will work out the SC/LV
Combined Operation Procedure based on Spacecraft
Operation Plan and SC/LV Combined Operations Plan.
Both sides will jointly review this procedure.
Both
Sides
2 months
before
launch
18
Approval of the spacecraft and ground support
equipment for fuelling and combined operations.
Customer
Before SC
fuelling

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Documents
Provider
Due Date
19
SC’s Mass Property Report (Final)
The spacecraft's mass properties are measured and
calculated after all tests, fueling and closeout operations
are completed. The data shall be provided to CGWIC
one day before SC/LV integration.
Customer
After the SC
is fueled
20
Launch Rehearsal Procedure
This document specifies the detailed operation steps
during the launch rehearsal.
Both
Sides
15 days
before
launch
21
GO/NO GO Criteria
This document specifies the baseline criteria for the
GO/NO-GO and the orders issued by the relevant
Program Managers of the mission team. The Go/No Go
criteria for the launch vehicle are standard but the criteria
will be updated with the spacecraft requirements.
The operation steps have been specified inside SC/LV
Combined Operation Procedure.
Both
Sides
15 days
before
launch
22
Countdown Procedure
The countdown procedure details the countdown
sequence, the communication links, all communication
exchanges, e.g. commands, status reports, test values,
etc., as conducted on the launch day. It also details the
responsibilities of the various participants and the inputs
required from the customer. The contingency procedures
will also be included for unplanned launch holds and
launch abort.
Both
Sides
15 days
before
launch
23
Orbit Injection Data Report
The initial orbit injection data of the spacecraft will be
provided 30 minutes after SC/LV separation. This
document will either be handed to the customer's
representative in XSLC or sent via telex, facsimile or
e-mail to a destination selected by the customer. Both
sides will sign on this document.
CGWIC
30 minutes
after SC/LV
separation

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Documents
Provider
Due Date
24
Orbital Tracking Report
The customer is required to provide the first round of
orbital tracking data after SC separation. This data is
used to verify the launch vehicle performance.
Customer
20 days
after launch
25
Launch Mission Evaluation Report
Using injection parameters and telemetry data obtained
from the launch vehicle, the launch vehicle side will
provide an assessment of the launch vehicle's
performance. This will include a comparison of flight data
with preflight predictions. The report will be submitted 60
days after a successful launch or 15 days after a launch
failure or a launch anomaly.
CGWIC
60 days
after launch

[ End of the Manual ]