KZ-1A Solid Launch Vehicle User’s Manual
KZ-1A Solid Launch Vehicle 2016.1
User’s Manual Release 1.1
Approved for Public Release ©2016 EXPACE
Distribution Unlimited All Rights Reserved
KZ-1A Solid Launch Vehicle User’s Manual
FOREWORD
This KZ-1A Solid Launch Vehicle User’s Manual is intended to provide the essential technical and process information for the customers’ spacecraft design and mission planning. The Manual also provides information of the launch site facilities, the documentations required, typical launch processing, the mission analysis requirements and other information. This Manual provides essential information of the use of launch vehicle, and the related technology details is shown in technology documents. This Manual will be updated as necessary. The updated information will be provided to existing customers. Users of this Manual are encouraged to contact the related units listed below to discuss the use of launch vehicle.
For issues: Expace Technology Co.,Ltd No.9, Jin Shan Avenue, Dongxihu District,Wuhan P. R. China P.C.:430040, Tel: +86 027 5939 3580 Fax: +86 027 5939 3500 E-mail: [email protected] Web: http://www.expace.com.cn
KZ-1A Solid Launch Vehicle User’s Manual
Revision
Approved Version Date Notes
1.0 2015.7 The first release
2.2.5, 3.3.5, 3.4, 1.1 2016.1 4.2.2, 4.3.1, 4.3.2, 4.3.3, 4.3.8, 6.2.2, 6.3,
KZ-1A Solid Launch Vehicle User’s Manual
TABLE OF CONTENTS
TABLE OF CONTENTS ...... LIST OF FIGURES ...... LIST OF TABLES ...... ABBREVIATIONS AND ACRONYMS ...... 1 Preface ...... 1 2 KZ-1A Solid Launch System ...... 1 2.1 INTRODUCTION ...... 1 2.2 KZ-1A SOLID LAUNCH VEHICLE ...... 1 2.2.1 Solid Propulsion System ...... 4 2.2.2 Vehicle Structure System ...... 4 2.2.3 Control System ...... 8 2.2.4 Integrated Propulsion & Attitude Control Subsystem ...... 9 2.2.5 Communication & Tracking System ...... 10 2.2.6 Safety Self-Destruct System ...... 11 2.3 MOBILE LAUNCH PLATFORM DESCRIPTION ...... 11 3 Vehicle Performance ...... 12 3.1 INTRODUCTION ...... 12 3.2 FLIGHT MISSION PROFILES ...... 12 3.3 LEO PERFORMANCE CAPABILITY ...... 14 3.3.1 Launches from JSLC ...... 16 3.3.2 Launches from XSLC ...... 17 3.3.3 Launches from WSLC ...... 18 3.3.4 Launches from TSLC ...... 19 3.3.5 Orbit Injection Accuracy ...... 20 3.4 PAYLOAD DEPLOYMENT ...... 20 3.4.1 Separation Attitude ...... 20 3.4.2 Separation Tip-Off ...... 20 4 Satellite Environment ...... 21 4.1 INTRODUCTION ...... 21 4.2 HORIZONTAL TRANSPORTATION ...... 21 4.2.1 Satellite Payload Environment ...... 21 4.2.2 Temperature Environment ...... 22 4.3 FLIGHT SECTION ...... 22 4.3.1 Steady State Acceleration ...... 22 4.3.2 Sinusoidal Vibration Environment ...... 23 4.3.3 Random Vibration Environment ...... 24 4.3.4 Acoustic Environment ...... 24 4.3.5 Shock Environment ...... 26 4.3.6 Thermal Environment ...... 26 4.3.7 Static Pressure Environment Under the Fairing During Ascent Phase ...... 27 4.3.8 Electromagnetic Environment ...... 27 KZ-1A Solid Launch Vehicle User’s Manual
4.3.9 Cleanliness Control ...... 28 5 Payload Interfaces ...... 29 5.1 INTRODUCTION ...... 29 5.2 REFERENCE AXES ...... 29 5.3 DYNAMIC ENVELOPE OF FAIRING ...... 31 5.4 MECHANICAL INTERFACE FOR SINGLE SATELLITE ...... 32 5.4.1 The mechanical interface of separation bolt and disengaging springs ...... 33 5.4.2 The mechanical interface of clamp band and disengaging springs ...... 34 5.5 MECHANICAL INTERFACE FOR MULTI SATELLITES ...... 38 5.6 ELECTRICAL INTERFACES ...... 39 5.7 SATELLITE DESIGN CONSTRAINT ...... 40 5.7.1 Mass Properties Accuracy Constraint ...... 40 5.7.2 Satellite Fundamental Frequency Requirement ...... 40 6 Satellite and Launch Vehicle Integration and Test ...... 41 6.1 INTRODUCTION ...... 41 6.2 FACILITIES OF JSLC ...... 41 6.2.1 Satellite Testing Facilities ...... 41 6.2.2 Vehicle Integration Facility ...... 42 6.2.3 Launch Site Facilities ...... 42 6.3 LAUNCH OPERATIONS ...... 43 6.3.1 Satellite Status Checking and Test ...... 44 6.3.2 Launch Vehicle Status Checking and Test ...... 44 6.3.3 Satellite, Vehicle and Mobile Launch Platform Integration ...... 45 6.3.4 Launch ...... 45 7 Mission Integration and Management ...... 47 7.1 INTRODUCTION ...... 47 7.2 MANAGEMENT OVERVIEW ...... 47 7.2.1 SC/LV Interface Control ...... 47 7.2.2 Launch Services Schedule ...... 47 7.2.3 Launch Campaign Schedule ...... 48 7.2.4 Launch License and Permits ...... 49 7.2.5 Satellite Technology Safeguard ...... 49 7.3 LAUNCH SERVICES PROGRAM ...... 49 7.4 INTERFACE CONTROL DOCUMENT (ICD) ...... 49 7.5 MISSION ANALYSIS ...... 50 7.5.1 Introduction...... 50 7.5.2 Preliminary Mission Analysis ...... 50 7.5.3 Final Mission Analysis ...... 53 7.6 VERIFICATION OF SATELLITE DESIGN ...... 56 7.6.1 Satellite Environmental Test Plan ...... 56 7.6.2 Satellite Environmental Test Program ...... 56 7.6.3 Satellite Environmental Test Report ...... 57 7.7 MEETINGS AND REVIEWS ...... 57 7.7.1 Kick-Off Meeting (KOM) ...... 57 KZ-1A Solid Launch Vehicle User’s Manual
7.7.2 Phase 1/2/3 Safety Submission ...... 58 7.7.3 Initial Mission Analysis Review (PMAR) ...... 58 7.7.4 Technical Interchange Meeting (TIM) ...... 58 7.7.5 Final Mission Analysis Review (FMAR) ...... 58 7.7.6 Launch Site Survey ...... 58 7.7.7 Launch Vehicle Pre-shipment Review (PSR) ...... 59 7.7.8 Launch Site Operation Meetings ...... 59 7.7.9 Combined Operations Procedure Review ...... 59 7.7.10 Launch Readiness Review (LRR) ...... 59 7.8 LAUNCH REPORTS AFTER SC/LV SEPARATION ...... 59 7.8.1 Orbit Injection Parameters Report ...... 60 7.8.2 Orbit Tracking Report ...... 60 EXPACE also requires the customer confirming acquisition of the satellite signal and the orbit parameters as soon as possible after SC/LV separation...... 60 7.8.3 Launch Mission Evaluation Report ...... 60 7.9 THE CUSTOMER COMMITMENT ...... 60 7.9.1 Documentation ...... 60 7.9.2 Hardware for Integration and Launch ...... 61 7.10 LAUNCH CAMPAIGN ...... 61 7.10.1 Introduction ...... 61 7.10.2 Launch Campaign Preparation ...... 62 7.10.3 Launch Campaign Documentation ...... 62 7.11 LAUNCH SERVICES DOCUMENTATION ...... 63 APPENDIX A ...... 70 Additional Data: ...... 75
KZ-1A Solid Launch Vehicle User’s Manual
LIST OF FIGURES
Figure 2.1 Launch vehicle theoretical configuration ...... 2 Figure 2.2 Launch vehicle structural configuration ...... 3 Figure 2.3 Front fairing configuration (Φ1.2m) ...... 5 Figure 2.4 Rear fairing configuration ...... 5 Figure 2.5 Liquid booster control segment ...... 6 Figure 2.6 Transition section ...... 6 Figure 2.7 Inter-stage Ⅱ ...... 7 Figure 2.8 Inter-stage Ⅰ ...... 8 Figure 2.9 Tail section ...... 8 Figure 2.10 Integrated propulsion & attitude control subsystem ...... 10 Figure 2.11 Erection state of mobile launch platform ...... 12 Figure 3.1 Typical Flight Profile ...... 13 Figure 3.2 Axial overload ...... 14 Figure 3.3 Launch Centers in China ...... 15 Figure 3.4 Carrying capability launches from JSLC (fairing diameter Ф1.2m) ...... 16 Figure 3.5 Carrying capability launches from JSLC (fairing diameter Ф1.4m) ...... 16 Figure 3.6 Carrying capability launches from XSLC (fairing diameter Ф1.2m) ...... 17 Figure 3.7 Carrying capability launches from XSLC (fairing diameter Ф1.4m) ...... 17 Figure 3.8 Carrying capability launches from WSLC (fairing diameter Ф1.2m) ...... 18 Figure 3.9 Carrying capability launches from WSLC (fairing diameter Ф1.4m) ...... 18 Figure 3.10 Carrying capability launches from TSLC (fairing diameter Ф1.2m) ...... 19 Figure 3.11 Carrying capability launches from TSLC (fairing diameter Ф1.4m) ...... 19 Figure 4.1 The vibration in vertical as the satellite is in transportation ...... 22 Figure 4.2 The vibration in axis and in side and as the satellite is in transportation ...... 22 Figure 4.3 Axial acceleration during the launch vehicle flight ...... 23 Figure 4.4 Sinusoidal vibration levels in axis X ...... 错误！未定义书签。 Figure 4.5 Sinusoidal vibration levels in axis Y and Z ...... 错误！未定义书签。 Figure 4.6 Random vibration levels ...... 24 Figure 4.7 Acoustic environment ...... 25 Figure 4.9 Static pressure under the fairing during ascent phase ...... 27 Figure 5.1 Coordinate system ...... 30 Figure 5.2 Dynamic envelope of Φ1.2m fairing ...... 31 Figure 5.3 Dynamic envelope of Φ1.4m fairing ...... 32 Figure 5.4 The outside view of the separator with separation bolt ...... 33 Figure 5.5 The interface view of the separator and the satellite ...... 34 Figure 5.6 The interface view of the separator in clamp band ...... 35 Figure 5.7 The interface size of the separator inΦ300 mm ...... 35 Figure 5.8 The interface size of the separator inΦ660 mm (for satellite) ...... 36 Figure 5.9 The interface size of the separator inΦ660 mm (for vehicle) ...... 37 Figure 5.10 The separator for 15 satellites in one vehicle...... 38 KZ-1A Solid Launch Vehicle User’s Manual
Figure 5.11 The separator for 2U satellite ...... 39 Figure 6.1 Flow chart of satellite status checking and test ...... 44 Figure 6.2 Flow chart of launch vehicle status checking and test ...... 45 Figure 6.3 Flow chart of satellite, vehicle and mobile launch platform integration ...... 45 Figure 7.1 Program organization ...... 错误！未定义书签。 Figure 7.2 Typical launch services program schedule ...... 47 Figure 7.3 The typical schedule ...... 48 Figure 7.4 Typical launch campaign flowchart ...... 62
KZ-1A Solid Launch Vehicle User’s Manual
LIST OF TABLES
Table 2.1 Major technical parameters of the three solid vehicle motors ...... 4
Table 3.1 Typical pre-separation payload pointing and spin rate accuracies ...... 20
Table 4.1 Typical maximum static acceleration ...... 23
Table 4.2 Payload interface sinusoidal vibration levels in axis X ...... 23
Table 4.3 Payload interface sinusoidal vibration levels in axis Y and Z ...... 24
Table 4.4 Payload interface random vibration levels ...... 24
Table 4.5 Payload acoustic environment ...... 24
Table 4.5 Shock response spectrum ...... 26
Table 4.8 Related parameters of wireless devices ...... 27
Table 4.9 Working electromagnetic radiation intensity of each device ...... 28
Table 6.1 Typical Pre-Launch Countdown Procedure ...... 46
Table 7.1 Meeting schedule...... 57
Table 7.2 Launch services documentation submission date ...... 63
KZ-1A Solid Launch Vehicle User’s Manual
ABBREVIATIONS AND ACRONYMS
Acronym Meaning
Ω Ohm(s)
° Degree(s)
°C Degree(s) Celsius
A
A Ampere(s) a Semi-major Axis
B
BD BeiDou Navigation Satellite System
C
CAN Controller area network
CASIC China Aerospace Science & Industry Corporation
EXPACE China Volant Industry CO., LTD
EXPACE China Space Sanjiang Group Corporation
D dB Decibel(s) dBW Decibel(s) Relative to 1 Watt dBμV Decibel(s) Relative to 1 Microvolt
E e Eccentricity
F
F Frequency
FM Frequency Modulation
G g Acceleration of Gravity
GPS Global Position System KZ-1A Solid Launch Vehicle User’s Manual
Acronym Meaning
Grms Root-Mean-Square Acceleration
H
H Flight Height
Hr Hour(s)
Hz Hertz
I i Inclination
IMU Inertial Measurement Unit
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
LEO Low Earth Orbit
M m Meter(s)
Ma March number ms Millisecond(s) min Minute(s) mm Millimeter(s)
Mpa Mega Pascal(s) m/s Meter(s) per second
N
N Newton(s) KZ-1A Solid Launch Vehicle User’s Manual
Acronym Meaning
N/A Not Applicable
Nx Axial Overload
O
Oct Octave
P
Pa Pascal(s)
Q
Q Quality Factor
R rad Radian(s)
S
SSO Sun-synchronous Orbit s, sec Second(s)
PSD Power Spectral Density
SPL Sound Pressure Level
T t Ton(s)
TSLC Taiyuan Satellite Launch Center
U
UPS Uninterrupted Power Supply
V
V Volt, Velocity
VF Flight Velocity
W
W Watt(s)
W/m2 Watt(s) Per Square Meter
WSLC Wenchang Spacecraft Launch Center
X KZ-1A Solid Launch Vehicle User’s Manual
Acronym Meaning
XSLC Xichang Satellite Launch Center
KZ-1A Solid Launch Vehicle User’s Manual
1 Preface
KZ-1A is a kind of high reliability, high precision and low cost solid launch vehicle developed by
China Space Sanjiang Group Corporation (EXPACE). The launch vehicle can send 200kg payload into 700km sun-synchronous orbit. It mainly offers the service of sending small satellite into Low
Earth Orbit (LEO) to domestic and international customers.
KZ-1A solid launch vehicle adopts mobile launch platform, integrated power supply equipment, test and launch control facilities, aiming facility and temperature control facility, to carry vehicle from technical support center to launch site, complete temperature control of payload, vehicle test and launch.
This user’s manual describes the system design of KZ-1A and the service offered to the user. The information in this manual is only effective in the initial mission design. The detailed research and design is determined by the engineering analysis of specific mission. The analysis results would be recorded in the interface control file for the users to develop or design satellite.
2 KZ-1A Solid Launch System
2.1 Introduction
This chapter introduces KZ-1A solid launch Vehicle system, including its major subsystems, and the mobile launch platform.
2.2 KZ-1A Solid Launch Vehicle
KZ-1A solid launch vehicle is 20m long with lift-off mass of 30tons, and its maximum diameter is
1.4m. The vehicle’s power is provided by three solid motors and one liquid motor. The vehicle theoretical configuration is shown in Figure 2.1. The vehicle constitutes from 9 segments, including front fairing, propulsion control cabin, rear fairing, transition section, stage-3 motor, inter-stage 2, stage-2 motor, inter-stage 1, stage-1 motor and tail section. The structural
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Figure 2.1 Launch vehicle theoretical configuration
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Figure 2.2 Launch vehicle structural configuration
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KZ-1A solid launch vehicle consists of the following major subsystems: solid propulsion system, vehicle structure system, control system, integrated propulsion & attitude control subsystem, communication & tracking system, and safety self-destruct system.
2.2.1 Solid Propulsion System
The solid propulsion system consists of three solid vehicle motors to provide power during first stage flight, second stage flight and third stage flight. All of the three solid motors choose single fixed nozzle, and do not shut off until the propellant is exhausted. The technical parameters of three motors are shown in Table 2.1. Table 2.1 Major technical parameters of the three solid vehicle motors
Stage-1 Stage-2 Stage-3 S/N Parameter (unit) motor motor motor
1 Diameter（mm） φ1400 φ1400 φ1202
2 Total mass(kg) 16621 8686 3183
3 Burning time(s) 65 62 55
4 Impulse (N.s/kg) 2352 2810 2850
2.2.2 Vehicle Structure System
The function of the vehicle structure system 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.
2.2.2.1 Fairing
Front fairing is available for two different diameter Φ1200mm andΦ1400mm according to the space demand of satellite. The front fairing structural configuration is shown in Figure 2.3.
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Figure 2.3 Front fairing configuration (Φ1.2m)
The same two half-hoods mating constitute the cylindrical shell of rear fairing. They are connected or unlocked by the separation bolt in the radial direction. Front end is connected with the propulsion control cabin. Rear end is connected with transition section. The rear fairing provides an excellent aerodynamic shape for the vehicle, bears the heat and vibration during the flight and supports the propulsion control cabin. The rear fairing structural configuration is shown in Figure
2.4.
Figure 2.4 Rear fairing configuration
2.2.2.2 Propulsion Control Cabin
Propulsion control cabin is composed of the integrated liquid propulsion & attitude control subsystem and the control system of the vehicle. The payload and the ipropulsion control cabin are connected by separator. Propulsion control cabin is shown in Figure 2.5.
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Figure 2.5 Liquid booster control segment
2.2.2.3 Transition Section
Transition section is located between propulsion control cabin and stage-3 motor. The front segment is in rear fairing, in which there is the telemetry device, external measure device and control system device. Transition section is shown in Figure 2.6.
Figure 2.6 Transition section
2.2.2.4 Inter-stage Ⅱ
Inter-stage Ⅱ connects Stage-2 motor and Stage-3 motor, and bears the load of the process of hoisting, transportation, erection and the flight loading and realizes the separation according to the schedule. Inter-stage Ⅱ is shown in Figure 2.7.
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Figure 2.7 Inter-stage Ⅱ
2.2.2.5 Inter-stage Ⅰ
Inter-stage Ⅰ is a cylinder segment which connects stage-1 motor and stage-2 motor. It bears the load of the process of hoisting, transportation, erection and the flight loading and realizes the separation according to the schedule. Inter-stage Ⅰ is shown in Figure 2.8.
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Figure 2.8 Inter-stage Ⅰ
2.2.2.6 Tail Section
Tail section is a cylinder segment located in the rear of vehicle, which is the protector of the electronic devices and the main load-bearing structure when the vehicle is launched. Lattice rudders and the connectors are settled outside. Tail section is shown in Figure 2.9.
Figure 2.9 Tail section
2.2.3 Control System
The control system accomplishes all kinds of static tests cooperated with ground support
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The vehicle avionics employs bus organization. All the electronic devices are miniaturization.The guidance plan employs rate strap down Inertial Navigation System (INS) with fiber gyroscopes and GPS/BD2. The first stage is operated under the preinstalled program, and the initial flight adopts lateral jet stabilization control. When the speed reaches a certain value, lattice rudders take over the control. The next two stages adopt perturbation guidance and lateral jet as executor. The final stage employs closed-loop guidance which adopts liquid propulsion and attitude control power system to control the attitude and achieve high precision orbit finally. Security mechanism and pyrotechnic control mechanism are employed to ensure the safety of the vehicle.
2.2.4 Integrated Propulsion & Attitude Control Subsystem
Integrated propulsion & attitude control subsystem adopts N2O4 and MMH as propellant. The main function of the integrated propulsion & attitude control subsystem is to control attitude of the vehicle from stage Ⅱ to Ⅳ, and provide power into orbit for the final stage. Integrated propulsion & attitude control subsystem is shown in Figure 2.10.
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Figure 2.10 Integrated propulsion & attitude control subsystem
2.2.5 Communication & Tracking System
Communication & tracking system based on communication satellites and space measurement & control station accomplish all the telemetry measurement, exterior measurement, remote control and data communication from the vehicle launch to the separate of satellite and launch vehicle.
Telemetry subsystem is responsible for measuring and sending the environmental data, control system data, attitude control subsystem and exterior subsystem data. Exterior subsystem is responsible for tracking and measuring the flight trajectory of vehicle power flight phase, providing vehicle real-time-data and post trajectory data, which is valuable in monitor the performance of the vehicle. Remote control subsystem can upload the control command for the last stage and safety command of the vehicle. Communication subsystem offer temporary communication service after the separate of the launch vehicle and satellite, download the key telemetry parameters and upload key remote control command for satellite.
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2.2.6 Safety Self-Destruct System
Safety self-destruct system is used to destroy the launch vehicle when it is broken. It is based on vehicle autonomous safety control and ground radio control as supplementary.
2.3 Mobile Launch Platform Description
Mobile launch platform mainly includes transport and launch vehicle, test and fire control equipment, aiming equipment and etc., which transfers the vehicle from technological area to launching area. The platform accomplish the temperature and environment control of payload, vehicle test and launching by using power supply equipment, test and fire control equipment, aiming equipment, temperature control device and etc, which are integrated in platform. The transport and launch vehicle employs semi-trailer which is usually used for transporting container.
It is composed of frame, suspension, axle, wheel, brake, steering, electronic devices, semi-support equipment, drag element and etc. The vehicle and erecting equipment are installed in a container whose length is over 21700 mm. The mobile launch platform is shown in Figure 2.11.
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Figure 2.11 Erection state of mobile launch platform
3 Vehicle Performance
3.1 Introduction
This chapter introduces the vehicle performance.
3.2 Flight Mission Profiles
This section describes low earth orbit mission profiles. The profile of a typical mission performed by KZ-1A is shown in Figure 3.1. Launch vehicle flight Sequence includes 1st stage powered-flight phase, 1st coasting-flight phase, 2nd stage powered-flight phase, 2nd stage coasting-flight, 3rd stage coasting-flight phase I, 3rd stage powered-flight phase, 3rd stage coasting-flight phase II and terminal boost phase, the typical flight sequence is shown in Figure
3.1.
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Flight Profile Time(s) Height(km)
1st stage separation 83.0 36.0
2nd stage separation 161.0 105.0
Fairing jettison 176.1 120.0
3rdstage ignition 192.1 133.8
3rd stage separation 284.2 245.5
4th stage ignition 287.2 249.7
4th stage shut down 1052.2 700.3
SC/LV separation 1060.2 700.3
Figure 3.1 Typical Flight Profile
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The levels of overload are shown in Figure 3.2.
Figure 3.2 Axial overload
3.3 LEO Performance Capability
There are four launch centers in China, i.e. Xichang Satellite Launch Center (XSLC), Jiuquan
Satellite Launch Center (JSLC), Taiyuan Satellite Launch Center (TSLC) and Wenchang
Spacecraft Launch Center (WSLC), shown in Figure 3.3.
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Figure 3.3 Launch Centers in China
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3.3.1 Launches from JSLC
Figure 3.4 Carrying capability launches from JSLC (fairing diameter Ф1.2m)
Figure 3.5 Carrying capability launches from JSLC (fairing diameter Ф1.4m)
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3.3.2 Launches from XSLC
Figure 3.6 Carrying capability launches from XSLC (fairing diameter Ф1.2m)
Figure 3.7 Carrying capability launches from XSLC (fairing diameter Ф1.4m)
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3.3.3 Launches from WSLC
Figure 3.8 Carrying capability launches from WSLC (fairing diameter Ф1.2m)
Figure 3.9 Carrying capability launches from WSLC (fairing diameter Ф1.4m)
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3.3.4 Launches from TSLC
Figure 3.10 Carrying capability launches from TSLC (fairing diameter Ф1.2m)
Figure 3.11 Carrying capability launches from TSLC (fairing diameter Ф1.4m)
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3.3.5 Orbit Injection Accuracy
KZ-1A could correct the trajectory deviation caused by the first three solid motors to achieve a high orbit injection accuracy, which is shown as follows: a) Semi-major Axis Deviation: |Δa|≤2km; b) Inclination Deviation: |Δi|≤0.05º; c) Eccentricity Deviation: e≤0.002.
3.4 Payload Deployment
For typical missions, KZ-1A coasts for approximately 8 seconds after Stage 4 motor shuts down
(orbit injection) to meet the required angular velocities for deployment. When the angular velocity has been met, satellite separation starts, after 10 seconds, the collision avoidance program begins to work.
3.4.1 Separation Attitude
The separation attitude is selected by the satellite mission planners to satisfy payload requirements.
KZ-1A provides a range of deployment direction options, including inertial, orbit-track-relative, and sun-pointing. Table 3.1 provides the typical KZ-1A payload pre-separation attitude angle accuracy and angular velocity accuracy.
Table 3.1 Typical pre-separation payload pointing and spin rate accuracies
Parameter Angle Rate
Pitch ±1° ±0.3°/s
3-Axis Yaw ±1° ±0.3°/s
Roll ±1° ±0.3°/s
3.4.2 Separation Tip-Off
Payload tip-off refers to the angular velocities imparted to the payload by the separation event.
These velocities are typically the result of initial body rates and unevenly distributed forces and
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Typically, the KZ-1A separation system provides a separation velocity greater than 0.5m/s and a total angular tip-off rate of 1.5 °/s or less. Note that these performance predictions are highly dependent on the satellite mass properties.
4 Satellite Environment
4.1 Introduction
The following section presents the maximum satellite environment levels during KZ-1A transportation and launch. The levels defined in this section are intended to be representative of a typical mission (transportation, flight). Mission-specific analyses are performed as a standard service and are documented in the mission ICD.
4.2 Horizontal Transportation
4.2.1 Satellite Payload Environment
The overload of the satellite or vehicle would not exceed 2g in horizon and 1g in axis. The specific value can be evaluated by the braking distance of the launch platform. The braking distance is inverse to the overload in axis when the speed is at the same level.
When the satellite is in transportation, the vibration caused by the vehicle would affect the satellite.
The details are shown in Figure 4.1 and Figure 4.2.
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PSD(g2/Hz) 0.01
0.002
5 50 200 f(Hz)
Figure 4.1 The vibration in vertical as the satellite is in transportation (effective order: 1.03g) PSD(g2/Hz) 0.005
0.001
5 50 200 f(Hz)
Figure 4.2 The vibration in axis and in side and as the satellite is in transportation (effective
order: 0.73g) Note: The description above is from practical measurement. It is not the same as the traditional transportation simulating vibration requirements. The transportation requirements is that the maximum speed should be lower than 50 km/h and the average speed is lower than 35 km/h and the distance is lower than 50 km.
4.2.2 Temperature Environment The temperature would be controlled between 15 ℃ and 25 ℃ when the satellite in the fairing is in transportation.
4.3 Flight Section
4.3.1 Steady State Acceleration
The launch vehicle flight generates external forces on the satellite due to the engine thrust and aerodynamic forces. The typical maximum axial steady state acceleration during the launch vehicle powered flight is shown in Table 4.1 and Figure 4.3. It can be seen that the maximum axial static acceleration occurs during the third stage flight. The maximum static acceleration will vary slightly with the satellite mass. The typical maximum lateral steady state acceleration is within
±1.0g.
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Table 4.1 Typical maximum static acceleration
Axial(g) Lateral(g)
During First Stage Flight +4.6
During Second Stage Flight +7.3 ±1.0 During Third Stage flight +11.8
During Fourth Stage flight +0.4
Figure 4.3 Axial acceleration during the launch vehicle flight
4.3.2 Sinusoidal Vibration Environment
Low level sinusoidal vibration environment during the flight is mainly present at the significant vibration events such as during engine ignition and engine shut-off, transonic flight and the stages separations. The environmental adaptability to this low level sinusoidal vibration can be tested by sinusoidal vibration test. The sinusoidal vibration at the payload interface is shown in Table 4.2,
4.3. Table 4.2 Payload interface sinusoidal vibration levels in axis X
Frequency (Hz) 5~8 8~75 75~100
Level A 4.7mm 1.2g 0.9g Amplitude Level C 7.05mm 1.8g 1.35g
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Table 4.3 Payload interface sinusoidal vibration levels in axis Y and Z
Frequency(Hz) 5~8 8~75 75~100
Level A 3.9mm 1.0g 0.9g Amplitude Level C 5.85mm 1.5g 1.35g
4.3.3 Random Vibration Environment
Random vibration environment during the flight mainly comes from the dynamical environment duo to aerodynamic acoustics, thrust ripple of solid vehicle motors and jet acoustics. The maximum expected random vibration levels at the payload interface are shown in Table 4.4 and
Figure 4.4. Time of duration in level A and level C is 1.5 minutes and 3 minutes. Table 4.4 Payload interface random vibration levels
Frequency(Hz) 10-80 80-200 300-800 800-2000 value
Level A +3dB/oct 0.045 0.02 -6dB/oct 5.46g PSD(g2/Hz) Level C +3dB/oct 0.09 0.04 -6dB/oct 7.72g
PSD(g2/Hz) 0.09
0.045 0.04 +3dB/Oct Group C 0.02 -6dB/Oct Group A
10 80 200 300 800
2000 f(Hz) Figure 4.4 Random vibration levels
4.3.4 Acoustic Environment
The acoustic noise is come from the first stage solid motor. The maximum expected acoustic levels within 1.4m diameter fairing are shown in Table 4.5 and Figure 4.5.
Table 4.5 Payload acoustic environment
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Frequency 1/3 Oct Band SPL Frequency 1/3 Oct Band SPL （Hz） （dB） （Hz） （dB） 20 86.5 500 121.6
25 91.7 630 130.1
31.5 91.9 800 130.4
40 94.7 1000 130.9
50 101.3 1250 130.0
63 103.1 1600 127.8
80 102.6 2000 129.6
100 101.3 2500 126.9
125 104.5 3150 124.8
160 109.3 4000 122.6
200 109.7 5000 118.2
250 112.1 6300 116.5
315 119.0 8000 106.1
400 121.1 10000 100.4 Overall SPL 138.6 （dB） Note: 0 dB corresponds to 2×10-5Pa.
Figure 4.5 Acoustic environment
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4.3.5 Shock Environment
Shock environment during the flight is mainly presented at the significant events such as stage-3 separation, fairing separate and attitude controlling engine working. The maximum expected shock response spectrum at the payload interface is shown in Table 4.6 and Figure 4.. Table 4.6 Shock response spectrum
Frequency（Hz） SRS（g）
100 50
1000 3000
10000 3000
Note:Q=10
Figure 4.8 Shock response level
4.3.6 Thermal Environment
4.3.6.1 Ground Operations
The typical thermal environment of payload is kept around 20 ℃ ±5 ℃ (controlled in air-conditioned cabin).
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4.3.6.2 Flight Environment
1) The net flux density radiated by the fairing does not exceed 300 W/m2 before fairing separation.
2) The amount of free molecule heating does not exceed 1135 W/m2 after fairing separation.
4.3.7 Static Pressure Environment Under the Fairing During Ascent Phase
During the ascent phase, the static pressure evolution under the fairing is shown in Figure 4.6.
100000
80000
60000
40000 Pressure(Pa)
20000
0 0 20 40 60 80 100 120 140 160 180 Flight Time(s)
Figure 4.6 Static pressure under the fairing during ascent phase
4.3.8 Electromagnetic Environment
All the wireless device are mounted in interstage section and transition section. The parameters of the devices are listed in Table 4.8. The electromagnetic radiation intensity are listed in Table 4.9.
Table 4.7 Related parameters of wireless devices Transmitter Receptor Frequency Device Name Output Sensitivity Remarks （MHz） Power（W） （dBW） Safety Control Instruc- *** ≤-128 Upgoing tion Receptor
Telemetry Transmitter **** ≥7 Downgoing
Pulse Coherent Resp- Upgoing / **** ≥100 ≤-90 onder Beacon Downgoing
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Transmitter Receptor Frequency Device Name Output Sensitivity Remarks （MHz） Power（W） （dBW） GPS Receptor **** ≤-130 Downgoing
Remote Detection and Upgoing / Control Terminal Proc- **** ≥20 ≤-105 Downgoing essor
Table 4.8 Working electromagnetic radiation intensity of each device Work Frequency Section Radiation Intensity Device Name （MHz） dBuV/m Telemetry Transmitter **** 143
Pulse Coherent Responder **** 81 Beacon
Remote Detection and Control **** 153 Terminal Processor
4.3.9 Cleanliness Control
A clean environment superior to the cleanliness Class 100,000 is provided for the mating of satellite and launch vehicle.
The cleanliness of fairing is guaranteed by its airtight design during the short time of explosion to natural environment prior to launch.
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5 Payload Interfaces
5.1 Introduction
This chapter introduces satellite interfaces, including mechanical interfaces, electrical interfaces, and satellite design constraints.
5.2 Reference Axes
The launch vehicle coordinate system is OaXaYaZa (a coordinate) : a coordinate is connected to the launch vehicle; origin Oa locate in the mass center of vehicle; OaXa is along the vehicle longitudinal direction, OaYa is along the normal direction of IMU lens. It is shown in Figure 5.1.
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Figure 5.1 Coordinate system
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5.3 Dynamic Envelope of Fairing
Satellite is located in front fairing. Dynamic Envelope of Φ1.2m and Φ1.4m fairing is shown in
Figure 5.3 and Figure 5.3.
Figure 5.2 Dynamic envelope of Φ1.2m fairing
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Front Fairing
y=700*(x/1920)
Satellite
0.56
Satellite Docking Interface
Figure 5.3 Dynamic envelope of Φ1.4m fairing
5.4 Mechanical Interface for single satellite
Satellite could be connected to the launch vehicle through standard or nonstandard separator. The two main interfaces are separation bolt and clamp band with the help of disengaging springs.
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5.4.1 The mechanical interface of separation bolt and disengaging springs
This type of separator is shown in Figure 5.4. A docking ring is applied for connecting the satellite and the separator.
Figure 5.4 The outside view of the separator with separation bolt
The mechanical interface of the separator, the docking ring and the satellite is shown in Figure 5.5.
The installation position of the satellite could be fine-tuned owing to the docking ring.
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Figure 5.5 The interface view of the separator and the satellite
5.4.2 The mechanical interface of clamp band and disengaging springs
This type of separator is shown in Figure 5.46. The diameter in Φ300 mm is shown in Figure 5.7.
The diameter in 660 mm is only used in the front fairing of Φ1.4m motor. The interface is shown in Figure 5.8 and 5.9.
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Figure 5.6 The interface view of the separator in clamp band
Figure 5.7 The interface size of the separator inΦ300 mm
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Figure 5.8 The interface size of the separator inΦ660 mm (for satellite)
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Figure 5.9 The interface size of the separator inΦ660 mm (for vehicle)
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5.5 Mechanical Interface for multi satellites
The separator for multi satellites is usually custom-made. The standard 2U satellite and 10U satellite can choose standard installation interface. The separator for 15 satellites in one vehicle is shown in Figure 5.10. There are 6 standard 10U satellites, 8 standard 2U satellites and one main satellite. The separate interface of the main satellite is the separation bolt, which is shown in
Figure 5.5. The separate interface for standard 2U satellite is shown in Figure 5.11, which is compose of the separator frame, closed cabin door, compressed power spring subgroup and single spring hinge.
Figure 5.10 The separator for 15 satellites in one vehicle
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Figure 5.11 The separator for 2U satellite
5.6 Electrical Interfaces
KZ-1A could offer 14 channels for satellite test cooperating with ground launch systems.
Detachable plugs are installed on the interface between satellite and the launch vehicle and the rear stage of the launch vehicle. The test signal of the satellite could be sent from the last stage to the rear stage of the vehicle. Then the ground test equipment offer power and receive the data from the satellite. Once the vehicle is launched, the plug at the rear stage falls off. The other plug detach when the satellite is released.
The satellite test information could be transmitted from liquid propulsion & control system to the rear of the ground system with the help of an additional adapter between the satellite and the launch vehicle. The adapter between the launch vehicle and the ground system would be disconnected when the rocket is launched and the adapter between the satellite and the launch vehicle would be disconnected when the satellite is separating from the rocket.
The material of front fairing is electromagnetic wave transparent material. Satellite can be tested
Release 1.1 39 KZ-1A Solid Launch Vehicle User’s Manual wirelessly with front fairing. The launch vehicle provides an excellent electrical grounding process for satellite. The separation state of satellite can be estimated by state of pressure switch of satellite vehicle connection structure.
5.7 Satellite Design Constraint
5.7.1 Mass Properties Accuracy Constraint
The final mass properties statement shall specify the weight of the mass to an accuracy of 0.5%t, the center of mass to an accuracy to 5mm in each axis, and the products of inertia to 10 %.
5.7.2 Satellite Fundamental Frequency Requirement
The satellite should be designed to ensure that the fundamental frequency of the satellite, fixed at the payload interface, is greater than 22Hz.
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6 Satellite and Launch Vehicle Integration and Test
6.1 Introduction
This chapter introduces satellite and launch vehicle integration and test.
6.2 Facilities of JSLC
6.2.1 Satellite Testing Facilities
Jiuquan Satellite Launch Center (JSLC) provides the primary Satellite Processing Facility (SPF) for commercial KZ-1A launch vehicles.
a)General Assembly Test Hall
The testing hall is a 18m×36m×22.9(hight)m cleanroom with a 8m×15.5m door.
b) Environmental Condition
The temperature of the testing hall is maintained 20℃±5 ℃, the relative humidity is kept between
35% to 55%, and the cleanliness reaches class 100,000.
c) Lifting Capacity
A overhead crane which quality is 16/3.2t and height is 18m is located in the testing hall. The minimum lifting speed of main hook and auxiliary hook is not more than 1m/min and the connecting speed of them is 0.5m/min. The minimum horizontal speed of crane is not more than
2m/min.
d) Power Supply and Grounding
The testing hall could provide power which is 380V/220V±10, 50Hz±1Hz. And with a UPS power system, it has a 50kW capacity and could keep supplying power more than 15 minutes.
There are two socket boxes in the general assembly test hall. Every socket box contains a 380V,
25A four cores socket, two 220V, 16A three cores sockets and two double cores sockets. Besides, there are two 380V, 16A sockets and 8 to 10 220V, 16A sockets on the both side of hall. The testing hall and other testing facilities are connected by a 0.5m×0.5m(H)cable trench which is used
Release 1.1 41 KZ-1A Solid Launch Vehicle User’s Manual for communication between testing and dispatch.
The lightning protection level of the testing hall reaches the 1st class. Three independent grounding systems are connected with grounding lattice respectively. A,B and C refer to the grounding system of satellite, the grounding system of the AC stabilified-voltage power and the testing equipments’ shell and the grounding system of the testing equipments’ signal. The grounding resistance is not more than 1.
e)Gas Supply
There are three kinds of gas for supplying, i.e. compressed air, nitrogen and high purity nitrogen.
The compressed air could provide 23Mpa or 0.1 to 1.6Mpa pressure, the nitrogen could provide
23Mpa or 5Mpa pressure, and the high purity nitrogen could provide 23Mpa pressure.
The dew-point temperature of the compress air is not more than –55℃, the diameter of particles is not more than 20μm. The purity of the high purity nitrogen exceeds 99.998％（ V/V）,the dew-point is not more than –65℃, the total amount of CO,CO2 and CH4 is not more than 3×10-6（V/V）,the diameter of particles is not more than 10μm, and the oil content is not more than 0.3×10-6（V/V）.
6.2.2 Vehicle Integration Facility
The Vehicle Integration Facility (VIF) for all KZ-1A launches is on JSLC. KZ-1A uses this facility to horizontally integrate and test the launch vehicle components prior to their delivery to the launch site for final launch operations through a mobile-launch-platform. The facility for KZ-1A includes a testing hall and a control room.
Following the completion of launches tests, the fairing, liquid boost control segment and transition section are transferred to Satellite Testing Facility for satellite mating. And then the assembly is transferred back to the VIF for integration of satellite and launch vehicle.
6.2.3 Launch Site Facilities
Launch site facilities are mainly consist of the launch site and facilities of power supply and communication.
a) launch Site
The launch site is used for parking the mobile launch platform. The area of the site is 40m×15m, and the ground which is in 2m×2m region around the launching point keeps more than 5MPa
Release 1.1 42 KZ-1A Solid Launch Vehicle User’s Manual loading capacity.
b) Power and Grounding
Oil machines are used for supplying power, which could provide 380V/220V, 30kW power. Some grounding terminal boxes are installed in the launching site, and the grounding resistance is not more than 4Ω.
The lightning protection level of the launch site reaches the 1st class, and a lightning tower is built.
c) Communication
A communication interface box which contains 4 telephones is located in the launch site.
6.3 Launch Operations
KZ-1A ground and launch operations are conducted in four major phases:
Satellite status checking and test: the appearance, single item self-testing and status checking
of the satellite should be accomplished at SPF;
Launch vehicle status checking and test: the appearance, single item self-testing and status
checking of the satellite should be accomplished at VIF;
Vehicle, Payload and Mobile Launch Platform Integration: The satellite is assembled on the
front fairing and support plate at the PPF at first, and then is transferred to the vehicle testing
hall for docking with the vehicle. After being docked, the satellite and the vehicle are
transferred to a mobile launch platform for integration.
Launch: The mobile launch platform transfers the vehicle to the launch site. the task of
fueling of Stage Ⅳ should be finished the day before launch. Once the work in the launch
day, include transferring the launch platform, erection, final test of satellite and vehicle are
completed, the vehicle is fired.
Each of these phases is more fully described in the following sections. EXPACE maintains the responsibilities of launch site management and test scheduling throughout the entire integration cycle. Additionally, all integration activities are controlled by a comprehensive set of KZ-1A program files, which describe everything about integration of KZ-1A and its payload. Concretely
Release 1.1 43 KZ-1A Solid Launch Vehicle User’s Manual program files are also edited for special tasks or special operations.
6.3.1 Satellite Status Checking and Test
Satellite status checking and test is independent to the vehicle status checking and test and mainly accomplished by the satellite provider. The test date is also determined by the satellite provider.
The integration of launch vehicle and satellite requires 1 day. The flow chart of satellite status checking and test is shown as Figure 6.1.
X days before launch 6 days before launch Satellite The satellite Satellite is sent Satellite test checking and provider arrives to SPF finishes test
Figure 6.1 Flow chart of satellite status checking and test
6.3.2 Launch Vehicle Status Checking and Test
All the subsystems of KZ-1A are transmitted from the factory of EXPACE to the VIF and finish their appearance checking, single item self-test and status test. The main responsibility of the employee is as follows:
Check the appearance of all the subsystems ;
Carry through the single item test of some of the subsystems
Test the status of all the elements.
The function checking of all the subsystems are conducted through the vehicle status checking before they are assembled. The computer of the launch vehicle has specific software to every test.
The whole flight of the vehicle is simulated based on the true control program. All the actor, fire command and commands of flight computer are maneuvered. The launch vehicle would stay still until the test data is checked once a test is finished.
When the simulated flight is finished, fairing and liquid propulsion & control subsystem are sent to SPF for assembling. The flow chart of launch vehicle status checking and test is shown in
Figure 6.2.
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6 days before launch 5 days before launch Stage I, Stage II and Stage III stay in VIF
The launch Appearance checking and Cabin status vehicle is sent test of all the subsystems Overall examination and to VIF and the vehicle test
Fairing and the upper level are sent to VIF
Figure 6.2 Flow chart of launch vehicle status checking and test
6.3.3 Satellite, Vehicle and Mobile Launch Platform Integration
The joint of fairing and liquid propulsion & control subsystem, the checking of interface and the encapsulation of the satellite are accomplished at SPF. Once the satellite, fairing and liquid propulsion & control subsystem are assembled together, they are transmitted to VIF. The encapsulated satellite is lifted and rotated to a proper direction to connect to the launch vehicle.
Integrated test is carry out after all the electrical interface is connected. All the interface, including the interface between the satellite and the vehicle, are checked. If the launch vehicle passes the test, all the assembly jobs are finished.
The launch vehicle would be transmitted to the mobile launch platform after assembling. The flow chart of satellite, vehicle and mobile Launch Platform Integration is shown as Figure 6.3.
5 days before launch
Fairing and the upper level are sent Interface to SPF connection and Assemble the Encapsulated test of the satellite and the satellite is sent to satellite and launch vehicle VIF launch vehicle Satellite tests finish. Electrical System test and connection of Interface checking the satellite and launch vehicle 4 days before launch 4 days before launch 3 days before launch
Final checking and The whole launch The whole checking and test Assembling test on the launch vehicle is sent to the of the subsystems vehicle launch platform
Figure 6.3 Flow chart of satellite, vehicle and mobile launch platform integration
6.3.4 Launch
A full flow of maneuver should be taken before launch. If the result can’t meet the requirements, the maneuver should be taken a second time. The aim of maneuver is familiar with the
Release 1.1 45 KZ-1A Solid Launch Vehicle User’s Manual communication system, launch flow, operation command flow and emergency plan. the work of fueling the Stage Ⅳ should be finished one day before launch. Then the count-down timer starts. Table 6.1 lists the key KZ-1A procedure during the final countdown on the Launch
Day. Table 6.1 Typical Pre-Launch Countdown Procedure
Time Procedure
-240min Prepare for launch; launch vehicle erection
-200min The whole checking of the launch vehicle
Satellite test (The test is allocated 1 hours, but the time could be -190min adjusted based on the actual situation)
-130min Install lattice rudders
-95min Aiming
-70min Correction of guidance before launch
-25min Communication & Tracking System power-on
-15min Recall satellite with GPS/BD2
-15min Demolition of defend breeze blot
-10min Communication & Tracking System starts work
-9min30sec Start the flow of launch
-3min Control System power switch-over
Automatically ignition procedure start;steering engine、safety control -1min turning to power of launch vehicle
0sec Ignition
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7 Mission Integration and Management
7.1 Introduction
This Chapter outlines the management process applied to contract implementation, the customer responsibilities, the interface requirements and documentation required.
7.2 Management Overview
7.2.1 SC/LV Interface Control
The SC/LV interface is controlled by the Interface Control Document (ICD) based on the template provided by EXPACE and details all of the customer interfaces, the customer’s technical, documentation and the schedule requirements.
7.2.2 Launch Services Schedule
The launch services schedule should be in line with the SC mission process, the typical program schedule is shown in Figure 7.1 and Figure 7.2. LV Manufacture LV AIT LV Activities
LV Lick-Off PMAR TIM#1 FMAR Site Survey
1st Site Survey* PMA Cycle FMA Cycle
LSC Launch EDC Launch Campaign Preliminary Critical Design Design
Vibration, SC Kick-Off PDR CDR Acoustic, Fit-check& Separation Shock Test
SC Activities SC Manufacture SC AIT
* for SC manufacturers that first work with ‘Feitian 1’ launch vehicle Figure 7.1 Typical launch services program schedule
The typical schedule is based on a 12 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 12 month
Release 1.1 47 KZ-1A Solid Launch Vehicle User’s Manual schedule, which is shown in Figure 7.2. The schedule for the specific launch is developed in conjunction with the customer and satellite manufacturer based on the satellite schedule.
The typical schedule can be shorter if requested and supported by the customer.
MONTHS -24 -23 -22 -21 -20 -19 -18 -17 -16 -15 -14 -13 -12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0
Effect Day of Launch ■ Service Contract(EDC) Mission Integration User´s Manual ▲ Safety Document Submission ▲ ▼ ▲ ▼ ▲ ▲ ▼ ▲ Launch Services Application ▼ Interface Control Document(ICD) ▼ ▲ ▲ ▼ ▲ ICD Verification ▲ ▲ Mission Analysis Review ■ ■ Launch Site Readiness Review ▼ LV Pre-shipment Review(PSR) ■ Launch Readiness Review ■ PMAR ▲ FMAR ▲ Mission Analysis Trajectory Analysis ▲ ▲ Separation Analysis ▲ ▲ RF/EMC Compatibility ▲ ▲ Mechanical Interface ▲ ▲ Electrical Interface ▲ ▲ Coupled Loads Analysis ▲ ▼ ▲ ▼ ▼ ▲ Thermal Analysis ▲ ▼ ▲ ▼ ▼ ▲ Rocket Motor Production Long Lead Procurement ▲ ▲ Production ▲ ▲ Buy-Off , Transport and Test ▲ ▲ Vehicle Production Avionics Components ▲ ▲ Harness Manufacturing ▲ ▲ Interstage Production ▲ ▲ Fairing Production ▲ ▲ Fairing Test ▲ ▲ Fairing Final Assembly ▲——LV Side ▲ ▲ Interstage Final Assembly ▲ ▲ ▼——User Avionics Testing ▲ ▲ —— Carry to JSLC ■ Both Side ▲
-24 -23 -22 -21 -20 -19 -18 -17 -16 -15 -14 -13 -12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 MONTHS
Figure 7.2 The typical schedule
7.2.3 Launch Campaign Schedule
The typical Launch Campaign period for KZ-1A launch vehicle in JSLC is 7 days. The activities assigned to each day are described in Paragraph 6.3 and 7.10.1.
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7.2.4 Launch License and Permits
EXPACE shall be responsible for ensuring that all the requisite Chinese government permits and approvals for the launch services and the launch of the customer’s satellite have been obtained and are in place.
7.2.5 Satellite Technology Safeguard
EXPACE commits to meet customer’s requirement to ensure the satellite technical security during the whole process of contract implementation.
EXPACE 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 Jiuquan for the satellite and its support equipment.
When the SC arrives at JSLC, the satellite processing facility is handed over to and will be exclusively controlled by the customer during the whole launch campaign.
7.3 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 Satellite Fairing (STF), Satellite Adapter (STA), 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.
7.4 Interface Control Document (ICD)
The ICD is based on the standard ICD developed by EXPACE 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 7.9 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
Release 1.1 49 KZ-1A Solid Launch Vehicle User’s Manual requirements, and the mission parameters.
The ICD is developed jointly by the customer and EXPACE 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 Analysis
Review (FMAR). All revisions are jointly agreed and signed by the customer and EXPACE.
The ICD is the master document for the launch services program and takes precedence over all other technical documents.
7.5 Mission Analysis
7.5.1 Introduction
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 launch services are based on a standard mission profile as defined in the User’s 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 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.
7.5.2 Preliminary Mission Analysis
The preliminary mission analysis is based on the initial injection requirements according to the
Interface Requirement Document (IRD, see Paragraph 7.11) 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
Release 1.1 50 KZ-1A Solid Launch Vehicle User’s Manual 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.
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.
7.5.2.1 Trajectory Analysis
The following topics should been completed after the preliminary mission analysis has been confirmed. 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
7.5.2.2 Separation Analysis
The following topics should been completed for Separation Analysis after the preliminary mission analysis has been confirmed. a) The satellite attitude at SC/LV separation; b) Analysis of the SC/LV separation velocity;
Release 1.1 51 KZ-1A Solid Launch Vehicle User’s Manual c) Analysis of the collision avoidance maneuver; d) Analysis of the separation dynamics, including impact of fuel slosh.
7.5.2.3 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 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.
7.5.2.4 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 EXPACE and JSLC to verify that the values provided in the ICD are correct.
7.5.2.5 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.
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7.5.2.6 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.
7.5.2.7 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.
7.5.2.8 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.
7.5.3 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
Release 1.1 53 KZ-1A Solid Launch Vehicle User’s Manual 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.
7.5.3.1 Trajectory Analysis
The following topics should been completed after the final mission analysis has been confirmed. 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 the contracted lift-off mass e) Propellant reservation f) Predicted injection accuracy
7.5.3.2 Separation Analysis
The following topics should been completed for Separation Analysis after the final mission analysis has been confirmed. a) The satellite attitude at SC/LV separation b) Analysis of SC/LV separation velocity c) Analysis of the collision avoidance maneuver d) Analysis of the separation dynamics, including impact of fuel slosh.
7.5.3.3 Coupled Loads Analysis (CLA)
The final CLA should complete the following topics, using the final satellite dynamic model provided by the customer: a) Modal analysis of the satellite and launch vehicle
Release 1.1 54 KZ-1A Solid Launch Vehicle User’s Manual 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 EXPACE 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.
7.5.3.4 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 platform. The ICD is used as the baseline for this analysis and allows the EXPACE and JSLC to verify that the values provided in the ICD are correct.
7.5.3.5 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.
7.5.3.6 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.
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7.5.3.7 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.
7.6 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 EXPACE review and approval.
7.6.1 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 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.
7.6.2 Satellite Environmental Test Program
A ST/STA fit check shall be performed to verify the interface and may be combined with a separation shock test.
EXPACE 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.
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7.6.3 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.
7.7 Meetings and Reviews
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 7.1 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 one month before the start of Launch Vehicle Pre-shipment Review (PSR) launch campaign Combined Operations Procedure Review During the launch campaign Launch Readiness Review (LRR) 2 days before launch
7.7.1 Kick-Off Meeting (KOM)
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
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The first issue of the ICD will be jointly drafted by EXPACE and customer after this KOM.
7.7.2 Phase 1/2/3 Safety Submission
The safety submission is required to show that the satellite design and customer operations at the launch site meet all the safety requirements of JSLC and EXPACE. The safety submission shall contain a description of all the hazardous systems, testing and materials the customer will utilize while processing the satellite at JSLC.
7.7.3 Initial 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.
7.7.4 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 EXPACE and the customer; however, TIM do not replace the scheduled technical meetings and are only held when a specific issue requires discussion between scheduled reviews.
7.7.5 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.
7.7.6 Launch Site Survey
The review is held in JSLC around six months before launch. The satellite team will be invited to this survey to verify that the facilities and equipment of JSLC on-site, such as the MCCC, BS, clinic facility and launch platform, 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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7.7.7 Launch Vehicle Pre-shipment Review (PSR)
This review will be held at least one month before launch. The purpose of this review is to confirm that the launch vehicle meets the specific requirements in the process of launch vehicle design, integration and testing. The launch vehicle delivery date to JSLC will be discussed in the meeting. EXPACE 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.
7.7.8 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.
7.7.9 Combined Operations Procedure Review
This review will be held at JSLC following the submission of the Combined Operation Procedure prepared by JSLC.
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.
7.7.10 Launch Readiness Review (LRR)
The LRR will be held at JSLC after the launch rehearsal. The review will cover the status of satellite, launch vehicle, launch facilities and TT&C network.
7.8 Launch Reports after SC/LV Separation
EXPACE 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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7.8.1 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.
7.8.2 Orbit Tracking Report EXPACE also requires the customer confirming acquisition of the satellite signal and the orbit parameters as soon as possible after SC/LV separation.
7.8.3 Launch Mission Evaluation Report
EXPACE 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 accuracy, the separation attitude and roll rates, and the comparison with the predictions.
7.9 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.
7.9.1 Documentation
The following documents shall be provided to EXPACE by customer: 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 g) Satellite Mechanical Environment Test Results h) Safety Submissions i) Satellite Mass Properties
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Environment Test Plan)
7.9.2 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
The customer shall also provide the following umbilical connector halves to EXPACE: Two pairs of connectors halves for flight (1 nominal + 1 spare)
7.10 Launch Campaign
7.10.1 Introduction
The launch campaign starts with the arrival of the customer’s advance team in JSLC. 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 JSLC 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 JSLC with the customer’s support for interface verification. The final preparation and launch are also managed by JSLC with customer’s support.
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The typical launch vehicle working period (from the launch vehicle to the launch site to ignition) for the launch campaign is 7 days. The typical launch campaign flowchart (assuming the launch campaign period for satellite is 26 days) is listed in Figure 7.3.
-X -6 -5 -4 -3 -2 -1 0
SC advance team arrival ▲ Launch facilities check and acceptance review ▲ SC arrival, transported to satellite test factory ▲ SC statue check test ▲ LV arrival, transfer to VIF ▲ LV statue check test, LV transfer to SC workshop, LV and SC mating. ▲ LV/SC cabin statue check test, final assembly. ▲ LV/SC transfer to mobile launch platform ▲ Launch rehearsals ▲ LV fueling ▲ Transfer and Luanch ▲
-6 -5 -4 -3 -2 -1 0
Figure 7.3 Typical launch campaign flowchart
7.10.2 Launch Campaign Preparation
In support of the customer preparation for the launch campaign, EXPACE and JSLC will issue a series of documents covering the activities and facilities. The key documents are as follows: a) JSLC User’s Manual b) Combined Operations Plan c) Combined Operations Procedures d) Countdown Procedure e) Go/No Go Criteria f) Launch Rehearsal Procedures
EXPACE will organize the launch site acceptance review at JSLC 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.
7.10.3 Launch Campaign Documentation
This section outlines the documents to be prepared by the customer and EXPACE that are required to plan and implement the launch campaign.
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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
7.11 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 7.2. Table 7.2 Launch services documentation submission date S.N. Documents Provider Due Date Launch Vehicle’s Introductory Technical Documents Launch Vehicle User’s Manual JSLC User’s Manual 1 month 1 EXPACE Launch Vehicle Safety Requirement Documents after EDC Required Format of Spacecraft Dynamic Model Required Format of Spacecraft Thermal Model Interface Requirement Document (IRD) The customer shall prepare the IRD, which generally includes the following information: General mission technical requirements 2 months 2 Launch Safety and Security Requirements Customer after EDC 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. 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. EXPACE will perform the 2 months 3 dynamic Coupled Load Analysis (CLA) with the Customer after No.1 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.
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S.N. Documents Provider Due Date Dynamic Coupled Load Analysis (Preliminary and Final) EXPACE will integrate the SC model, launch vehicle model and flight characteristics together to calculate 3 months 4 loads on SC/LV interface at selected critical points. The EXPACE after No.3 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. Spacecraft Thermal Model (Preliminary and Final) The customer shall provide hard copies or soft copies(such as a CD)of the spacecraft thermal model 2 months 5 according to Required Format of Spacecraft Thermal Customer after No.1 Model. EXPACE will use the model for thermal environment analysis. The analysis output requirement should be specified in written form. Thermal Analysis (Preliminary and Final) This analysis determines the spacecraft thermal 3 months 6 EXPACE environment from the arrival of the spacecraft to its after No.5 separation from the launch vehicle. 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 3 months 7 interfaces, electrical interfaces and RF characteristics; Customer after EDC c) Requirements and constraints for combined operations. The customer shall provide the spacecraft configuration drawings to EXPACE. For all minimal or potential protrusion out of the fairing envelope, agreement has to be reached with EXPACE on its acceptability one year before launch.
S.N. Documents Provider Due Date Safety Phase 1 – Design This submission is for the design of the spacecraft and Before the 8 ground support equipment with details of all hazardous Customer PMAR systems or components. It will provide details of all safety measures implemented, warning devices included
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and a list clearly showing the risks for each hazardous system and component. Mission Analysis Reports (Preliminary and Final) a) Trajectory Analysis To optimize the launch mission; To determine the launch sequence, flight trajectory and performance margin. b) 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 3 months 9 separation; EXPACE after No.7 To review if the LV collision avoidance maneuvers can satisfy the spacecraft requirements. c) Interface Compatibility Analysis (EMC Analysis, Electrical Interface Analysis and Mechanical Interface Analysis) To ensure that the SC/LV interfaces are compatible. d) 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. Safety Phase 2 – AIT and Qualification Status This submission provides the qualification 12 months status,manufacturing status and acceptance verification of Customer 10 before all hazardous systems and components for the spacecraft launch and ground support equipment. This submission will also include the initial spacecraft operations procedures. S.N. Documents Provider Due Date 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 One month translated into a qualification and acceptance test 11 Customer before the program. The plan shall provide an overview of the test 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
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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. Spacecraft Environmental Test Report The reports provided by the customer should include the detailed environmental test and some related analysis 15 days Customer 12 conclusions. The adaptability and the margins of the after the
spacecraft should also be included. The document will be test jointly reviewed by both sides to ensure the compatibility of SC and LV. Safety Phase 3 – Final Acceptance and Hazardous Operations The final submission includes the results of 6 months the AIT program for the hazardous systems plus the final 13 Customer before details of all hazardous operations to be conducted at the launch launch site. This includes the spacecraft, ground support equipment and deliveries of hazardous materials.
S.N. Documents Provider Due Date 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 Jiuquan airport through to mating to the Payload Adapter (PLA). The plan will also include all the JSLC interface requirements, the 8 months 14 support required for the standalone spacecraft operations Customer before and communications requirements. Although the plan is launch 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 JSLC. Both sides will jointly review this document. Part of the document will be incorporated into ICD and written into SC/LV Combined Operation Procedure. Spacecraft Operation Procedure These procedures will be used during the 2 months launch campaign by the customer in processing the 15 Customer before spacecraft. The hazardous procedures shall detail every launch step required including the constraints, safety requirements, handling requirements and personnel
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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 JSLC interfaces for communications and monitoring the spacecraft.
S.N. Documents Provider Due Date SC/LV Combined Operations Plan JSLC 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 8 months spacecraft with specific emphasis on the following: 16 EXPACE before a) SC/LV joint operations launch 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. SC/LV Combined Operations Procedure JSLC 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, 2 months handling requirements, personnel requirements and the Both 17 before responsibilities. These procedures are to ensure that the Sides launch 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. Approval of the spacecraft and ground support Before SC 18 Customer equipment for fuelling and combined operations. fuelling
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S.N. Documents Provider Due Date SC’s Mass Property Report (Final) The spacecraft's mass properties are measured and After the SC 19 Customer calculated after all tests completed. The data shall be is fueled provided to EXPACE one day before SC/LV integration. Launch Rehearsal Procedure 15 days Both 20 This document specifies the detailed operation steps before Sides during the launch rehearsal. launch GO/NO GO Criteria This document specifies the baseline criteria for the GO/NO-GO and the orders issued by the relevant 15 days Program Managers of the mission team. The Go/No Go Both 21 before criteria for the launch vehicle are standard but the criteria Sides launch will be updated with the spacecraft requirements. The operation steps have been specified inside SC/LV Combined Operation Procedure. Countdown Procedure The countdown procedure details the countdown sequence, the communication links, all communication exchanges, e.g. commands, status reports, test values,etc., 15 days Both 22 as conducted on the launch day. It also details the before Sides responsibilities of the various participants and the inputs launch required from the customer. The contingency procedures will also be included for unplanned launch holds and launch abort. Orbit Injection Data Report The initial orbit injection data of the spacecraft will be provided 30 minutes after SC/LV separation. This 30minutes 23 document will either be handed to the customer's EXPACE After SC/LV representative in JSLC or sent via telex, facsimile or separation e-mail to a destination selected by the customer. Both sides will sign on this document.
S.N. Documents Provider Due Date Orbital Tracking Report The customer is required to provide the first round of 20 days 24 Customer orbital tracking data after SC separation. This data is used after launch to verify the launch vehicle performance. Launch Mission Evaluation Report Using injection parameters and telemetry data obtained 60 days 25 EXPACE from the launch vehicle, the launch vehicle side will after launch provide an assessment of the launch vehicle's
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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.
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APPENDIX A
The launch vehicle party requires the satellite party to fill in an effective payload investigation chart in the initial mission analysis. This effective payload investigation chart is the initial document in the mission cycle which should be well prepared 11 months before the required launch date. PAYLOAD QUESTIONNAIRE
Payload Information
Full Name:
Acronym:
Owner/Operator:
Integrator(s):
Points of Contact and Contact Information：
Payload Mission Information
Desired Launch Date/Timeframe:
Launch Window Constraints:
Mission Timeline Description:
Payload Trajectory Requirements
Parameter Value SI Units
Desired Semi-Major Axis ______±______km
Desired Eccentricity: ______≤e≤______
Desired Orbit Inclination ______±______deg
Desired Right Ascension of Ascending Node ______±______deg
Desired Argument of Perigee ______±______deg
Payload Orbital Injection Conditions
Parameter Value SI Units
Maximum Allowable Tip-Off Rate deg/s
Desired Spin-Up Rate rpm
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Pointing Requirement
(Please Specify)
Maximum Allowable Pointing Error deg
Describe Any Other Separation Requirements:
Other
Payload Mass Properties
Parameter Value SI Units
Spacecraft Mass (Maximum) kg Describe the Origin and Orientation of the spacecraft reference coordinate system, including its orientation with respect to the launch vehicle (provide illustration if available):
Spacecraft Coordinate System
Item Stowed Configuration Tolerance
Xcg ± Center of Gravity Ycg ± (mm) Zcg ±
Moment of Inertia Ixx ±
2 (kg. mm ) Iyy ±
Izz ±
Product of Inertia Ixy ±
2 (kg. mm ) Iyz ±
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Ixz ±
Payload Mechanical Interface
Parameter Value SI Units
Spacecraft Length/Height (Maximum) mm
Spacecraft Diameter (Maximum) mm
Describe any appendages/antennas/etc which extend beyond the basic spacecraft envelope:
*Note: If available, provide dimensioned drawings for stowed configurations.
Do you have a Spacecraft Separation System? If so, provide details here:
Note: KZ-1A can design/provide the Spacecraft Separation System if desired.
Payload Electrical Interface
Bonding Requirements:
Are Launch Vehicles Supplied? Yes / No If Yes, describe:
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Discrete Commands Required? Yes / No
Is Electrical Access to the Satellite Required... After Encapsulation? Yes / No
Is Satellite Battery Charging Required... After Encapsulation? Yes / No
Is a Telemetry Interface with the Launch Vehicle Flight Computer Required?
Yes / No
If Yes, describe:
Other Electrical Requirements:
*Note: Please complete attached sheet of required pass-through signals.
Payload Thermal Environment
Parameter Value SI Units
Pre-launch Temperature Range °C
grains/lb Pre-launch Allowable Water Vapor in Air dry air
Maximum Ascent Heat Flux W/m2
Maximum Free-Molecular Heat Flux W/m2
Maximum Fairing Ascent Depressurization Rate mbar/s
Payload Contamination Control
Parameter Value SI Units
Desired Payload Processing Capabilities Class
Desired Fairing Air Cleanliness Class
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Payload Dynamic Environment
Parameter Value SI Units
Maximum Allowable Acoustic Sound Pressure Level dB OA
Maximum Allowable Sine Vibration Grms
Maximum Allowable Shock g
Maximum Lateral Acceleration g
Maximum Axial Acceleration g
Fundamental Frequency - Lateral Hz
Fundamental Frequency - Longitudinal Hz
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Additional Data: 1. Please provide a description of the payload testing planned during payload processing at the launch site, as well as any testing planned while encapsulated. Please describe each test in terms of personnel required, duration of test, tools/GSE required, and any possible safety concerns that should be considered.
2. Please describe any safety issues associated with the spacecraft.
3. Please describe the propulsion systems to be used on the spacecraft.
4. Please describe the pressure vessels to be used on the spacecraft.
5. Please describe the power systems (batteries, solar cells, etc).
6. Please describe the RF systems to be used on the spacecraft. Please detail each RF transmitter or receiver, its function, frequency, sensitivity, power output, and bandwidth.
7. Please provide the spacecraft allowable or test acoustic profile, random vibration spectrum, shock spectrum, and sine vibration curve.
8. Please provide Dimensional Drawings and/or CAD models of the spacecraft if available.These drawings/models should include the spacecraft separation system. Rather than attaching to this PDF.
9. Please describe any security concerns or requirements you have.
10. Please describe any additional spacecraft requirements that we should be made aware of.
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