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GJB 9632-2019   Specification for low orbit spacecraft high-voltage power supply system (English Version)
Standard No.: GJB 9632-2019 Status:valid remind me the status change

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Word Count: 11000 words Price(USD):400.0 remind me the price change

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Implemented on:2020-1-1 Delivery: via email in 1 business day
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Standard No.: GJB 9632-2019
English Name: Specification for low orbit spacecraft high-voltage power supply system
Chinese Name: 低轨航天器高压母线电源系统规范
Professional Classification: GJB    Professional Standard - Military
Issued on: 2019-12-08
Implemented on: 2020-1-1
Status: valid
Language: English
File Format: PDF
Word Count: 11000 words
Price(USD): 400.0
Delivery: via email in 1 business day
Codeofchina.com is in charge of this English translation. In case of any doubt about the English translation, the Chinese original shall be considered authoritative. This specification was proposed by the Equipment Department of Aerospace Systems Department of PLA Strategic Support Force. Specification for low orbit spacecraft high-voltage power supply system 1 Scope This specification specifies the technical requirements, quality assurance regulations, delivery preparation and instructions of 100V high-voltage power supply system and its products for low orbit manned spacecraft. This specification is applicable to the development, test and acceptance of high-voltage fully regulated bus power supply system (hereinafter referred to as power supply system) with bus voltage of 100V for low orbit manned spacecraft. 2 Normative references The following normative documents contain provisions which, through reference in this text, constitute provisions of this specification. For dated or version indicated reference, subsequent amendments (excluding corrections), or revisions, of any of these publications do not apply to this specification. However, parties to agreements based on this specification are encouraged to investigate the possibility of applying the most recent editions of the normative documents indicated below. For undated references or references with version not indicated, the latest edition of the normative document referred to applies. GB/T 191 Packaging — Pictorial Marking for Handling of Goods GJB 150.5A-2009 Laboratory environmental test methods for military materiel — Part 5: Temperature shock test GJB 150.15A-2009 Laboratory environmental test methods for military materiel — Part 15: Acceleration test GJB 150.16A-2009 Laboratory environmental test methods for military materiel — Part 16: Vibration test GJB 150.17A-2009 Laboratory environmental test methods for military materiel — Part 17: Acoustic noise test GJB 150.18A-2009 Laboratory environmental test methods for military materiel — Part 18: Impact test GJB 151A-1997 Electromagnetic emission and susceptibility requirements for military equipment and subsystems GJB 152A Electromagnetic emission and susceptibility measurements for military equipment and subsystems GJB 1181 Packaging, handling, storage and transportability program requirements (for systems and equipment) GJB 2042A-2012 General specification for electrical power system of satellite GJB 2602-1996 General specification for space solar cell arrays GJB 2831A-2009 General specification for hermetically sealed nickel-hydrogen rechargeable cells in spacecraft GJB 2998 Mark of satellite products GJB 4038-2000 General specification for solar cell array mechanisms GJB 5174-2003 General specification for solar array drive assembly of tracking the sun GJB 6789-2009 General specification for lithium-ion rechargeable cells in spacecraft GJB/Z 35-1993 Derating criteria for electrical, electronic and electromechanical parts GJB/Z 1391 Guide to failure mode, effects and criticality analysis QJ 1019 Measurement method for electrical characteristics of solar cells QJ 2630.1 Space environment test methods for satellite components — Thermal vacuum test QJ 2630.3 Space environment test methods for satellite components — Vacuum discharge test 3 Requirements 3.1 Composition The composition of power supply system shall include: a) power generator: semi-rigid or flexible solar cell array; b) energy storage device: nickel-hydrogen battery pack or lithium-ion battery pack; c) sun-tracking device: including driving mechanism and actuator; d) power supply control device: including control equipment to realize the functions of main error amplification signal, shunt regulation, discharge regulation, charging control and bus filtering. 3.2 Performance requirements 3.2.1 Power bus 3.2.1.1 Bus voltage Nominal bus voltage: 100V. 3.2.1.2 Adjustment degree of bus voltage In regulation domain, the bus voltage regulation degree is within ±3%. 3.2.1.3 Bus voltage jump rate when entering in or exiting from shadow The voltage jump rate of bus shall be no more than 3.5V/ms and the voltage disturbance caused by bus shall be no more than 10%. 3.2.1.4 Output impedance of bus voltage Output impedance of bus shall be no more than 70 mΩ. 3.2.1.5 Bus ripple voltage In the frequency range of 10 kHz ~ 10 MHz and under rated output voltage and rated resistive load, the peak-to-peak ripple voltage of power supply system shall be no more than 500mV(p-p). 3.2.1.6 Transient characteristics of bus voltage When the transient of load current is less than 50% of the rated load, the bus output voltage shall not exceed 5% of the nominal bus voltage, and the adjustment time to dropping the voltage to within 1% of the nominal voltage shall not exceed 20ms. When the transient of load current exceeds 50% of the rated load, the bus output voltage shall not exceed 10% of the nominal bus voltage, and the adjustment time to dropping the voltage to within 1% of the nominal voltage shall not exceed 20ms. 3.2.1.7 Surge current The surge current provided by the bus shall meet the requirements of the maximum surge current when multiple loads are powered at the same time. The surge current rising slope is generally not less than 1×106A/s, and its duration is not less than 5ms, and the surge amplitude is not more than 50% of the rated load current of the bus. Under the above surge current, the voltage disturbance caused by bus shall not exceed 5%. 3.2.1.8 Power margin The output power margin of the power bus shall be at 5% ~ 10%. 3.2.2 Power supply of initiating explosive device Battery pack tap method is generally adopts for large pulse current power supply device like initiating explosive devices, generally adopts cell tap mode, and can meet the voltage requirements of initiating explosive devices under specified working temperature conditions. 3.2.3 Power generator 3.2.3.1 Output power The output power of solar cell array is determined by loading power, charging power, transmission loss and power margin of the spacecraft. Meanwhile, factors such as system test error, combination loss, optimal operating point voltage and current, effective illumination and on-orbit operating temperature shall be considered. The output power of end-of-life solar cell array is determined by the maximum output power at its early life, the working temperature of orbital solar cell array, effective illumination, particle radiation, attenuation coefficient of ultraviolet radiation and loss at high-low temperature cycle, etc., which shall meet the requirements of the provisions of special technical documents. 3.2.3.2 Output voltage The output voltage of end-of-life solar cell array is determined by bus voltage, battery pack charging voltage, isolation diode of power supply circuit, sampling resistance, cable voltage drop and other factors. The determination shall be leave with a certain margin. 3.2.3.3 Mechanical characteristics Mechanical characteristics of solar cell array are as follows: a) folded state of solar cell array: 1) the requirements of strength and stiffness under the launching mechanics environment shall be met, the solar cell array shall be free from relative sliding, and the solar panels shall not collide with each other; 2) the first-order natural frequency of the solar cell array in the folded state shall not be coupled with the natural frequency of the spacecraft, and shall meet the requirements of special technical documents of the spacecraft. b) unfolded state of solar cell array: 1) the impact load caused by the unfolding and locked of the solar cell array shall not exceed the impact allowable range of the solar cell array driving device and other components; 2) the first-order bending natural frequency after the solar cell array is unfolding and locked shall not be coupled with the loop frequency of the spacecraft control system; 3) the first-order torsion natural frequency after the solar cell array is unfolding and locked shall not be coupled with the driving frequency of the solar cell array driving device. 3.2.3.4 Thermal characteristics The solar cell array shall be able to withstand the influence of high-low temperature cycle in orbit. The temperature of solar cells shall be no higher than 110°C in illumination area and no lower than -100°C in shadow area. 3.2.3.5 Cell covering rate The requirements for the cell covering rate of solar cell array are as follows: a) the cell covering rate of the unfolding semi-rigid solar cell array is not less than 85%; b) the cell covering rate of flexible solar cell array shall meet the requirements of special technical documents for spacecraft. 3.2.3.6 Anti-blocking For the solar cell array circuit, protective measures shall be taken to prevent hot spots caused by local blocking, and the influence of blocking on the output power of the solar cell array shall be analyzed. 3.2.3.7 Remnant magnetic torques The remnant magnetic torques produced by the solar cell array shall not be greater than 0.4A·m2. 3.2.3.8 Anti-space-radiation and anti-atomic-oxygen The requirements of solar cell array for anti-space-radiation and anti-atomic-oxygen are as follows: a) the total dose equivalent to 1 MeV electron damage during the on-orbit life of the solar cell shall be determined. The ratio of the failure dose to the total dose of electron damage during the on-orbit life shall be no less than 2 times; b) protective measures shall be taken against ultraviolet rays, and the power loss caused by ultraviolet irradiation shall be considered; c) protective measures shall be taken against the corrosion caused by atomic oxygen for materials like polyimide film. 3.2.3.9 Anti-static Anti-static requirements of solar cell array are as follows: a) according to the requirements of 3.7.2, the solar cell array structure shall be grounded with high resistance; b) if the potential difference between adjacent circuits of the solar cell string is not more than 70V, the interval shall be not less than 2mm. 3.2.4 Energy storage device 3.2.4.1 Capacity The actual capacity of the battery pack at the beginning of its life shall be 110% of the rated capacity. The capacity deviation of cell in each battery pack shall not exceed 3% of the rated capacity. 3.2.4.2 Working voltage and charging efficiency 3.2.4.2.1 Working voltage of battery pack Under the worst working conditions (end-of-life, open circuit failure of a cell or a single parallel unit, maximum discharge depth and longest shadow period), the minimum discharge voltage of the battery pack after deducting the line voltage drop shall not be less than the minimum value required by the bus voltage or discharge regulator. Under the worst working conditions (end-of-life, open circuit failure of a cell or a single parallel unit), the highest charging voltage of the battery pack after adding the line voltage drop shall not be greater than the voltage of best working point of the solar cell array or the highest value required by the charging regulator. 3.2.4.2.2 Discharge voltage, charging voltage and charging efficiency of cell The average discharge voltage, charging voltage and ampere-hour charging efficiency of cell on-orbit operation for 3a shall meet the requirements in Table 1. If the cell has been in orbit for more than 3a, it is necessary to analyze and determine the average discharge voltage, charging voltage and charging efficiency according to the ground test data.   Table 1 Average discharge voltage, maximum end-of-charge voltage and charging efficiency of battery pack (on-track 3a) Item Nickel-hydrogen cell Lithium-ion cell Average discharge voltage 1.25V 3.6Va end-of-charge voltage ≤1.65V ≤4.2V Ampere-hour charging efficiency ≥90% ≥95% Average discharge depth ≤30% ≤20% a for lithium cobalt oxides system cell. 3.2.4.3 Temperature gradient The maximum temperature difference between cells in the same cell module shall not exceed 3°C, and the maximum temperature difference between modules in the same pack shall not exceed 5°C. 3.2.4.4 Redundant backup Redundancy is conducted by hot backup with one or two cells. The battery pack shall be able to work if a cell is failed, and still able to work normally if a cell is short-circuited or open-circuited. 3.2.4.5 Anti-open-circuit When a cell in the battery pack is open-circuited, the battery pack shall be able to charge and discharge normally. The requirements for different battery packs are as follows: a) for nickel-hydrogen battery pack, the charging and discharging channels of the battery packs shall be provided when the open circuit fails; b) for lithium-ion battery pack, single parallel connection mode or anti-open-circuit with bypass device may be adopted. 3.2.4.6 Discharge depth, working temperature and cycle life The discharge depth, working temperature and cycle life of the battery pack in orbit for 3a shall meet the requirements in Table 2: Table 2 Average discharge depth, optimum working temperature and cycle life Item Nickel-hydrogen cell Lithium-ion cell Average discharge depth ≤30% ≤20% Optimum working temperature 0°C~10°C 10°C~30°C Cycle life 18000 times 18000 times For the cell in orbit for more than 3a, it is necessary to analyze and determine the discharge depth and working temperature according to the ground life test data. 3.2.4.7 Equalization processing For the power supply system using lithium-ion battery pack, the voltage of each cell or parallel single unit in the battery pack shall be controlled within 60mV deviation by means of equalization processor. 3.2.5 Sun-tracking device 3.2.5.1 Electrical property The electrical property of the drive mechanism include: a) the electric transmission capacity (current and voltage) of the conductive ring; b) the transmission voltage drops of conductive ring; c) transmission noise. The electrical performance of the driving mechanism shall meet the requirements of special technical documents for spacecraft. 3.2.5.2 Driving ability 3.2.5.2.1 Drive work mode According to the driving pulse input by the driver, the driving mechanism shall be able to drive the solar cell wing to complete the working modes such as tracking, capturing, zeroing, stalling, increment, fixed angle holding, etc. 3.2.5.2.2 Driving torque The maximum output torque on the output shaft of the driving mechanism shall not be less than 2 times of the total resistance torque. 3.2.5.3 Load The driving mechanism shall be able to bear the unfolding load of solar cell array and the load of spacecraft in various flight conditions such as orbit change, maintenance and docking. The bearing capability reserve margin meets to the requirements of spacecraft special technical documents.
Foreword i 1 Scope 2 Normative references 3 Requirements 3.1 Composition 3.2 Performance requirements 3.3 Reliability 3.4 Safety 3.5 Environmental adaptability 3.6 Electromagnetic compatibility 3.7 Power consumption 3.8 Interface 3.9 Weight 3.10 Life span 3.11 Appearance 3.12 Maintainability 3.13 Interchangeability 3.14 Marking 4 Specification for quality assurance 4.1 Inspection classification 4.2 Inspection conditions 4.3 Appraisal inspection 4.4 Acceptance inspection 4.5 Inspection methods 5 Delivery preparation 5.1 Sealing and packaging 5.2 Packing 5.3 Transportation and storage 5.4 Marking 6 Instructions 6.1 Intended use 6.2 Ordering file information
Referred in GJB 9632-2019:
*GB/T 191-2008 Packaging - Pictorial Marking for Handling of Goods
*GJB 150.5A-2009 Laboratory environmental test methods for military materiel -Part 5:Temperature shock test
*GJB 150.15A-2009 Laboratory environmental test methods for military materiel - Part 15: Acceleration test
*GJB 150.16A-2009 Laboratory environmental test methods for military materiel-Part 16: Vibration test
*GJB 150.17A-2009 Laboratory environmental test methods for military materiel - Part 17: Acoustic noise test
*GJB 150.18A-2009 Laboratory environmental test methods for military materiel - Part 18: Shock test
*GJB151A-1997
*GJB152A-
*GJB1181-
*GJB2042A-2012
*GJB2602-1996
*GJB 2831A-2009 General specification for hermetically sealed Nickel-Hydrogen rechargeable cells in spacecraft
*GJB 2998-1997 Mark of satellite products
*GJB4038-2000
*GJB5174-2003
*GJB6789-2009
*GJBZ35-1993
*GJBZ1391-
*QJ 1019-1986
*QJ2630.1-
*QJ2630.3-
GJB 9632-2019 is referred in:
*QJ 20422.4-2016 Environment test methods for spacecraft unit- Part 4: Magnetic test
*GJB 1198.8A-2004 Telemetry tracking command and data handling for spacecraft Part 8:Onboard data handling interface
*GB/T 8145-1987 Gum rosin
*GB/T 8145-2003 Gum rosin
*GB/T 8145-2021 Gum rosin
*GB/T 20491-2006 Steel slag powder used for cement and concrete
*GB 11244-1989 General specifications for medical fiber endoscope
*GB 11244-2005 General requirements for the medical endoscope and endoscope accessories
*GB 19778-2005 Packaging glass containers—Release of lead cadmium arsenic and antimony—Permissible limits
*GB/T 13344-1992 Downhole drill hammers and bils
*GB/T 13344-2010 Downhole drill hammers and bits
*GB/T 13344-2019 Down-the-hole hammers and bits
*YY/T 0466.1-2009 Medical devices - Symbols to be used with medical device labels, labelling and information to be supplied - Part 1: General requirements
*YY/T 0466.1-2016 Medical devices - Symbols to be used with medical device labels, labelling and information to be supplied - Part 1: General requirements
*YY/T 0466.1-2023 Medical devices―Symbols to be used with information to be supplied by the manufacturer―Part 1:General requirements
*GB/T 14480.3-2008 Non-destructive testing - Equipment for eddy current examination - Part 3: System characteristics and verification
*GB/T 14480.3-2020 Non-destructive testing instruments—Equipment for eddy current examination—Part 3:System characteristics and verification
*YY 0666-2008 Method for the test of sharpness and strength of needles tips
*GB/T 7064-1996 Requirements for turbine type synchronous machine
*GB/T 7064-2002 Requirements for turbine type synchronous machine
*GB/T 7064-2017 Specific requirements for cylindrical rotor synchronous machines
*FZ/T 01026-1993 Quantitative Chemical Analysis of Quaternary Fibre Mixtures
*FZ/T 01026-2009 Textiles—Quantitative chemical analysis—Quaternary fibre mixtures
*FZ/T 01026-2017 Textiles - Quantitative chemical analysis - Multi-component fibre mixtures
*GM/T 0005-2012 Randomness Test Specification
*GB/T 9766.6-2008 Test method for tyre valve - Part 6: Test method for core
*GB/T 9766.6-2021 Test method for tyre valve - Part 6: Test method for core
*GB/T 18801-2002 Air cleaner
*GB/T 246-1997 Metallic materials--Tube--Flattening test
*GB/T 246-1982 Method for flattening test on tubes of metals
*GB/T 246-2007 Metal materials – Tube - Flattening test
Code of China
Standard
GJB 9632-2019  Specification for low orbit spacecraft high-voltage power supply system (English Version)
Standard No.GJB 9632-2019
Statusvalid
LanguageEnglish
File FormatPDF
Word Count11000 words
Price(USD)400.0
Implemented on2020-1-1
Deliveryvia email in 1 business day
Detail of GJB 9632-2019
Standard No.
GJB 9632-2019
English Name
Specification for low orbit spacecraft high-voltage power supply system
Chinese Name
低轨航天器高压母线电源系统规范
Chinese Classification
Professional Classification
GJB
ICS Classification
Issued by
Issued on
2019-12-08
Implemented on
2020-1-1
Status
valid
Superseded by
Superseded on
Abolished on
Superseding
Language
English
File Format
PDF
Word Count
11000 words
Price(USD)
400.0
Keywords
GJB 9632-2019, GJB/T 9632-2019, GJBT 9632-2019, GJB9632-2019, GJB 9632, GJB9632, GJB/T9632-2019, GJB/T 9632, GJB/T9632, GJBT9632-2019, GJBT 9632, GJBT9632
Introduction of GJB 9632-2019
Codeofchina.com is in charge of this English translation. In case of any doubt about the English translation, the Chinese original shall be considered authoritative. This specification was proposed by the Equipment Department of Aerospace Systems Department of PLA Strategic Support Force. Specification for low orbit spacecraft high-voltage power supply system 1 Scope This specification specifies the technical requirements, quality assurance regulations, delivery preparation and instructions of 100V high-voltage power supply system and its products for low orbit manned spacecraft. This specification is applicable to the development, test and acceptance of high-voltage fully regulated bus power supply system (hereinafter referred to as power supply system) with bus voltage of 100V for low orbit manned spacecraft. 2 Normative references The following normative documents contain provisions which, through reference in this text, constitute provisions of this specification. For dated or version indicated reference, subsequent amendments (excluding corrections), or revisions, of any of these publications do not apply to this specification. However, parties to agreements based on this specification are encouraged to investigate the possibility of applying the most recent editions of the normative documents indicated below. For undated references or references with version not indicated, the latest edition of the normative document referred to applies. GB/T 191 Packaging — Pictorial Marking for Handling of Goods GJB 150.5A-2009 Laboratory environmental test methods for military materiel — Part 5: Temperature shock test GJB 150.15A-2009 Laboratory environmental test methods for military materiel — Part 15: Acceleration test GJB 150.16A-2009 Laboratory environmental test methods for military materiel — Part 16: Vibration test GJB 150.17A-2009 Laboratory environmental test methods for military materiel — Part 17: Acoustic noise test GJB 150.18A-2009 Laboratory environmental test methods for military materiel — Part 18: Impact test GJB 151A-1997 Electromagnetic emission and susceptibility requirements for military equipment and subsystems GJB 152A Electromagnetic emission and susceptibility measurements for military equipment and subsystems GJB 1181 Packaging, handling, storage and transportability program requirements (for systems and equipment) GJB 2042A-2012 General specification for electrical power system of satellite GJB 2602-1996 General specification for space solar cell arrays GJB 2831A-2009 General specification for hermetically sealed nickel-hydrogen rechargeable cells in spacecraft GJB 2998 Mark of satellite products GJB 4038-2000 General specification for solar cell array mechanisms GJB 5174-2003 General specification for solar array drive assembly of tracking the sun GJB 6789-2009 General specification for lithium-ion rechargeable cells in spacecraft GJB/Z 35-1993 Derating criteria for electrical, electronic and electromechanical parts GJB/Z 1391 Guide to failure mode, effects and criticality analysis QJ 1019 Measurement method for electrical characteristics of solar cells QJ 2630.1 Space environment test methods for satellite components — Thermal vacuum test QJ 2630.3 Space environment test methods for satellite components — Vacuum discharge test 3 Requirements 3.1 Composition The composition of power supply system shall include: a) power generator: semi-rigid or flexible solar cell array; b) energy storage device: nickel-hydrogen battery pack or lithium-ion battery pack; c) sun-tracking device: including driving mechanism and actuator; d) power supply control device: including control equipment to realize the functions of main error amplification signal, shunt regulation, discharge regulation, charging control and bus filtering. 3.2 Performance requirements 3.2.1 Power bus 3.2.1.1 Bus voltage Nominal bus voltage: 100V. 3.2.1.2 Adjustment degree of bus voltage In regulation domain, the bus voltage regulation degree is within ±3%. 3.2.1.3 Bus voltage jump rate when entering in or exiting from shadow The voltage jump rate of bus shall be no more than 3.5V/ms and the voltage disturbance caused by bus shall be no more than 10%. 3.2.1.4 Output impedance of bus voltage Output impedance of bus shall be no more than 70 mΩ. 3.2.1.5 Bus ripple voltage In the frequency range of 10 kHz ~ 10 MHz and under rated output voltage and rated resistive load, the peak-to-peak ripple voltage of power supply system shall be no more than 500mV(p-p). 3.2.1.6 Transient characteristics of bus voltage When the transient of load current is less than 50% of the rated load, the bus output voltage shall not exceed 5% of the nominal bus voltage, and the adjustment time to dropping the voltage to within 1% of the nominal voltage shall not exceed 20ms. When the transient of load current exceeds 50% of the rated load, the bus output voltage shall not exceed 10% of the nominal bus voltage, and the adjustment time to dropping the voltage to within 1% of the nominal voltage shall not exceed 20ms. 3.2.1.7 Surge current The surge current provided by the bus shall meet the requirements of the maximum surge current when multiple loads are powered at the same time. The surge current rising slope is generally not less than 1×106A/s, and its duration is not less than 5ms, and the surge amplitude is not more than 50% of the rated load current of the bus. Under the above surge current, the voltage disturbance caused by bus shall not exceed 5%. 3.2.1.8 Power margin The output power margin of the power bus shall be at 5% ~ 10%. 3.2.2 Power supply of initiating explosive device Battery pack tap method is generally adopts for large pulse current power supply device like initiating explosive devices, generally adopts cell tap mode, and can meet the voltage requirements of initiating explosive devices under specified working temperature conditions. 3.2.3 Power generator 3.2.3.1 Output power The output power of solar cell array is determined by loading power, charging power, transmission loss and power margin of the spacecraft. Meanwhile, factors such as system test error, combination loss, optimal operating point voltage and current, effective illumination and on-orbit operating temperature shall be considered. The output power of end-of-life solar cell array is determined by the maximum output power at its early life, the working temperature of orbital solar cell array, effective illumination, particle radiation, attenuation coefficient of ultraviolet radiation and loss at high-low temperature cycle, etc., which shall meet the requirements of the provisions of special technical documents. 3.2.3.2 Output voltage The output voltage of end-of-life solar cell array is determined by bus voltage, battery pack charging voltage, isolation diode of power supply circuit, sampling resistance, cable voltage drop and other factors. The determination shall be leave with a certain margin. 3.2.3.3 Mechanical characteristics Mechanical characteristics of solar cell array are as follows: a) folded state of solar cell array: 1) the requirements of strength and stiffness under the launching mechanics environment shall be met, the solar cell array shall be free from relative sliding, and the solar panels shall not collide with each other; 2) the first-order natural frequency of the solar cell array in the folded state shall not be coupled with the natural frequency of the spacecraft, and shall meet the requirements of special technical documents of the spacecraft. b) unfolded state of solar cell array: 1) the impact load caused by the unfolding and locked of the solar cell array shall not exceed the impact allowable range of the solar cell array driving device and other components; 2) the first-order bending natural frequency after the solar cell array is unfolding and locked shall not be coupled with the loop frequency of the spacecraft control system; 3) the first-order torsion natural frequency after the solar cell array is unfolding and locked shall not be coupled with the driving frequency of the solar cell array driving device. 3.2.3.4 Thermal characteristics The solar cell array shall be able to withstand the influence of high-low temperature cycle in orbit. The temperature of solar cells shall be no higher than 110°C in illumination area and no lower than -100°C in shadow area. 3.2.3.5 Cell covering rate The requirements for the cell covering rate of solar cell array are as follows: a) the cell covering rate of the unfolding semi-rigid solar cell array is not less than 85%; b) the cell covering rate of flexible solar cell array shall meet the requirements of special technical documents for spacecraft. 3.2.3.6 Anti-blocking For the solar cell array circuit, protective measures shall be taken to prevent hot spots caused by local blocking, and the influence of blocking on the output power of the solar cell array shall be analyzed. 3.2.3.7 Remnant magnetic torques The remnant magnetic torques produced by the solar cell array shall not be greater than 0.4A·m2. 3.2.3.8 Anti-space-radiation and anti-atomic-oxygen The requirements of solar cell array for anti-space-radiation and anti-atomic-oxygen are as follows: a) the total dose equivalent to 1 MeV electron damage during the on-orbit life of the solar cell shall be determined. The ratio of the failure dose to the total dose of electron damage during the on-orbit life shall be no less than 2 times; b) protective measures shall be taken against ultraviolet rays, and the power loss caused by ultraviolet irradiation shall be considered; c) protective measures shall be taken against the corrosion caused by atomic oxygen for materials like polyimide film. 3.2.3.9 Anti-static Anti-static requirements of solar cell array are as follows: a) according to the requirements of 3.7.2, the solar cell array structure shall be grounded with high resistance; b) if the potential difference between adjacent circuits of the solar cell string is not more than 70V, the interval shall be not less than 2mm. 3.2.4 Energy storage device 3.2.4.1 Capacity The actual capacity of the battery pack at the beginning of its life shall be 110% of the rated capacity. The capacity deviation of cell in each battery pack shall not exceed 3% of the rated capacity. 3.2.4.2 Working voltage and charging efficiency 3.2.4.2.1 Working voltage of battery pack Under the worst working conditions (end-of-life, open circuit failure of a cell or a single parallel unit, maximum discharge depth and longest shadow period), the minimum discharge voltage of the battery pack after deducting the line voltage drop shall not be less than the minimum value required by the bus voltage or discharge regulator. Under the worst working conditions (end-of-life, open circuit failure of a cell or a single parallel unit), the highest charging voltage of the battery pack after adding the line voltage drop shall not be greater than the voltage of best working point of the solar cell array or the highest value required by the charging regulator. 3.2.4.2.2 Discharge voltage, charging voltage and charging efficiency of cell The average discharge voltage, charging voltage and ampere-hour charging efficiency of cell on-orbit operation for 3a shall meet the requirements in Table 1. If the cell has been in orbit for more than 3a, it is necessary to analyze and determine the average discharge voltage, charging voltage and charging efficiency according to the ground test data.   Table 1 Average discharge voltage, maximum end-of-charge voltage and charging efficiency of battery pack (on-track 3a) Item Nickel-hydrogen cell Lithium-ion cell Average discharge voltage 1.25V 3.6Va end-of-charge voltage ≤1.65V ≤4.2V Ampere-hour charging efficiency ≥90% ≥95% Average discharge depth ≤30% ≤20% a for lithium cobalt oxides system cell. 3.2.4.3 Temperature gradient The maximum temperature difference between cells in the same cell module shall not exceed 3°C, and the maximum temperature difference between modules in the same pack shall not exceed 5°C. 3.2.4.4 Redundant backup Redundancy is conducted by hot backup with one or two cells. The battery pack shall be able to work if a cell is failed, and still able to work normally if a cell is short-circuited or open-circuited. 3.2.4.5 Anti-open-circuit When a cell in the battery pack is open-circuited, the battery pack shall be able to charge and discharge normally. The requirements for different battery packs are as follows: a) for nickel-hydrogen battery pack, the charging and discharging channels of the battery packs shall be provided when the open circuit fails; b) for lithium-ion battery pack, single parallel connection mode or anti-open-circuit with bypass device may be adopted. 3.2.4.6 Discharge depth, working temperature and cycle life The discharge depth, working temperature and cycle life of the battery pack in orbit for 3a shall meet the requirements in Table 2: Table 2 Average discharge depth, optimum working temperature and cycle life Item Nickel-hydrogen cell Lithium-ion cell Average discharge depth ≤30% ≤20% Optimum working temperature 0°C~10°C 10°C~30°C Cycle life 18000 times 18000 times For the cell in orbit for more than 3a, it is necessary to analyze and determine the discharge depth and working temperature according to the ground life test data. 3.2.4.7 Equalization processing For the power supply system using lithium-ion battery pack, the voltage of each cell or parallel single unit in the battery pack shall be controlled within 60mV deviation by means of equalization processor. 3.2.5 Sun-tracking device 3.2.5.1 Electrical property The electrical property of the drive mechanism include: a) the electric transmission capacity (current and voltage) of the conductive ring; b) the transmission voltage drops of conductive ring; c) transmission noise. The electrical performance of the driving mechanism shall meet the requirements of special technical documents for spacecraft. 3.2.5.2 Driving ability 3.2.5.2.1 Drive work mode According to the driving pulse input by the driver, the driving mechanism shall be able to drive the solar cell wing to complete the working modes such as tracking, capturing, zeroing, stalling, increment, fixed angle holding, etc. 3.2.5.2.2 Driving torque The maximum output torque on the output shaft of the driving mechanism shall not be less than 2 times of the total resistance torque. 3.2.5.3 Load The driving mechanism shall be able to bear the unfolding load of solar cell array and the load of spacecraft in various flight conditions such as orbit change, maintenance and docking. The bearing capability reserve margin meets to the requirements of spacecraft special technical documents.
Contents of GJB 9632-2019
Foreword i 1 Scope 2 Normative references 3 Requirements 3.1 Composition 3.2 Performance requirements 3.3 Reliability 3.4 Safety 3.5 Environmental adaptability 3.6 Electromagnetic compatibility 3.7 Power consumption 3.8 Interface 3.9 Weight 3.10 Life span 3.11 Appearance 3.12 Maintainability 3.13 Interchangeability 3.14 Marking 4 Specification for quality assurance 4.1 Inspection classification 4.2 Inspection conditions 4.3 Appraisal inspection 4.4 Acceptance inspection 4.5 Inspection methods 5 Delivery preparation 5.1 Sealing and packaging 5.2 Packing 5.3 Transportation and storage 5.4 Marking 6 Instructions 6.1 Intended use 6.2 Ordering file information
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Keywords:
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