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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.
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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.
