Spacecraft test energy adaptive control system

Abstract

The invention relates to a spacecraft test energy self-adaptive control system, which comprises a telemetering subscription and processing unit, which is used for obtaining bus and load working conditions of a spacecraft and charging and discharging conditions of a storage battery through telemetering data, and counting and calculating an energy list of a ground power supply; the energy list of the ground power supply comprises volt-ampere characteristics and maximum power supply current of the ground power supply system corresponding to each typical working condition; the data analysis and statistics unit is used for calculating the state of charge (SOC) of the storage battery by adopting an extended Kalman filter algorithm based on a second-order RC circuit model and a calculation model of the state of charge of the storage battery; the energy planning unit is used for comparing the output voltage Ubat of the storage battery with a preset charging cut-off voltage UGC and a preset discharging cut-off voltage UGF, and carrying out energy distribution on each ground power supply according to the state of charge SOC of the storage battery; and the energy management unit is used for carrying out balance control on each storage battery monomer according to the charge state of the storage battery and the monomer voltage dispersion.

Description

Technical Field

[0001] The invention relates to a spacecraft test energy adaptive control system, belonging to the technical field of spacecraft integrated testing, and in particular to a ground test platform and a test method for a power supply of a spacecraft with a fast load change speed and a large range. Background Art

[0002] In the traditional spacecraft testing process, the energy adjustment method of the entire satellite is to supply energy based on the actual load demand by manually selecting a specific volt-ampere curve under the typical operating conditions of the spacecraft to meet the energy security under different test conditions.

[0003] However, with the rapid development of aerospace technology, the number of concurrent models is large, the comprehensive test cycle is short, and the cost constraints are strong, which puts higher requirements on the safety and reliability of the power supply and distribution subsystem test. The original energy adjustment method requires manual evaluation of the ground power supply capacity and the load of the entire satellite, and dynamically selects the appropriate volt-ampere curve to meet the power distribution of the entire system. This method requires real-time attention to the working conditions of the spacecraft for manual adjustment, which greatly reduces the test efficiency, and lacks the means to evaluate the energy usage of the spacecraft in real time, including battery charge state, ground power supply capacity, spacecraft load power consumption, etc., making it difficult to achieve intelligent management of energy during model testing. Summary of the invention

[0004] The technical problem solved by the present invention is: in view of the many shortcomings of the traditional mode of manually adjusting the energy supply capacity of spacecraft testing, a spacecraft test energy adaptive control system is provided, which is applied in the daily testing stage of spacecraft. While releasing human resources, it can effectively carry out full-cycle load power consumption analysis during the spacecraft testing process and realize intelligent energy management.

[0005] The solution to the technical problem of the present invention is: a spacecraft test energy adaptive control system, the system includes a telemetry subscription and processing unit, a data analysis and statistics unit, an energy planning unit, and an energy management unit;

[0006] The telemetry subscription and processing unit establishes a connection with the spacecraft master control test system, subscribes to the telemetry data of the ground power supply and the spacecraft master control test system, obtains the bus and load conditions of the spacecraft, the charging and discharging conditions of the battery through the telemetry data, and counts and calculates the energy list of the ground power supply; the energy list of the ground power supply includes the volt-ampere characteristics and the maximum supply current of the ground power supply system corresponding to each typical working condition;

[0007] The data analysis and statistics unit constructs a second-order RC circuit model of the battery, uses the ampere-hour integration method to establish a calculation model for the battery state of charge, and uses the extended Kalman filter algorithm based on the second-order RC circuit model and the calculation model for the battery state of charge to calculate the battery state of charge SOC;

[0008] Energy planning unit, compares the battery output voltage U bat With the preset charging cut-off voltage U GC , discharge cut-off voltage U GF , if U bat >U GC When U bat GF When the ground power source selects the volt-ampere characteristic output corresponding to the maximum power supply current in the energy list and gives a warning; otherwise, the energy is distributed to each ground power source according to the battery state of charge SOC;

[0009] The energy management unit prohibits the balancing management function and stops all balancing actions if the battery is in the charge and discharge state. Otherwise, the balancing management function is enabled to perform balancing control on the battery cells according to the battery charge state and cell voltage dispersion.

[0010] Preferably, the spacecraft test energy adaptive control system is time synchronized with the aerospace master control test system.

[0011] Preferably, the telemetry data includes the operating time t, the voltage U at the maximum operating point of the ground power supply, mp and current I mp , spacecraft bus voltage U bus The total load current I bus , battery cell voltage, battery output voltage U bat 、Battery charge and discharge current I bat , Battery charging voltage level setting value U set .

[0012] Preferably, the maximum power supply capacity of the ground power supply equipment corresponding to each typical working condition is obtained by the following method:

[0013] By comparing the maximum operating point voltages of all ground power supplies under different typical operating conditions with the bus voltage of the spacecraft, and accumulating the maximum operating point currents of the ground power supplies that are higher than the bus voltage, the corresponding maximum supply currents of the ground power supplies under different typical operating conditions can be obtained.

[0014] Preferably, the second-order RC circuit model of the battery includes resistors R0, R b , R p , capacitor C b , C​p , equivalent power supply;

[0015] Resistor R b and capacitor C b connected in parallel to form the first RC parallel circuit; the resistor R p and capacitor C p The positive pole of the equivalent power supply is connected to one end of the resistor R0, and the other end of the resistor R0 is connected in series with the first RC parallel circuit and the second RC parallel circuit. The output end of the second RC parallel circuit is the positive pole of the battery terminal voltage, and the negative pole of the equivalent power supply is the negative pole of the battery terminal voltage.

[0016] Preferably, the R b and C b The values ​​are the double layer capacitance and resistance of the battery; R p and C p The values ​​are the capacitance and resistance of the diffusion phenomena inside the battery.

[0017] Preferably, the state of charge SOC of the battery is calculated by the following formula:

[0018]

[0019] Among them, Q n is the rated capacity of the battery, I bat is the battery charge and discharge current, η is the charge and discharge rate, Δt is the calculation cycle, SOC k is the battery state of charge of the kth calculation cycle, SOC k+1 is the battery state of charge SOC of the k+1th calculation cycle.

[0020] Preferably, the state quantity of the Kalman filter equation is:

[0021] The state quantity is: [SOC U b U p ] T

[0022] The state equation is:

[0023]

[0024] The input quantity is: I bat

[0025] The observed quantity is: U bat

[0026] The observation equation is:

[0027]

[0028] Among them, SOC is the battery state of charge, Ts is the calculation period, U bat is the battery terminal voltage; U OC is the open circuit voltage of the battery; R0 represents the battery pack; U b and U p are the voltages of the first RC parallel circuit and the second RC parallel circuit respectively; R b and C b Respectively, the resistance R b The resistance and capacitance C b Capacitance value; R p and C p Respectively, the resistance R p The resistance and capacitance C p Capacitance value; U bat is the output voltage of the battery, I bat is the charge and discharge current of the battery, k and k+1 represent the kth calculation cycle and the k+1th calculation cycle respectively, Q n Calibrate the capacity of the battery.

[0029] Preferably, the energy management method is:

[0030] Calculate the average total load current I within the preset time period before the current moment bus , compare the battery's current state of charge SOC with the preset spacecraft energy management monitoring point. When the battery's current state of charge SOC is greater than 1.05, the ground power system selects a power supply lower than I from the maximum power supply capacity list. bus The volt-ampere characteristic curve corresponding to the maximum supply current is output. If it does not exist, the ground power supply output is prohibited and a warning is given. If the current state of charge SOC of the battery is less than 0.95, the ground power system selects a higher voltage than I from the maximum power supply capacity list. bus The volt-ampere characteristic curve output corresponding to the maximum supply current is output. If it does not exist, the current state is maintained and no action is taken.

[0031] Preferably, the spacecraft is provided with a balancing switch for each battery cell. When the balancing switch is turned on, the output voltage of the battery remains unchanged. When the balancing switch is turned on, the output voltage of the battery cell is reduced. The balancing switch state of the battery cell is obtained by telemetry data collection.

[0032] The balancing method is: monitoring all effective battery cell voltages, selecting the cell with the lowest voltage in the battery, and the difference between the voltage of other cells and the lowest voltage is greater than a first threshold value U JHK When the number of battery cells that have been connected to the balancing switch is ≤ N MAX , then the balancing switch of the battery cell with higher output voltage is turned on, and the number of connected cells is increased by 1; when the voltage difference between the cell turned on by the balancing switch and the lowest cell is less than the second threshold value U JHGWhen the number of battery cells with equalizing switches turned on is ≥ 0, the equalizing switches of the battery cells with lower output voltages are disconnected, and the number of battery cells with equalizing switches turned on is -1, N MAX is the total number of battery cells.

[0033] The beneficial effects of the present invention compared with the prior art are:

[0034] (1) The spacecraft test energy adaptive control system provided by the present invention establishes a communication interface with the master control test system, and based on real-time satellite-to-ground telemetry information, calculates the current ground energy list, spacecraft load conditions, and battery charge status in real time, and adopts spacecraft energy management technology based on multi-criteria limitation to achieve autonomous control of the flow of satellite-to-ground energy.

[0035] (2) Based on the battery state of charge and single cell voltage dispersion, it can autonomously complete the control of balancing instructions to increase the service life of the spacecraft battery pack.

[0036] (3) The entire system can not only reduce labor costs and achieve energy adaptation during spacecraft testing, but also complete autonomous management of batteries, thereby improving the efficiency of spacecraft comprehensive testing and the ability of automated testing. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Figure 1 This is a platform architecture diagram for testing an energy adaptive control system for a spacecraft according to an embodiment of the present invention;

[0038] Figure 2 This is a diagram of an equivalent circuit model of a double RC battery according to an embodiment of the present invention;

[0039] Figure 3 This is a flow chart of spacecraft energy management based on multiple criteria according to an embodiment of the present invention. DETAILED DESCRIPTION

[0040] The present invention will be further described below in conjunction with the embodiments.

[0041] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0042] In order to overcome the situation of relying on manual interpretation and manual adjustment of power supply curves during spacecraft comprehensive testing, the present invention provides a spacecraft test energy adaptive control system to analyze, count and manage the energy flow of the entire system, and realize adaptive adjustment of energy during model testing.

[0043] The spacecraft test energy adaptive control system is composed of ① telemetry subscription and processing unit, ② data analysis and statistics unit, and ③ energy planning and management unit. Its characteristics are:

[0044] The telemetry subscription and processing unit ① establishes a connection with the spacecraft master control test system through the TCP / IP protocol, subscribes to the telemetry data of the ground power supply and the spacecraft master control test system, obtains the bus and load conditions of the spacecraft and the charge and discharge conditions of the battery through the telemetry data, and counts and calculates the energy list of the ground power supply; the energy list of the ground power supply includes the volt-ampere characteristics and the maximum supply current of the ground power supply system corresponding to each typical working condition;

[0045] The data analysis and statistics unit ② constructs a second-order RC circuit model of the battery, uses the ampere-hour integration method to establish a calculation model for the battery state of charge, and based on the second-order RC circuit model and the calculation model for the battery state of charge, uses an extended Kalman filter algorithm (Extended Kalman Filtering, EKF) to calculate the battery state of charge (SOC). However, the accuracy of this method is based on the precise measurement of the battery current. If you want to use this method to obtain a higher estimation accuracy, you must ensure a higher current measurement accuracy. Therefore, according to the characteristics of the spacecraft comprehensive test, calibration measures are added to eliminate the impact of cumulative errors on the estimation of the battery state of charge.

[0046] The energy planning and management unit ③ adopts the spacecraft energy management technology based on multi-criteria limitation, plans and manages the use of the ground energy curve according to the current ground energy list, spacecraft load conditions, and battery charge state, and then completes the balance control according to the battery charge state and single cell voltage dispersion.

[0047] The energy planning and management unit includes an energy planning unit and an energy management unit.

[0048] Energy planning unit, compares the battery output voltage U bat With the preset charging cut-off voltage U GC , discharge cut-off voltage U GF , if U bat >U GC When U bat GF When the ground power source selects the volt-ampere characteristic output corresponding to the maximum power supply current in the energy list and gives a warning; otherwise, the energy is distributed to each ground power source according to the battery state of charge SOC;

[0049] ​The energy management unit prohibits the balancing management function and stops all balancing actions if the battery is in the charge and discharge state. Otherwise, the balancing management function is enabled to perform balancing control on the battery cells according to the battery charge state and cell voltage dispersion.

[0050] According to this design, the spacecraft test energy adaptive control system is synchronized with the entire satellite time. Through the telemetry data between the satellite and the ground, the actual working status of the ground power supply is judged, and the adaptive control of the energy and the balanced control of the spacecraft battery pack in the comprehensive test phase are realized.

[0051] The core idea of ​​the present invention is to use the power supply capacity of the ground power supply, the bus system and load conditions of the spacecraft, the charging and discharging conditions of the battery and other data as external inputs of the spacecraft test energy adaptive control system, and bring them into ① telemetry subscription and processing unit, ② data analysis and statistics unit, and ③ energy planning and management unit for simulation calculation, calculate the current ground energy list, spacecraft load conditions, battery charge state and single cell voltage dispersion, and realize energy adaptive management of spacecraft testing and autonomous balancing control of battery packs.

[0052] The implementation process of the present invention is described in detail below. The present invention provides an energy flow management design for spacecraft comprehensive testing, such as Figure 1 As shown, building a spacecraft test energy adaptive control system includes the following steps:

[0053] (1) The spacecraft energy adaptive control system is built using the LabVIEW platform, and the telemetry parameter subscription configuration is performed in the background, including the spacecraft time, the spacecraft bus and load conditions, the battery working status, etc. The TCP / IP protocol is used to establish a connection with the spacecraft master control test system, and the data acquisition of the corresponding telemetry code is completed through the RTS data service interface.

[0054] (2) Obtaining original telemetry data, which includes the real-time running time t, the voltage U at the maximum operating point of the ground power supply mp and current I mp , spacecraft bus voltage U bus The total load current I bus , the single cell voltage and the whole group voltage U bat 、Battery charge and discharge current I bat , Battery charging voltage level setting value U set .

[0055] Based on the information from telemetry data, the maximum power supply capacity of the ground power equipment corresponding to each typical working condition is obtained by the following method:

[0056] By comparing the maximum operating point voltages of all ground power supplies under typical operating conditions with the bus voltage of the spacecraft, and accumulating the maximum operating point currents of the ground power supplies that are higher than the bus voltage, the corresponding maximum supply current of the ground power supplies under typical operating conditions can be obtained.

[0057] (3) Construct a second-order RC circuit model of the battery

[0058] like Figure 2 As shown, the second-order RC circuit model of the battery includes resistors R0, R b , R p , capacitor C b , C p , equivalent power supply;

[0059] Resistor R b and capacitor C b connected in parallel to form the first RC parallel circuit; the resistor R p and capacitor C p The positive pole of the equivalent power supply is connected to one end of the resistor R0, and the other end of the resistor R0 is connected in series with the first RC parallel circuit and the second RC parallel circuit. The output end of the second RC parallel circuit is the positive pole of the battery terminal voltage, and the negative pole of the equivalent power supply is the negative pole of the battery terminal voltage.

[0060] The zero-state response equation of the battery terminal voltage is:

[0061]

[0062] Among them, U bat is the battery terminal voltage; U OC is the open circuit voltage of the battery; R0 represents the internal resistance of the battery, including the resistance generated by various film layers on the electrode surface; U b and U p are the voltages of the two RC loops respectively; b and C b The values ​​are the double-layer capacitance and resistance of the battery. When two phases with higher conductivity come into contact, the charge will be redistributed. This process can be simulated by the double-layer capacitance and resistance. p and C p The values ​​are the capacitance and resistance of the diffusion phenomena inside the battery.

[0063] In practical engineering applications, the only thing that can be directly measured by the battery is the battery terminal voltage U bat and current I bat and temperature, while other parameters of the battery can be obtained through identification provided by the manufacturer or using experimental data.

[0064] (4) The battery state of charge calculation model is established using the ampere-hour integration method. Assuming that the initial SOC value of the battery at time t0 before operation is SOC0, then the SOC value of the battery at time t is:

[0065]

[0066] Among them, Q n is the rated capacity of the battery, I bat is the battery charge and discharge current, τ is the time integral, and η is the charge and discharge rate. For ease of calculation, the above formula can be discretized to obtain the discrete equation expression of the ampere-hour integration method:

[0067]

[0068] Among them, Q n is the rated capacity of the battery, I bat is the battery charge and discharge current, η is the charge and discharge rate, Δt is the calculation cycle, SOC k is the battery state of charge of the kth calculation cycle, SOC k+1 is the battery state of charge SOC of the k+1th calculation cycle.

[0069] The above model can be used to calculate the battery power change at each moment to estimate the remaining battery power in real time.

[0070] (5) Establish a system model for estimating SOC based on the EKF algorithm of the second-order RC circuit, and use SOC and U b , U p As the state vector of the algorithm system, the battery terminal voltage equation is used as the observation equation of the system. The equivalent model and the ampere-hour integral definition are discretized to obtain the discretized state equation of the model:

[0071]

[0072] The input quantity is: I bat

[0073] The observed quantity is: U bat

[0074] The discretized observation equation of the model is:

[0075]

[0076] Among them, SOC is the battery state of charge, T s is the calculation period, U bat is the battery terminal voltage; U OC is the open circuit voltage of the battery; R0 represents the battery pack; U b and U pare the voltages of the first and second RC parallel circuits respectively; R b and C b are the resistance value of resistor R b and the capacitance value of capacitor C b respectively; R p and C p are the resistance value of resistor R p and the capacitance value of capacitor C p respectively; U bat is the output voltage of the storage battery, I bat is the charge and discharge current of the storage battery, k and k + 1 respectively represent the k-th and (k + 1)-th calculation cycles, and Q n is the rated capacity of the storage battery.

[0077] Substitute the above equations into the EKF algorithm to estimate a relatively accurate SOC value.

[0078] (6) Energy management method

[0079] To ensure the safety of the comprehensive test of the spacecraft, set the charging cut-off voltage U GC and the discharging cut-off voltage U GF : ① When U bat > U GC , the ground power supply is prohibited from outputting and a warning is given; ② When U bat < U GF , the ground power supply selects the volt-ampere characteristic corresponding to the maximum supply current in the energy list for output and gives a warning; otherwise, perform energy allocation for each ground power supply according to the state of charge SOC of the storage battery:

[0080] Based on the current state of charge SOC of the battery, set the energy management monitoring points of the spacecraft. When the estimated SOC reaches the monitoring point, calculate the average total load current I bus within the current 5 minutes and select the power supply curve: ① When SOC > 1.05, the ground power supply system needs to select the maximum volt-ampere characteristic curve lower than I bus from the maximum supply capacity list for output; ② When SOC < 0.95, the ground power supply system needs to select the minimum volt-ampere characteristic curve higher than I bus from the maximum supply capacity list for output; ③ When 0.95 < SOC < 1.05, then keep the current state without taking any action.

[0081] (8) Equalization management

[0082] Equalization management needs to be carried out when the storage battery is not in the charge and discharge state, that is, |I bat | < 2A. If not satisfied, the equalization management function is prohibited and all equalization actions are stopped; if satisfied, the equalization management function is enabled.

[0083] The spacecraft is equipped with a balancing switch for each battery cell. When the balancing switch is turned off, the battery output voltage remains unchanged. When the balancing switch is turned on, the battery output voltage is reduced. The balancing switch status of the battery cell is obtained through telemetry data collection.

[0084] When the balancing management function is enabled, the balancing management process, such as Figure 3 As shown: monitor all valid battery cell voltages, select the cell with the lowest voltage in the battery, and the difference between the voltage of other cells and the lowest voltage is greater than the first threshold value U JHK When the number of battery cells that have been connected to the balancing switch is ≤ N MAX , then the balancing switch of the battery cell with higher output voltage is turned on, and the number of connected cells is increased by 1; when the voltage difference between the cell turned on by the balancing switch and the lowest cell is less than the second threshold value U JHG When the number of battery cells with equalizing switches turned on is ≥ 0, the equalizing switches of the battery cells with lower output voltages are disconnected, and the number of battery cells with equalizing switches turned on is -1, N MAX is the total number of battery cells.

[0085] If the current spacecraft integrated test is completed, exit; otherwise, continue to execute the energy adaptive management of the spacecraft test.

[0086] The present invention has been successfully applied to the whole-satellite system-level test of a certain traditional remote sensing satellite with good effect. The present invention can be further extended to the test of other remote sensing satellites or other fields.

[0087] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A spacecraft test energy adaptive control system, characterized in that It includes telemetry subscription and processing unit, data analysis and statistics unit, energy planning unit and energy management unit; The telemetry subscription and processing unit establishes a connection with the spacecraft master control test system, subscribes to the telemetry data of the ground power supply and the spacecraft master control test system, obtains the bus and load conditions of the spacecraft, the charging and discharging conditions of the battery through the telemetry data, and counts and calculates the energy list of the ground power supply; the energy list of the ground power supply includes the volt-ampere characteristics and the maximum supply current of the ground power supply system corresponding to each typical working condition; The data analysis and statistics unit constructs a second-order RC circuit model of the battery, uses the ampere-hour integration method to establish a calculation model for the battery state of charge, and uses the extended Kalman filter algorithm based on the second-order RC circuit model and the calculation model for the battery state of charge to calculate the battery state of charge SOC; Energy planning unit, compares the battery output voltage U bat With the preset charging cut-off voltage U GC , discharge cut-off voltage U GF If U bat >U GC When U bat GF When the ground power source selects the volt-ampere characteristic corresponding to the maximum supply current in the energy list, it outputs and gives a warning;​ Otherwise, energy is distributed to various ground power sources according to the battery state of charge SOC; The energy management unit prohibits the balancing management function and stops all balancing actions if the battery is in the charge and discharge state. Otherwise, the balancing management function is enabled to perform balancing control on the battery cells according to the battery charge state and cell voltage dispersion.

2. A spacecraft test energy adaptive control system according to claim 1, characterized in that The spacecraft test energy adaptive control system is time synchronized with the aerospace master control test system.

3. A spacecraft test energy adaptive control system according to claim 1, characterized in that The telemetry data includes the running time t, the voltage U at the maximum operating point of the ground power supply mp and current I mp , spacecraft bus voltage U bus With the total load current I bus , battery cell voltage, battery output voltage U bat 、Battery charge and discharge current I bat , Battery charging voltage level setting value U set .

4. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The maximum power supply capacity of the ground power equipment corresponding to each typical working condition is obtained by the following method: By comparing the maximum operating point voltages of all ground power supplies under typical operating conditions with the bus voltage of the spacecraft, and accumulating the maximum operating point currents of the ground power supplies that are higher than the bus voltage, the corresponding maximum supply current of the ground power supplies under typical operating conditions can be obtained.

5. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The second-order RC circuit model of the battery includes resistors R0, R b , R p , capacitor C b , C p , equivalent power supply; Resistor R b and capacitor C b connected in parallel to form the first RC parallel circuit; the resistor R p and capacitor C p The first and second RC parallel circuits are connected in parallel to form a second RC parallel circuit; the positive electrode of the equivalent power supply is connected to one end of the resistor R0, and the other end of the resistor R0 is connected in series with the first RC parallel circuit and the second RC parallel circuit, the output end of the second RC parallel circuit is the positive electrode of the battery terminal voltage, and the negative electrode of the equivalent power supply is the negative electrode of the battery terminal voltage.

6. The spacecraft test energy adaptive control system according to claim 5, characterized in that: The R b and C b The values ​​are the double-layer capacitance and resistance of the battery; R p and C p The values ​​are the capacitance and resistance of the diffusion phenomena inside the battery.

7. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The state of charge SOC of the battery is calculated by the following formula: Among them, Q n is the rated capacity of the battery, I bat is the battery charge and discharge current, η is the charge and discharge rate, Δt is the calculation cycle, SOC k is the battery state of charge of the kth calculation cycle, SOC k+1 is the battery state of charge SOC of the k+1th calculation cycle.

8. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The state quantity of the Kalman filter equation is: The state quantity is: [SOC U b U p ] T The state equation is: The input quantity is: I bat The observed quantity is: U bat The observation equation is: Among them, SOC is the battery state of charge, T s is the calculation period, U bat is the battery terminal voltage; U OC is the open circuit voltage of the battery; R0 represents the battery pack; U b and U p are the voltages of the first RC parallel circuit and the second RC parallel circuit respectively; R b and C b Respectively, the resistance R b The resistance and capacitance C b Capacitance value; R p and C p Respectively, the resistance R p The resistance and capacitance C p Capacitance value; U bat is the output voltage of the battery, I bat is the charge and discharge current of the battery, k and k+1 represent the kth calculation cycle and the k+1th calculation cycle respectively, Q n Calibrate the capacity of the battery.

9. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The energy management method is: Calculate the average total load current I within the preset time period before the current moment bus , compare the battery's current state of charge SOC with the preset spacecraft energy management monitoring point. When the battery's current state of charge SOC is greater than 1.05, the ground power system selects a power supply lower than I from the maximum power supply capacity list. bus The volt-ampere characteristic curve corresponding to the maximum supply current is output. If it does not exist, the ground power supply output is prohibited and a warning is given. If the current state of charge SOC of the battery is less than 0.95, the ground power system selects a higher voltage than I from the maximum power supply capacity list. bus The volt-ampere characteristic curve output corresponding to the maximum supply current is output. If it does not exist, the current state is maintained and no action is taken.

10. The spacecraft test energy adaptive control system according to claim 1, characterized in that: The spacecraft is equipped with a balancing switch for each battery cell. When the balancing switch is turned on, the battery output voltage remains unchanged. When the balancing switch is turned on, the output voltage of the battery cell is reduced. The balancing switch status of the battery cell is obtained through telemetry data collection. The balancing method is: monitoring all effective battery cell voltages, selecting the cell with the lowest voltage in the battery, and the difference between the voltage of other cells and the lowest voltage is greater than a first threshold value U JHK When the number of battery cells that have been connected to the balancing switch is ≤ N MAX , then the balancing switch of the battery cell with higher output voltage is turned on, and the number of connected cells is increased by 1; when the voltage difference between the cell turned on by the balancing switch and the lowest cell is less than the second threshold value U JHG When the number of battery cells with equalizing switches turned on is ≥ 0, the equalizing switches of the battery cells with lower output voltages are disconnected, and the number of battery cells with equalizing switches turned on is -1, N MAX is the total number of battery cells.

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