Design of ignition control system for failed spacecraft array solid thruster

The life of micro-nano spacecraft is generally short. In order to solve the problem of space debris endangering space safety after its failure and make it deorbit quickly, this paper has carried out the hardware system design and software system design of the ignition control system of the failed spacecraft attitude adjustment solid thruster. In order to facilitate the realization of high integration, the hardware system adopts the array arrangement circuit and selects the MOSFET as the driving switch. The software system receives the ignition command from the spacecraft platform and enables the row and column control terminals of the ignition circuit to achieve addressing ignition. The results show that the control module can correctly receive the instructions from the spacecraft platform and make the correct ignition action, successfully complete the attitude adjustment task, and improve the reliability of the deorbiting of the failed spacecraft.


Introduction
In recent years, micro-nano spacecrafts have been continuously expanded in the civil and military fields due to their significant advantages such as low cost, short development cycle and formation flight.The number of launches has shown an ' explosive ' growth, forming a large-scale constellation development trend.However, the life of these spacecrafts is generally within 5 years.After failure, they will stay in space for a long time to occupy orbital resources, and are likely to cause a large number of debris, causing serious accidents and even chain reactions, which will bring extremely adverse effects to high-value spacecrafts and normal aerospace activities.Therefore, in order to sustain the development of space resources and slow down the rapid increase of space debris density, it is urgent to develop a system technology that can realize the full autonomous, high reliable and fast maneuvering of the failed spacecraft to realize the autonomous and rapid deorbiting of the failed spacecraft [1,2].
At present, the common deorbiting methods at home and abroad are as follows: deorbiting sail deorbiting, tether deorbiting, solar sail deorbiting and electric propulsion deorbiting [3].However, the thrust of these de-orbiting methods is at the milli-cow level, the maneuverability is poor, and the deorbiting takes a long time.The solid propulsion system has the advantages of simple structure, high integration and reliable operation.It can produce great total impulse in a short time and realize fast ignition maneuver [4].Based on the above background, this paper carried out the design of the ignition control system of the attitude-adjustable solid thruster of the failed spacecraft.The ignition command is sent to the attitude-adjustable solid thruster at the end of the spacecraft 's life.The solid thruster is reliably ignited after receiving the command, completing the attitude adjustment of the failed spacecraft and improving the reliability of the fast maneuver deorbit of the failed spacecraft.

System project design
The ignition control system of the failed spacecraft attitude adjustment solid thruster mainly includes two parts : hardware circuit part and software control part.The system workflow is shown in Figure 1.Spacecraft deorbit is mainly divided into three stages.Firstly, the satellite computer calculates which thrusters need to be ignited according to the attitude information of the spacecraft, and transmits the ignition information packet to the core control module through RS422 serial communication.Then the core control module parses the received data packets and generates an enabling signal.Finally, the ignition circuit configures the relay and the MOSFET drive circuit under the control of the enabling signal, so that the lithium battery can discharge the thruster igniter, the thruster is successfully ignited, and the attitude of the spacecraft is adjusted, thereby improving the reliability of the fast maneuver deorbiting of the failed spacecraft.

Ignition control hardware system design
The ignition circuit and the drive circuit are the core components of the ignition control hardware system.Whether the attitude adjustment thruster can work normally also depends on the ignition circuit and the drive circuit.

Firing circuit
The attitude thruster described in this paper is composed of 16 thruster units, and the arrangement is 2×8, so the wiring method of the thruster ignition circuit adopts the array ignition circuit.The number of ignition leads in the array circuit is small, the circuit structure is simple, and it is of great benefit to integration, so it is very suitable for large-scale array ignition [5].However, this method will cause the ignition circuit to have a closed quadrilateral, resulting in the wrong ignition of the thruster.Therefore, a certain ignition rule needs to be set when controlling the ignition.The ignition rules will be explained in the subsequent software design.
The structure of the ignition circuit of the ignition system is shown in Figure 2. The ignition circuit uses the ignition head as the heating element.After the ignition head is energized, it heats the electric explosion to ignite the surrounding ignition powder.The explosion and combustion of the ignition powder makes the pressure and temperature conditions established inside the combustion chamber of the thruster, so that the propellant is ignited and the ignition operation is completed.The ignition head and the diode are connected in series.The diode can ensure the one-way flow of the current, prevent the occurrence of false ignition, and prevent the formation of parallel connection between the ignition head and the ignition head from reducing the resistance [6].The column end COLi is connected to the power supply end through the column drive control circuit, and the row end ROWi is connected to the ground end through the row drive control circuit.When it is necessary to ignite the ignition head at the intersection of the column COLi and the row ROWi, it is only necessary to turn off the driving circuit corresponding to the row and column, so that the ignition voltage, the ignition circuit and the ground form a loop to heat the ignition head.

Drive circuit
MOSFET is a semiconductor device that controls the current of the drain at the output end by the voltage added to the gate of the input end, so as to realize the function of small voltage controlling large current.MOSFET has high speed on / off response, good high frequency performance and high input impedance [7].Therefore, the driving circuit of the ignition system in this paper uses MOSFET as the control element.According to the principle that N-MOSFET is suitable for source grounding and P-MOSFET is suitable for high level of source connection, a single drive circuit is designed as shown in Figure 3.In Figure 3, the left side is the column drive circuit, which is mainly composed of a three-stage tube and two P-MOSFETs.In order to improve the safety and reliability of the line, the P-MOSFET adopts a redundant backup design, so that two P-MOSFETs are connected in parallel.One input side of the column drive circuit is connected to the ignition voltage (VCC), the other input side is connected to the IO port of the single-chip microcomputer as the column control enable end (COL1), and the output side is connected to the ignition circuit composed of the diode and the ignition head.The right side is the row drive circuit, which is mainly composed of two N-MOSFETs.One input side is connected to the tail end of the ignition circuit, the other input side is connected to the IO port of the single chip microcomputer as the row control enable end (ROW1), and the output side is connected to the ground.Only when COL1 and ROW1 are controlled by single chip microcomputer to input high level at the same time, can the ignition head be heated and exploded, so as to complete the ignition task.
Due to the existence of the parasitic capacitance and resistance of the MOSFET, when the MOSFET is turned off, if a voltage is suddenly added to the source, the parasitic capacitance and resistance of the gate source will couple the voltage difference between the gate source and the source, resulting in the misconduction of the MOSFET, figure 4(a) shows the current waveform at the moment of power-on.In the project, this situation will lead to the wrong ignition of the thruster when the ignition voltage is energized.In order to prevent this situation, a soft start capacitor is connected in parallel between the gate and source of the MOSFET of the column drive circuit, that is, C1 in Figure 3.

Ignition control software system design
The ignition circuit and the drive circuit are the core components of the ignition control hardware system.Whether the attitude adjustment thruster can work normally also depends on the ignition circuit and the drive circuit.

Ignition control system program design
The program design part of the ignition control system is based on the serial port receiving and sending function of STM32F407.In order to prevent the disturbance in the data transmission process from causing the difference between the ignition signal analysis result and the expected result, the packet format of the ignition command transmission is agreed in advance.In order to prevent the closed quadrangle of the ignition circuit when the row and column control signal is output, the ignition delay rule is agreed, that is, the row and column control signal is closed immediately after the ignition of one thruster, the power supply of the lithium battery is cut off, and the ignition enable control of the other thruster is carried out after the cut off [8].The ignition process of the control program is shown in Figure 5.The spacecraft satellite computer calculates the attitude adjustment thruster that needs to be ignited according to the spacecraft attitude information, and packages the ignition information and sends it to the ignition core control board.After the control board receives the ignition instruction packet, it parses the ignition instruction packet to determine whether the packet is correct.If it is not correct, it sends the information requesting the satellite computer to send the ignition instruction again.If it is correct, according to the thruster to be ignited specified in the instruction, the control enables the row and column control terminals of the thruster to be specified.The lithium battery supplies power to the thruster igniter to successfully ignite the thruster and realize the adjustment of the failed spacecraft attitude.So as to improve the reliability of fast maneuvering deorbit of failed spacecraft.

Design of ignition control host computer
In order to simulate the spacecraft satellite platform to send ignition instructions to the ignition core control board, the windows form application is developed based on C #.The developed host computer interface is shown in Figure 6.The host computer realizes the communication with the STM32 core control board by setting the relevant parameters.The ignition control can be carried out by clicking the button or manual input, and the test data can be collected and displayed.

Ssoftware system testing
In order to verify the effectiveness of the software system design, the software system is tested.The software test mainly includes two parts: the first is whether the ignition core control module can receive the ignition information sent by the host computer, and the second is whether the correct ignition action can be made after receiving the ignition information.
By clicking the "11" button of the host computer, the command of the thruster ignition in the first row and the first column is issued, and the logic analyzer is used to monitor the ignition of the core control module to control the output level of the IO port.The detected waveform is shown in Figure 7(a).It can be seen from the figure that the IO port that can be controlled outputs a high level after receiving the ignition signal, that is, the thruster is ignited, and the ignition mark "Fire11" is also set to 1.The "1112": ignition signal is sent in the sending area of the host computer, which indicates that the two thrusters in the first column of the first row and the second column of the first row are ignited.The monitored waveform is shown in Figure 7(b).It can be seen from the figure that when the ignition signal is given, the two thrusters are ignited in turn, and the ignition signals "Fire11" and "Fire22" are also set to 1 when enabling the IO port to output high levels.It can be seen from the test results that the ignition core control module can correctly receive the ignition information sent by the host computer, and can correctly make the ignition action after receiving the ignition signal to meet the task requirements.

Conclusion
In this paper, the ignition control system of the attitude adjustment thruster of the failed spacecraft is studied.The array ignition circuit driven by MOS tube is adopted for the hardware system, and the ignition delay is adopted for the software system to prevent the ignition circuit from forming a closed quadrilateral, which leads to mis-ignition.The ignition control host computer is developed.Finally, the feasibility and reliability of the ignition control system are verified by tests, and the reliability of the deorbit of the failed spacecraft is improved.

Figure 1 .
Figure 1.Ignition control system work flow chart.

Figure 4 (
b) shows the current flowing through the ignition head at the moment of power on under different soft start capacitance values.It can be seen from Figure 4(b) that the larger the soft-start capacitance value, the better the effect of suppressing the conduction of the MOSFET at the moment of power-on.Considering the selection of 47 nF capacitor as the soft start capacitor.(a).Current at the moment of energization with no soft start capacitor.(b).Current at the moment of energization with different soft start capacitors.

Figure 4 .
Figure 4. Current at the moment of energization.