Modularized integrated spacecraft recycling comprehensive electronic system

Abstract

The invention relates to a modularized integrated spacecraft recycling comprehensive electronic system which comprises a power supply and distribution module, two firer management modules, a recycling interface module, a data acquisition and processing module and three FPGA control circuit modules. The modular and integrated design is carried out on each functional module, redundant parts are reduced, the circuit design is optimized, the structural shell is made of light materials, cable connection between single machines is reduced, and power supply, fire work, control and data acquisition are integrated. Under the fault mode of the FPGA control circuit module, the cold backup time controller circuit is switched and started, and normal work of the firer circuit can be guaranteed.

Description

Technical Field

[0001] This invention relates to a modular integrated electronic system for spacecraft recovery, belonging to the field of intelligent spacecraft recovery control technology. Background Technology

[0002] Domestic spacecraft reentry capsules utilize numerous recovery electronic systems. However, traditional recovery electronic systems typically employ a "bottom-up" development model, with individual units for power distribution, pyrotechnics, control, and data acquisition operating relatively independently, each performing its own function. This decentralized design of multiple functional modules results in numerous redundant components, independent structural shells, and separate external electrical connectors, adding significant amounts of extra materials. Furthermore, the cables connecting these independent units bear considerable weight, meaning that this decentralized recovery electronic system leads to a substantial waste of the reentry capsule's weight and volume.

[0003] With the widespread application of integrated electronic systems in spacecraft systems, recovery electronic systems have also undergone modular and integrated design. The heavy-load recovery control system, through unified requirements analysis, unified functional integration, and unified resource allocation, urgently needs to design a modular and integrated integrated electronic system for spacecraft recovery, making a significant contribution to the overall space and weight reduction of the spacecraft. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a modular and integrated electronic system for spacecraft recovery, which realizes the functional modularization and resource integration of the electronic system for spacecraft recovery, and significantly reduces weight and volume. The technical solution of this invention is: A modular integrated electronic system for spacecraft recovery includes: a power supply and distribution module, two pyrotechnics management modules, a recovery interface module, a data acquisition and processing module, and three FPGA control circuit modules; The power supply module receives the recovery main switch opening signal and the recovery power-on unlocking signal provided by the electronic information subsystem on the spacecraft. When the recovery main switch opening signal and the recovery power-on unlocking signal are both valid, the power bus voltage is connected to the recovery integrated electronic system and secondary power conversion is performed to supply power to the three FPGA control circuit modules, two pyrotechnic management modules, the recovery interface module, and the data acquisition and processing module. The data acquisition and processing module acquires several external analog signals, converts the analog signals into digital signals, and outputs them to three FPGA control circuit modules; The recovery interface module receives external switch signals and positive pulse signals, processes them, and forwards them to three FPGA control circuit modules. The FPGA control circuit modules determine the spacecraft's status based on the switch signals and positive pulse signals received from the recovery interface module and the digital signals output by the data acquisition and processing module, and send recovery timing commands to the recovery interface module. Each FPGA control circuit module includes a 422 interface circuit for information exchange with the electronic information subsystem on the spacecraft. After receiving the recovery timing instructions from the three FPGA control circuit modules, the recovery interface module performs a two-out-of-three vote and outputs the voted recovery timing instructions to the two pyrotechnic management modules; the pyrotechnic management modules receive the recovery timing instructions from the recovery interface module and drive the corresponding pyrotechnic devices to detonate. The recovery interface module is also used to enable the three FPGA control circuit modules to send health frequency signals to the recovery interface module when they send information to the outside world through the 422 interface, thereby setting the priority of the 422 data communication of the three FPGA control circuit modules; when at least two FPGA control circuit modules fail, the recovery interface module switches to start the cold backup time controller, and simultaneously enables the detonation of the corresponding pyrotechnic device in the FPGA control circuit failure mode; the recovery interface module is also used for forwarding the spacecraft's telemetry signals and electronic information subsystem commands.

[0005] Furthermore, the data acquisition and processing module collects several analog signals from sources including: three landing sensors, a static pressure altitude controller, and six airbag pressure sensors. The overload signal of the landing sensor is also converted into a separate switching signal, which serves as the input signal for the secondary startup of the cold backup time controller when the FPGA control circuit module is in fault mode; the FPGA control circuit module being in fault mode means that at least two FPGA control circuit modules have failed.

[0006] Furthermore, the FPGA control circuit module uses an FPGA chip as the main chip, with a 20MHz crystal oscillator. After initialization, the FPGA chip starts working after receiving a reset signal from an external reset circuit.

[0007] Furthermore, the spacecraft is equipped with two battery packs, each corresponding to a pyrotechnic management module. The two pyrotechnic management modules are completely identical, both outputting pyrotechnic device detonation signals and are redundant with each other.

[0008] Furthermore, the recycling interface module includes: a switch signal processing circuit, a positive pulse signal processing circuit, a communication on-duty machine status discrimination circuit, a cold backup primary start circuit, a cold backup secondary start circuit, a three-out-of-two voting output circuit, a telemetry signal circuit, an instruction forwarding circuit, and a detonation pyrotechnic device circuit. After receiving the externally input switch signal, the switch signal processing circuit sets the level validity and sends the processed switch signal to three FPGA control circuit modules and a cold backup start-up circuit. After receiving the positive pulse signal from the external input, the positive pulse signal processing circuit holds the positive pulse signal and converts it into a switching quantity before sending it to the three FPGA control circuit modules and the cold backup start-up circuit. The two-out-of-three voting output circuit, upon receiving the recovery timing instructions from the three FPGA control circuit modules, performs a two-out-of-three vote and sends the result to the two pyrotechnic management modules. Simultaneously, the result of the two-out-of-three vote is transmitted to the ground via the telemetry signal circuit, and then forwarded to the electronic subsystem on the aircraft via the command forwarding circuit. The communication duty status determination circuit collects the health frequency signals of the three FPGA control circuit modules to determine the 422 communication duty, that is, to determine the priority of the data output by the three FPGA control circuit modules through the 422 interface; The cold backup one-time start circuit determines whether to start the cold backup time controller based on the received switch signal and positive pulse signal when the three FPGA control circuit modules are in fault mode. When the six conditions are met, the cold backup one-time start circuit sends a time controller one-time start signal to control the external cold backup time controller to start once. The cold backup time controller generates a timing instruction after one-time start and sends it to the detonation pyrotechnic device circuit. The cold backup secondary start circuit sends a time controller secondary start signal after receiving the time controller primary start signal and the landing sensor overload switch signal, thereby realizing the secondary start of the cold backup time controller. It is used to generate timing instructions after the secondary start of the time controller and send them to the detonation pyrotechnic device circuit. The detonation pyrotechnic device circuit enables itself based on the received first or second activation timing command, and detonates the corresponding pyrotechnic device according to the first or second activation timing command.

[0009] Furthermore, the six conditional criteria are as follows: 1) The aircraft is in high-altitude rescue mode; 2) Limit switch a is not open; 3) Limit switch b is not open; 4) Limit switch C is not open; 5) Diaphragm-type altitude controller with 10km range; 6) Static pressure height controller operates at 6km; The limit switches a, b, and c are mechanical switches installed on the spacecraft. Limit switches a and b are used for the spacecraft parachute detachment signal. When the parachute detaches, limit switches a and b are connected. When the parachute does not detach, limit switches a and b are disconnected. The initial state of limit switches a and b is disconnected. Limit switch C is used as a signal for the separation mechanism hatch cover to detach. When the separation mechanism hatch cover detaches, limit switch C is turned on, and when it is not detached, limit switch C is turned off. The initial state of limit switch C is turned off. Both the diaphragm-type altitude controller and the hydrostatic altitude controller are equipment on spacecraft. The diaphragm-type altitude controller is used for 10km altitude information acquisition, while the hydrostatic altitude controller is used for 6km altitude information acquisition. The 10km altitude output signal of the diaphragm-type altitude controller is a switching signal, where a value of 1 represents on and a value of 0 represents off. The 6km altitude output signal of the hydrostatic altitude controller is a continuous analog signal, which is converted into a switching signal, where a value of 1 represents on and a value of 0 represents off.

[0010] Furthermore, the cold backup one-time start circuit includes optocoupler isolation circuits 1-6, transistor drive circuits 1-4, optocoupler self-holding circuits 1-3, optocoupler isolation and three-out-of-two circuits, and time controller power supply input circuit; When the spacecraft is in high-altitude rescue mode, it receives status 1 and status 2 signals from the integrated business unit of the return capsule sealed compartment; when both status 1 and status 2 signals are 1, the return capsule sealed compartment is in high-altitude rescue mode; when either status 1 or status 2 signal is 1, it is in mid-altitude rescue mode; when both status 1 and status 2 signals are 0, it is in low-altitude rescue mode. The spacecraft's 28V time controller power supply is connected to the time controller power supply input circuit, and the output of the time controller power supply input circuit serves as the power input of the optocoupler isolation circuit 3; the limit switch a signal is amplified by the transistor drive circuit 1 and then sent to the optocoupler isolation circuit 3 and the optocoupler self-holding circuit 1, and the output signal of the optocoupler isolation circuit 3 serves as the power input of the optocoupler isolation circuit 4. The signal from limit switch b is amplified by transistor drive circuit 2 and then sent to optocoupler isolation circuit 4 and optocoupler self-holding circuit 2. The output signal of optocoupler isolation circuit 4 serves as the power input of optocoupler isolation circuit 5. The limit switch c signal is amplified by the transistor drive circuit 3 and then sent to the optocoupler isolation circuit 5 and the optocoupler self-holding circuit 3. The output signal of the optocoupler isolation circuit 5 is used as the power input of the optocoupler isolation circuit 6. The 10km switching signal is amplified by the transistor driver circuit 4 and then sent to the optocoupler isolation circuit 6. The output signal of the optocoupler isolation circuit 6 serves as the power input of the optocoupler isolation circuit 1. The state 1 signal is sent to the optocoupler isolation circuit 1, and the output signal of the optocoupler isolation circuit 1 serves as the power input of the optocoupler isolation circuit 2. The state 2 signal is sent to the optocoupler isolation circuit 2, and the output signal of the optocoupler isolation circuit 2 is used as the power input for the optocoupler isolation and three-out-of-two circuit. The 6KM-A signal, 6KM-B signal, and 6KM-C signal are sent to the optocoupler isolation and three-out-of-two circuit. After the three-out-of-two voting, the 28V time controller power supply is finally sent to the cold backup time controller as the time controller's first start signal. Since limit switches a, b, and c will activate according to the timing sequence, the time controller sends a start signal and feeds it back to optocoupler self-holding circuit 1, optocoupler self-holding circuit 2, and optocoupler self-holding circuit 3 to ensure the continuity of the circuit.

[0011] Furthermore, the integrated electronic recycling system adopts an integrated structural design, which includes a housing, a cover plate, a base plate, connecting strips, and a mounting frame. The upper and lower ends of the magnesium alloy shell are sealed and connected by a cover plate and a bottom plate. The power supply and distribution module, two pyrotechnic management modules, recycling interface module, data acquisition and processing module, and three FPGA control circuit modules are all fixed inside the shell by mounting frames. The mounting frames are provided with grooves, and all mounting frames are electrically connected and fixed by connecting strips.

[0012] The beneficial effects of this invention compared to the prior art are: (1) This invention uses modular and integrated design of various functional modules to reduce redundant components, optimize circuit design, use lightweight materials for the structural shell, reduce cable connections between individual units, integrate power supply, pyrotechnics, control and data acquisition, and the whole machine weighs 13kg, which is significantly lighter than traditional electronic recovery systems, and provides a favorable value for subsequent large-payload spacecraft recovery control systems.

[0013] (2) In the fault mode of the FPGA control circuit module in this invention, the cold backup time controller circuit is switched on to ensure the normal operation of the pyrotechnic circuit. The circuit uses optocouplers to isolate the primary and secondary power supplies, and the time controller is started once by connecting the above 6 conditions in series. When the spacecraft is in high-altitude rescue mode, it receives status 1 and status 2 signals from the integrated business unit 1 of the return capsule sealed compartment. After circuit conversion, these signals are transmitted to the circuit. These signals are active at a low level of 28V and are sent to optocoupler isolation circuit 1 and optocoupler isolation circuit 2. When limit switches a, b, and c are closed, these three signals are active at a low level of 28V and are connected to optocoupler isolation circuit 3, 4, and 5, respectively. Since these three signals will be turned on after the time controller sends the timing command, to maintain the state of the three limit switch signals, the state of the three limit switches is fed back through the cold backup time controller start signal, forming a self-holding circuit. These three signals are connected to optocoupler self-holding circuit 1, 2, and 3, respectively. The 10km signal is active at a high level of 28V and is connected to optocoupler isolation circuit 4. The 6km signal is a three-way signal. It is first processed by optocoupler isolation and a 3-out-of-2 circuit before being connected in series with the previous six circuits, ultimately realizing the start of the cold backup time controller. Attached Figure Description

[0014] Figure 1 This is a schematic diagram of the integrated electronic system architecture for recycling according to the present invention; Figure 2 This is an internal block diagram of the recycling interface module of the present invention; Figure 3 This is a schematic diagram of the connection methods between layers; Figure 4 This is a schematic diagram showing the connection method at the top and bottom; Figure 5 Block diagram of the startup circuit for the cold backup time controller; Figure 6 Installation frame structure diagram; Figure 7 This is a structural diagram of the connecting strip. Detailed Implementation

[0015] The specific embodiments of the present invention will now be described in further detail with reference to the accompanying drawings.

[0016] like Figure 1 As shown, the present invention provides a modular integrated electronic system for spacecraft recovery, including: a power supply and distribution module, two pyrotechnics management modules, a recovery interface module, a data acquisition and processing module, and three FPGA control circuit modules; The power supply module receives the recovery main switch opening signal and the recovery power-on unlocking signal provided by the electronic information subsystem on the spacecraft. When the recovery main switch opening signal and the recovery power-on unlocking signal are both valid, the power bus voltage is connected to the recovery integrated electronic system and secondary power conversion is performed to supply power to the three FPGA control circuit modules, two pyrotechnic management modules, the recovery interface module, and the data acquisition and processing module. The data acquisition and processing module acquires several external analog signals, converts the analog signals into digital signals, and outputs them to three FPGA control circuit modules; The recovery interface module receives external switch signals and positive pulse signals, processes them, and forwards them to three FPGA control circuit modules. The FPGA control circuit modules determine the spacecraft's status based on the switch signals and positive pulse signals received from the recovery interface module and the digital signals output by the data acquisition and processing module, and send recovery timing commands to the recovery interface module. Each FPGA control circuit module includes a 422 interface circuit for information exchange with the electronic information subsystem on the spacecraft. After receiving the recovery timing instructions from the three FPGA control circuit modules, the recovery interface module performs a two-out-of-three vote and outputs the voted recovery timing instructions to the two pyrotechnic management modules; the pyrotechnic management modules receive the recovery timing instructions from the recovery interface module and drive the corresponding pyrotechnic devices to detonate. The recovery interface module is also used to enable the three FPGA control circuit modules to send health frequency signals to the recovery interface module when they send information to the outside world through the 422 interface, thereby setting the priority of the 422 data communication of the three FPGA control circuit modules; when at least two FPGA control circuit modules fail, the recovery interface module switches to start the cold backup time controller, and simultaneously enables the detonation of the corresponding pyrotechnic device in the FPGA control circuit failure mode; the recovery interface module is also used for forwarding the spacecraft's telemetry signals and electronic information subsystem commands.

[0017] The data acquisition and processing module collects several analog signals from sources including: three landing sensors, a static pressure altitude controller, and six airbag pressure sensors. The overload signal of the landing sensor is also converted into a switch signal separately, which serves as the input signal for the secondary startup of the cold backup time controller when the FPGA control circuit module is in fault mode. The FPGA control circuit module being in fault mode means that at least two FPGA control circuit modules have failed.

[0018] Preferably, the FPGA control circuit module uses an FPGA chip as the main chip, and a 20MHz crystal oscillator is used. After the FPGA chip is initialized, it starts working after receiving the reset signal from the external reset circuit.

[0019] Preferably, the spacecraft is equipped with two battery packs, each corresponding to a pyrotechnic management module. The two pyrotechnic management modules are completely identical and both output pyrotechnic device detonation signals, which are redundant with each other.

[0020] like Figure 2 As shown, the recovery interface module includes: a switch signal processing circuit, a positive pulse signal processing circuit, a communication on-duty machine status discrimination circuit, a cold backup primary start circuit, a cold backup secondary start circuit, a three-out-of-two voting output circuit, a telemetry signal circuit, an instruction forwarding circuit, and a detonation pyrotechnic device circuit. After receiving the externally input switch signal, the switch signal processing circuit sets the level validity and sends the processed switch signal to three FPGA control circuit modules and a cold backup start-up circuit. After receiving the positive pulse signal from the external input, the positive pulse signal processing circuit holds the positive pulse signal and converts it into a switching quantity before sending it to the three FPGA control circuit modules and the cold backup start-up circuit. The two-out-of-three voting output circuit, upon receiving the recovery timing instructions from the three FPGA control circuit modules, performs a two-out-of-three vote and sends the result to the two pyrotechnic management modules. Simultaneously, the result of the two-out-of-three vote is transmitted to the ground via the telemetry signal circuit, and then forwarded to the electronic subsystem on the aircraft via the command forwarding circuit. The communication duty status determination circuit collects the health frequency signals of the three FPGA control circuit modules to determine the 422 communication duty, that is, to determine the priority of the data output by the three FPGA control circuit modules through the 422 interface; The cold backup one-time start circuit determines whether to start the cold backup time controller based on the received switch signal and positive pulse signal when the three FPGA control circuit modules are in fault mode. When the six conditions are met, the cold backup one-time start circuit sends a time controller one-time start signal to control the external cold backup time controller to start once. The cold backup time controller generates a timing instruction after one-time start and sends it to the detonation pyrotechnic device circuit. The cold backup secondary start circuit sends a time controller secondary start signal after receiving the time controller primary start signal and the landing sensor overload switch signal, thereby realizing the secondary start of the cold backup time controller. It is used to generate timing instructions after the secondary start of the time controller and send them to the detonation pyrotechnic device circuit. The detonation pyrotechnic device circuit enables itself based on the received first or second activation timing command, and detonates the corresponding pyrotechnic device according to the first or second activation timing command.

[0021] The key feature of this invention is that when at least two FPGA control circuit modules fail, referred to as a fault mode, the cold backup time controller function is switched on. The startup time controller requires the following six conditions to be met simultaneously, i.e., the six-condition criterion: 1) The aircraft is in high-altitude rescue mode; 2) Limit switch a is not open; 3) Limit switch b is not open; 4) Limit switch C is not open; 5) Diaphragm-type altitude controller with 10km range; 6) Static pressure height controller operates at 6km; The limit switches a, b, and c are mechanical switches installed on the spacecraft. Limit switches a and b are used for the spacecraft parachute detachment signal. When the parachute detaches, limit switches a and b are connected. When the parachute does not detach, limit switches a and b are disconnected. The initial state of limit switches a and b is disconnected. Limit switch C is used as a signal for the separation mechanism hatch cover to detach. When the separation mechanism hatch cover detaches, limit switch C is turned on, and when it is not detached, limit switch C is turned off. The initial state of limit switch C is turned off. Both the diaphragm-type altitude controller and the hydrostatic altitude controller are equipment on spacecraft. The diaphragm-type altitude controller is used for 10km altitude information acquisition, while the hydrostatic altitude controller is used for 6km altitude information acquisition. The 10km altitude output signal of the diaphragm-type altitude controller is a switching signal, where a value of 1 represents on and a value of 0 represents off. The 6km altitude output signal of the hydrostatic altitude controller is a continuous analog signal, which is converted into a switching signal, where a value of 1 represents on and a value of 0 represents off.

[0022] like Figure 5 As shown, the cold backup one-time start circuit includes optocoupler isolation circuits 1~6, transistor drive circuits 1~4, optocoupler self-holding circuits 1~3, optocoupler isolation and three-out-of-two circuits, and time controller power supply input circuit; When the spacecraft is in high-altitude rescue mode, it receives status 1 and status 2 signals from the integrated business unit of the return capsule sealed compartment; when both status 1 and status 2 signals are 1, the return capsule sealed compartment is in high-altitude rescue mode; when either status 1 or status 2 signal is 1, it is in mid-altitude rescue mode; when both status 1 and status 2 signals are 0, it is in low-altitude rescue mode. The spacecraft's 28V time controller power supply is connected to the time controller power supply input circuit, and the output of the time controller power supply input circuit serves as the power input of the optocoupler isolation circuit 3; the limit switch a signal is amplified by the transistor drive circuit 1 and then sent to the optocoupler isolation circuit 3 and the optocoupler self-holding circuit 1, and the output signal of the optocoupler isolation circuit 3 serves as the power input of the optocoupler isolation circuit 4. The signal from limit switch b is amplified by transistor drive circuit 2 and then sent to optocoupler isolation circuit 4 and optocoupler self-holding circuit 2. The output signal of optocoupler isolation circuit 4 serves as the power input of optocoupler isolation circuit 5. The limit switch c signal is amplified by the transistor drive circuit 3 and then sent to the optocoupler isolation circuit 5 and the optocoupler self-holding circuit 3. The output signal of the optocoupler isolation circuit 5 is used as the power input of the optocoupler isolation circuit 6. The 10km switching signal is amplified by the transistor driver circuit 4 and then sent to the optocoupler isolation circuit 6. The output signal of the optocoupler isolation circuit 6 serves as the power input of the optocoupler isolation circuit 1. The state 1 signal is sent to the optocoupler isolation circuit 1, and the output signal of the optocoupler isolation circuit 1 serves as the power input of the optocoupler isolation circuit 2. The state 2 signal is sent to the optocoupler isolation circuit 2, and the output signal of the optocoupler isolation circuit 2 is used as the power input for the optocoupler isolation and three-out-of-two circuit. The 6KM-A signal, 6KM-B signal, and 6KM-C signal are sent to the optocoupler isolation and three-out-of-two circuit. After the three-out-of-two voting, the 28V time controller power supply is finally sent to the cold backup time controller as the time controller's first start signal. Since limit switches a, b, and c will activate according to the timing sequence, the time controller sends a start signal and feeds it back to optocoupler self-holding circuit 1, optocoupler self-holding circuit 2, and optocoupler self-holding circuit 3 to ensure the continuity of the circuit.

[0023] The invention is characterized by a modular and integrated spacecraft recovery electronic system that adopts a modular, universal, and standardized design philosophy. Multiple functional modules are planned from subsystems, and different functional modules can be combined to form different stand-alone products. The modules are combined through stacking multiple internal connectors.

[0024] The integrated electronic recycling system adopts an integrated structural design, which includes a housing, cover plate, base plate, connecting strips, and mounting frame. The magnesium alloy housing is sealed at both ends via the cover plate and base plate. The power supply module, two pyrotechnics management modules, recycling interface module, data acquisition and processing module, and three FPGA control circuit modules are all fixed inside the housing via the mounting frame. (See attached image for mounting frame details.) Figure 6 As shown, the mounting frame has grooves, and all mounting frames are electrically connected and fixed via connecting strips. (See connecting strips...) Figure 7 As shown. The outer surface of the casing is anodized in black, and the contact areas between the plates are electrically oxidized. A total of 16 lug mounting surfaces are electrically oxidized, as shown. Figure 3 and Figure 4 As shown.

[0025] The shell material of this invention is made of magnesium alloy, which has a density of 1.8 g / cm³, only 2 / 3 that of aluminum alloy (2.7 g / cm³). Based on the tensile strength of magnesium alloy plates according to GBn250-1985, aluminum alloy of grade MB15-S is used, which has a tensile strength of up to 265 MPa.

[0026] To prevent the gap from widening and expanding due to vibration at the bottom of the product during operation, and to ensure a good connection between the entire casing, a connecting strip was designed. See the structural diagram for details. Figure 7 .

[0027] This invention proposes a modular and integrated electronic system for spacecraft recovery. Through the integrated design of various functional modules, redundant components are reduced, circuit design is optimized, and the structural shell uses lightweight materials to minimize cable connections between individual units. It integrates power supply, pyrotechnics, control, and data acquisition into a single unit, with a total weight of 13 kg. In contrast, traditional recovery electronic systems typically have power distribution equipment weighing approximately 4 kg, control equipment approximately 3.6 kg, pyrotechnics control equipment approximately 6.4 kg, data acquisition equipment approximately 2 kg, and cable connections between individual units and the ship approximately 20 kg, totaling 36 kg. The integrated electronic system designed in this study reduces the weight of traditional recovery electronic systems by approximately 23 kg, providing valuable insights for future heavy-payload spacecraft recovery control systems.

[0028] The parts of this invention not described in detail are common knowledge to those skilled in the art.

Claims

1. A modular integrated electronic system for spacecraft recovery, characterized in that, include: The system includes a power supply and distribution module, two pyrotechnics management modules, a recycling interface module, a data acquisition and processing module, and three FPGA control circuit modules. The power supply module receives the recovery main switch opening signal and the recovery power-on unlocking signal provided by the electronic information subsystem on the spacecraft. When the recovery main switch opening signal and the recovery power-on unlocking signal are both valid, the power bus voltage is connected to the recovery integrated electronic system and secondary power conversion is performed to supply power to the three FPGA control circuit modules, two pyrotechnic management modules, the recovery interface module, and the data acquisition and processing module. The data acquisition and processing module acquires several external analog signals, converts the analog signals into digital signals, and outputs them to three FPGA control circuit modules. The recovery interface module receives external switch signals and positive pulse signals, processes them, and forwards them to three FPGA control circuit modules. The FPGA control circuit modules determine the spacecraft's status based on the switch signals and positive pulse signals received from the recovery interface module and the digital signals output by the data acquisition and processing module, and send recovery timing commands to the recovery interface module. Each FPGA control circuit module includes a 422 interface circuit for information exchange with the electronic information subsystem on the spacecraft. After receiving the recovery timing instructions from the three FPGA control circuit modules, the recovery interface module performs a two-out-of-three vote and outputs the voted recovery timing instructions to the two pyrotechnic management modules; the pyrotechnic management modules receive the recovery timing instructions from the recovery interface module and drive the corresponding pyrotechnic devices to detonate. The recovery interface module is also used to enable the three FPGA control circuit modules to send health frequency signals to the recovery interface module when they send information to the outside world through the 422 interface, thereby setting the priority of the 422 data communication of the three FPGA control circuit modules; when at least two FPGA control circuit modules fail, the recovery interface module switches to start the cold backup time controller, and simultaneously enables the detonation of the corresponding pyrotechnic device in the FPGA control circuit failure mode; the recovery interface module is also used for forwarding the spacecraft's telemetry signals and electronic information subsystem commands.

2. The modular integrated spacecraft recovery electronic system according to claim 1, characterized in that: The data acquisition and processing module collects several analog signals from the following sources: three landing sensors, a static pressure altitude controller, and six airbag pressure sensors. The overload signal of the landing sensor is also converted into a separate switching signal, which serves as the input signal for the secondary startup of the cold backup time controller when the FPGA control circuit module is in fault mode; the FPGA control circuit module being in fault mode means that at least two FPGA control circuit modules have failed.

3. The modular integrated spacecraft recovery electronic system according to claim 1, characterized in that: The FPGA control circuit module uses an FPGA chip as the main chip and a 20MHz crystal oscillator. After initialization, the FPGA chip starts working after receiving a reset signal from an external reset circuit.

4. The modular integrated spacecraft recovery electronic system according to claim 1, characterized in that: The spacecraft is equipped with two battery packs, each corresponding to a pyrotechnic management module. The two pyrotechnic management modules are completely identical and both output pyrotechnic device detonation signals, which are redundant with each other.

5. The modular integrated spacecraft recovery electronic system according to claim 1, characterized in that: The recovery interface module includes: a switch signal processing circuit, a positive pulse signal processing circuit, a communication on-duty machine status discrimination circuit, a cold backup primary start circuit, a cold backup secondary start circuit, a three-out-of-two voting output circuit, a telemetry signal circuit, an instruction forwarding circuit, and a detonation pyrotechnic device circuit. After receiving the externally input switch signal, the switch signal processing circuit sets the level validity and sends the processed switch signal to three FPGA control circuit modules and a cold backup start-up circuit. After receiving the positive pulse signal from the external input, the positive pulse signal processing circuit holds the positive pulse signal and converts it into a switching quantity before sending it to the three FPGA control circuit modules and the cold backup start-up circuit. The two-out-of-three voting output circuit, upon receiving the recovery timing instructions from the three FPGA control circuit modules, performs a two-out-of-three vote and sends the result to the two pyrotechnic management modules. Simultaneously, the result of the two-out-of-three vote is transmitted to the ground via the telemetry signal circuit, and then forwarded to the electronic subsystem on the aircraft via the command forwarding circuit. The communication duty status determination circuit collects the health frequency signals of the three FPGA control circuit modules to determine the 422 communication duty, that is, to determine the priority of the data output by the three FPGA control circuit modules through the 422 interface; The cold backup one-time start circuit determines whether to start the cold backup time controller based on the received switch signal and positive pulse signal when the three FPGA control circuit modules are in fault mode. When the six conditions are met, the cold backup one-time start circuit sends a time controller one-time start signal to control the external cold backup time controller to start once. The cold backup time controller generates a timing instruction after one-time start and sends it to the detonation pyrotechnic device circuit. The cold backup secondary start circuit sends a time controller secondary start signal after receiving the time controller primary start signal and the landing sensor overload switch signal, thereby realizing the secondary start of the cold backup time controller. It is used to generate timing instructions after the secondary start of the time controller and send them to the detonation pyrotechnic device circuit. The detonation pyrotechnic device circuit enables itself based on the received first or second activation timing command, and detonates the corresponding pyrotechnic device according to the first or second activation timing command.

6. The modular integrated spacecraft recovery electronic system according to claim 5, characterized in that: The six conditional criteria are as follows: 1) The aircraft is in high-altitude rescue mode; 2) Limit switch a is not open; 3) Limit switch b is not open; 4) Limit switch C is not open; 5) Diaphragm-type altitude controller with 10km range; 6) Static pressure height controller operates at 6km; The limit switches a, b, and c are mechanical switches installed on the spacecraft. Limit switches a and b are used for the spacecraft parachute detachment signal. When the parachute detaches, limit switches a and b are connected. When the parachute does not detach, limit switches a and b are disconnected. The initial state of limit switches a and b is disconnected. Limit switch C is used as a signal for the separation mechanism hatch cover to detach. When the separation mechanism hatch cover detaches, limit switch C is turned on, and when it is not detached, limit switch C is turned off. The initial state of limit switch C is turned off. Both the diaphragm-type altitude controller and the hydrostatic altitude controller are equipment on spacecraft. The diaphragm-type altitude controller is used for 10km altitude information acquisition, while the hydrostatic altitude controller is used for 6km altitude information acquisition. The 10km altitude output signal of the diaphragm-type altitude controller is a switching signal, where a value of 1 represents on and a value of 0 represents off. The 6km altitude output signal of the hydrostatic altitude controller is a continuous analog signal, which is converted into a switching signal, where a value of 1 represents on and a value of 0 represents off.

7. A modular integrated spacecraft recovery electronic system according to claim 6, characterized in that: The cold backup one-time start circuit includes optocoupler isolation circuits 1-6, transistor drive circuits 1-4, optocoupler self-holding circuits 1-3, optocoupler isolation and three-out-of-two circuits, and time controller power supply input circuit; When the spacecraft is in high-altitude rescue mode, it receives status 1 and status 2 signals from the integrated operational unit of the return capsule sealed compartment. The spacecraft's 28V time controller power supply is connected to the time controller power supply input circuit, and the output of the time controller power supply input circuit serves as the power input of the optocoupler isolation circuit 3; the limit switch a signal is amplified by the transistor drive circuit 1 and then sent to the optocoupler isolation circuit 3 and the optocoupler self-holding circuit 1, and the output signal of the optocoupler isolation circuit 3 serves as the power input of the optocoupler isolation circuit 4. The signal from limit switch b is amplified by transistor drive circuit 2 and then sent to optocoupler isolation circuit 4 and optocoupler self-holding circuit 2. The output signal of optocoupler isolation circuit 4 is used as the power input of optocoupler isolation circuit 5. The limit switch c signal is amplified by the transistor drive circuit 3 and then sent to the optocoupler isolation circuit 5 and the optocoupler self-holding circuit 3. The output signal of the optocoupler isolation circuit 5 is used as the power input of the optocoupler isolation circuit 6. The 10km switching signal is amplified by the transistor driver circuit 4 and then sent to the optocoupler isolation circuit 6. The output signal of the optocoupler isolation circuit 6 serves as the power input of the optocoupler isolation circuit 1. The state 1 signal is sent to the optocoupler isolation circuit 1, and the output signal of the optocoupler isolation circuit 1 serves as the power input of the optocoupler isolation circuit 2. The state 2 signal is sent to the optocoupler isolation circuit 2, and the output signal of the optocoupler isolation circuit 2 is used as the power input for the optocoupler isolation and three-out-of-two circuit. The 6KM-A signal, 6KM-B signal, and 6KM-C signal are sent to the optocoupler isolation and three-out-of-two circuit. After the three-out-of-two voting, the 28V time controller power supply is finally sent to the cold backup time controller as the time controller's first start signal. Since limit switches a, b, and c will activate according to the timing sequence, the time controller sends a start signal and feeds it back to optocoupler self-holding circuit 1, optocoupler self-holding circuit 2, and optocoupler self-holding circuit 3 to ensure the continuity of the circuit.

8. A modular integrated spacecraft recovery electronic system according to any one of claims 1-7, characterized in that: The integrated electronic recycling system adopts an integrated structural design, which includes a housing, cover plate, base plate, connecting strips, and mounting frame. The upper and lower ends of the magnesium alloy shell are sealed and connected by a cover plate and a bottom plate. The power supply and distribution module, two pyrotechnic management modules, recycling interface module, data acquisition and processing module, and three FPGA control circuit modules are all fixed inside the shell by a mounting frame. The mounting frame is provided with grooves, and all mounting frames are electrically connected and fixed by connecting strips.

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