A design method and system for a spacecraft small debris impact monitoring system

Abstract

The present invention provides a design method and system for a spacecraft micro-debris impact monitoring system. A passive wireless sensor with a zigzag interdigital finger and four-reinforced grid structure is designed and fixed to the inner layer of the spacecraft bulkhead in an array form, so that electromagnetic waves can form secondary reflections when propagating inside the sensor. The reflection enhances the intensity of the electromagnetic waves and improves the sensitivity of the sensor. An impact positioning system is designed, and a laser reader is used to obtain changes in sensor impact signals to determine the strain of the bulkhead. A ground communication system is designed, and a ground laser reader is used to transmit laser beams for communication. The area size and position of the laser emitted by the laser reader correspond to the sensor array, and the system has the advantages of low cost, maintainability, and easy iterative upgrade. A digital flight companion system is designed, which realizes on-orbit health status monitoring and early warning of the spacecraft, realizes synchronous operation of the ground digital model and the actual spacecraft, and performs full-mission and full-time digital flight companionship.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft health monitoring, and in particular to a design method and system for a spacecraft micro-debris impact monitoring system. Background Art

[0002] Currently, space exploration is threatened by the increasing presence of space debris. Reducing the harm posed by space debris and ensuring the long-term sustainable development of humanity's space environment is of paramount importance to all countries worldwide. On-orbit monitoring and location of tiny space debris can effectively assess impact scenarios for spacecraft, enabling on-orbit health management and ensuring the health of spacecraft.

[0003] Currently, direct acoustic emission and pasting PVDF film onto the surface of spacecraft are mainly used to measure its strain. This method has the following defects:

[0004] In large spacecraft such as space stations, the elastic modulus of ordinary PVDF film and the adhesive after solidification far exceeds that of the measured bulkhead space, which restricts the expansion and contraction of the bulkhead space surface close to the PVDF film, resulting in local stress concentration. The test value cannot accurately reflect the actual situation.

[0005] The acoustic emission method is limited by the extravehicular environment and its principle. It can only measure hypervelocity impacts, has limited measurement accuracy, and can only measure flat plate structures.

[0006] Measurements using PVDF film and acoustic emission require a power supply, bridge, signal conditioning circuitry, and wired connections. The product is bulky in practice, and installation requires drilling holes in the spacecraft surface, making it difficult. Furthermore, the active method poses safety risks because the power supply must be installed and routed on the bulkhead.

[0007] PVDF film and acoustic emission sensors are not resistant to high temperatures. When a spacecraft is subjected to ablation due to high-speed impact of tiny debris, the surface temperature can reach 1300 degrees Celsius, making PVDF film and acoustic emission sensors ineffective.

[0008] At present, relevant passive wireless research for spacecraft has been carried out internationally. The research direction is mainly in the direction of sensor nodes, especially non-contact mechanical signal detection, and the main direction is LC sensors. However, internal energy collection and power supply are required, and the circuit will affect the surrounding electric field environment. In addition, the wireless transmission distance is short, and a larger impedance analyzer is required for signal acquisition. The actual use environment is obviously limited. Summary of the Invention

[0009] The purpose of the present invention is to overcome the deficiencies of the prior art and to provide a method and system for designing a space debris impact monitoring system for spacecraft.

[0010] The object of the present invention is achieved through the following technical solutions:

[0011] In a first aspect, the present invention provides a method for designing a space debris impact monitoring system for a spacecraft, comprising:

[0012] Step 1: Passive Wireless Sensor Design

[0013] Using multi-field coupling analysis software for calculations and simulations, a passive wireless sensor with a zigzag interdigital finger and four reinforcement grid structures was designed. The specific structure includes: a laser antenna, a reflective quad-mirror consisting of two primary and two secondary mirrors, four zigzag interdigital fingers connected to the laser antenna, and four reinforcement grids. Each zigzag interdigital finger has an angle of 120 degrees, and each reinforcement grid has two angles, each of which is 120 degrees.

[0014] Step 2: Passive Wireless Sensor Deployment

[0015] Passive wireless sensors are fixed to the inner layer of the spacecraft bulkhead in the form of an array. The passive wireless sensor array is installed according to the shape of the measured bulkhead. The spacecraft bulkhead is rectangular. The horizontal and vertical spacing of the passive wireless sensors are the same. When arranging the array, the sensor positions are calibrated to determine the position of each sensor.

[0016] Step 3: Design of impact monitoring system

[0017] When cracks and deformations occur in the spacecraft cabin area where the passive wireless sensor is located, the zigzag fingers and reinforcement grids deform, generating resonant wave signals carrying strain information. The resonant wave signals carrying strain information are acquired by a laser reader and processed using a convolutional neural network to determine the strain position and magnitude of the spacecraft cabin wall.

[0018] Step 4: Ground communication system design

[0019] Communication is achieved by emitting laser beams from a ground-based laser reader. The area and position of the laser emitted by the laser reader are identical to those of the passive wireless sensor array, and they correspond one-to-one. The laser emitted by the ground-based laser reader can penetrate the spacecraft cabin wall to reach the laser antenna, which then returns the backscattered signal generated by the passive wireless sensor.

[0020] Step 5: Design of digital flight companion system

[0021] The on-orbit data is transmitted to the spacecraft digital flight model of the ground flight control center. The digital flight system integrates historical health information, real-time health information, and predicted health information to monitor and warn of the spacecraft's on-orbit health status, formulate on-orbit operation strategies for the spacecraft, and achieve synchronous operation of the ground digital model and the actual spacecraft, conducting full-mission and full-time digital flight.

[0022] In a second aspect, the present invention provides a space micro-debris impact monitoring system for spacecraft, comprising: a passive wireless sensor array arranged on the inner layer of the spacecraft bulkhead, an impact monitoring system, a ground communication system, and a digital flight companion system arranged on the ground.

[0023] The passive wireless sensor array includes a plurality of passive wireless sensors installed laterally and vertically, with the lateral spacing between the passive wireless sensors being the same as the vertical spacing, and is used to generate a resonant wave signal carrying strain information when the spacecraft is hit by small debris;

[0024] The impact positioning system is used to obtain the resonant wave signal carrying strain information through a laser reader, and use a convolutional neural network to process the resonant wave signal carrying strain information to determine the strain position and strain magnitude;

[0025] The ground communication system is used to transmit a laser beam to the passive wireless sensor array through a laser reader and receive backscattered signals generated by each passive wireless sensor;

[0026] The digital companion flight system is used to obtain on-orbit data of the spacecraft as real-time health information of the spacecraft. By fusing historical health information, real-time health information, and predicted health information, it can monitor and warn the health status of the spacecraft on-orbit and formulate the spacecraft on-orbit operation strategy.

[0027] Optionally, the system further comprises a passive wireless sensor with an integrated temperature sensor arranged on the inner layer of the spacecraft bulkhead, for monitoring the temperature of the spacecraft bulkhead.

[0028] Optionally, the impact positioning system is further used to fit the strain distribution of the spacecraft bulkhead according to the strain position and strain magnitude of the spacecraft bulkhead, and draw a strain distribution diagram of the spacecraft bulkhead.

[0029] Furthermore, the system also includes a display component for performing three-dimensional visual display of the strain distribution diagram of the spacecraft cabin wall.

[0030] Compared with the prior art, the present invention has the following positive effects:

[0031] 1. The present invention specifically designs a passive wireless sensor and fixes it to the inner layer of the spacecraft cabin wall in the form of an array. The angle is designed to be 120 degrees so that the zigzag fingers and the reinforcement grid can form secondary reflections when the electromagnetic waves propagate inside the sensor. The reflections enhance the intensity of the electromagnetic waves and improve the sensitivity of the sensor.

[0032] 2. The present invention designs an impact positioning system, which uses a laser reader to obtain the sensor's reflected signal intensity (RSSI), phase, resonant frequency shift (RFS) and other characteristics as damage characteristics to determine the strain of the measured cabin and can measure cracks at the micron level.

[0033] 3. The present invention designs a ground communication system to monitor the health status of the spacecraft throughout its life cycle, including the launch phase, the on-orbit mission phase, and the return phase. Communication is achieved by emitting laser beams from a ground-based laser reader. The area size and position of the laser emitted by the laser reader are the same as those of the sensor array, and they correspond one-to-one. This method is low-cost, maintainable, and easy to iterate and upgrade.

[0034] 4. The present invention designs a digital flight companion system, which realizes the monitoring and early warning of the spacecraft's on-orbit health status, formulates the spacecraft's on-orbit operation strategy, realizes the synchronous operation of the ground digital model and the actual spacecraft, and conducts full-mission and full-time digital flight companion.

[0035] It should be understood that the contents described in the above summary of the invention are not intended to limit the key or important features of the embodiments of the present invention, nor are they intended to limit the scope of the present invention. Other features of the present invention will become readily understood through the following description. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following is a brief introduction to the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative labor.

[0037] Figure 1 A schematic diagram of the structure of a passive wireless sensor provided in an embodiment of the present disclosure;

[0038] Figure 2 Schematic diagram of strain distribution measured by passive wireless sensor;

[0039] Figure 3 A schematic diagram of the structure of a space debris impact monitoring system for spacecraft provided by an embodiment of the present disclosure;

[0040] Figure 4A schematic diagram of the structure of another space micro-debris impact monitoring system for spacecraft provided in an embodiment of the present disclosure. DETAILED DESCRIPTION

[0041] Embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. Although certain embodiments of the present invention are shown in the accompanying drawings, it should be understood that the present invention can be implemented in various forms and should not be construed as limited to the embodiments described herein. Rather, these embodiments are provided to provide a more thorough and complete understanding of the present invention. It should be understood that the drawings and embodiments of the present invention are for illustrative purposes only and are not intended to limit the scope of protection of the present invention.

[0042] The technical solution of the present invention is described in detail below with reference to specific embodiments. The following specific embodiments can be combined with each other, and the same or similar concepts or processes may not be described in detail in some embodiments.

[0043] The present disclosure provides a method for designing a space debris impact monitoring system for a spacecraft, including:

[0044] Step 1: Passive Wireless Sensor Design

[0045] Multi-field coupling analysis software is used for calculation and simulation to design a passive wireless sensor with a zigzag interdigital finger and four reinforced grid structures, such as Figure 1 As shown, the specific structure includes: a laser antenna, which is a reflective quad-mirror composed of two primary mirrors and two secondary mirrors, four zigzag fingers and four reinforcement grids connected to the laser antenna, each zigzag finger has an angle of 120 degrees, and each reinforcement grid has two angles, each angle is 120 degrees.

[0046] The sensor uses an ultra-high-temperature resistant ceramic material with a melting point above 1400°C as its substrate. The ceramic material is cleaned with deionized water and then evaporated in a constant temperature oven to remove the plasma water molecules on the substrate. A positive photoresist is sprayed on the surface of the ceramic material, which is then placed in a constant temperature oven for evaporation again. The photoresist is then placed in a photolithography machine for double-sided exposure and development, resulting in an etched pattern area. The ceramic material is then placed in a nitrogen temperature oven for preparation. In the etched pattern area, a metal platinum strain-sensitive layer is laser-written, followed by a layer of aluminum oxide. These two layers form a sensitive structure. A seven-layer sensitive structure is then built on the etched pattern area, and a high-temperature resistant material is added to form a passivation layer for bonding. This allows the sensor to reach a temperature sensitivity of 1300°C. To improve reliability, double-sided laser direct writing bonding is performed on the substrate. If one side of the structure fails due to the external environment, the entire sensor function remains intact. Finally, the laser mirror is adhered to the sensor body.

[0047] The sensor's millimeter-scale dimensions make it easy to install, regardless of mounting location or surface shape. The base is made of high-temperature-resistant ceramic, offering exceptional rigidity. The cured adhesive does not affect surface expansion and contraction, ensuring that test values ​​are unaffected by installation.

[0048] The angle is designed to be 120 degrees so that the zigzag fingers and the reinforcement grid can form secondary reflection when the electromagnetic wave propagates inside the sensor. The reflection enhances the intensity of the electromagnetic wave and improves the sensitivity of the sensor.

[0049] The sensor has a high quality factor Q, and a slight impact can generate a large signal output. The sensor was impact tested using an impact test bench. The sensor was installed on the test bench and the test bench was impacted at different speeds by setting the impact speed. The test results show that it can sense high-speed, low-speed and ultra-low-speed impacts.

[0050] Step 2: Passive Wireless Sensor Deployment

[0051] Passive wireless sensors are fixed to the inner layer of the spacecraft bulkhead in the form of an array. The passive wireless sensor array is installed according to the shape of the measured bulkhead. The spacecraft bulkhead is rectangular, and the horizontal and vertical spacing of the installation positions of each passive wireless sensor are the same. When arranging the array, the sensor position is calibrated to determine the position of each sensor.

[0052] Due to its high temperature resistance, it can be installed in high temperature areas. The passive wireless sensor array is installed according to the shape of the measured wall, such as Figure 2 As shown, the spacecraft cabin wall is rectangular, and the horizontal and vertical spacing of each sensor installation position are the same, that is, 0.1m. Figure 2 Where εxi=εyi=0.1m, the bulkhead is divided into several equal 0.01m 2 Area of ​​the region.

[0053] The sensors are arranged in a matrix array along the inner layer of the spacecraft cabin wall. Because the positions of the four reinforced grids of each passive wireless sensor are different, the reflected waves have their own phase information. When arranging the array, the sensor position can be obtained by measurement. After the array is arranged, the final sensor tag position can be obtained through calibration. All sensors in the array can be read and written using a reader.

[0054] Step 3: Design of impact monitoring system

[0055] When cracks and deformations occur in the spacecraft cabin area where the passive wireless sensor is located, the zigzag fingers and reinforcement grids are deformed, generating resonant wave signals carrying strain information. The resonant wave signals carrying strain information are obtained by a laser reader and processed using a convolutional neural network to determine the strain position and strain magnitude of the spacecraft cabin wall.

[0056] After the laser antenna on the sensor receives the laser emitted by the laser reader, it will generate a resonant wave signal with a frequency close to the laser on the zigzag interdigital fingers of the sensor's piezoelectric surface. The resonant wave signal will increase in signal intensity after being reflected by the reinforcement grids on both sides, thereby improving the Q value of the sensor. Impact will cause the zigzag interdigital fingers and reinforcement grids to change, which in turn causes changes in the signal strength (RSSI), phase, and resonant frequency shift (RFS) of the resonant wave signal. These changes are called strain information. The resonant wave signal carrying strain information is reflected back to the zigzag interdigital fingers and then returns to the laser reader through the laser antenna.

[0057] When strain acts on the sensor, it changes the speed S and wavelength L of the resonant wave. Strain causes the elastic constant, density and other parameters of the piezoelectric material to change. Changing the speed S causes the width of the zigzag interdigitated electrodes to be stretched or compressed, changing the wavelength L. In summary, the resonant frequency shift (RFS) is obtained according to the micromotion theory:

[0058]

[0059] Where S is the speed of the resonant wave, L is the wavelength of the resonant wave, ΔS is the change in the speed of the resonant wave, and ΔL is the change in the wavelength of the resonant wave.

[0060] According to the resonant frequency shift RFS of the laser signal received by the laser reader, ΔS is the speed change of the resonant wave, and ΔL is the wavelength change of the resonant wave. By inputting the parameters into the multi-field coupling analysis software, the strain can be obtained.

[0061] Step 4: Ground communication system design

[0062] Communication is achieved by emitting laser beams from a ground-based laser reader. The area size and position of the laser emitted by the laser reader are the same as those of the passive wireless sensor array, and they correspond one-to-one. The laser emitted by the ground-based laser reader can penetrate the spacecraft cabin wall to reach the laser antenna, and return the backscattered signal generated by the passive wireless sensor through the laser antenna.

[0063] In order to achieve full life cycle monitoring of the health status of the spacecraft during the launch phase, on-orbit mission phase and return phase, a ground-based laser reader is used to emit laser beams for communication. The area size and position of the laser emitted by the laser reader are the same as those of the sensor array, and they correspond one-to-one. This method is low-cost, maintainable and can be iteratively upgraded. The laser emitted by the ground-based laser reader can penetrate the cabin wall to reach the antenna, and return the backscattered signal from the sensor through the antenna.

[0064] Step 5: Design of digital flight companion system

[0065] The on-orbit data is transmitted to the spacecraft digital flight model of the ground flight control center. The digital flight system integrates historical health information, real-time health information, and predicted health information to monitor and warn of the spacecraft's on-orbit health status, formulate on-orbit operation strategies for the spacecraft, and achieve synchronous operation of the ground digital model and the actual spacecraft, conducting full-mission and full-time digital flight.

[0066] Figure 3 The embodiment of the present disclosure provides a space debris impact monitoring system for spacecraft, which is designed using the design method described in the above embodiment. Figure 3 As shown, the system includes: a passive wireless sensor array 1 arranged on the inner layer of the spacecraft bulkhead, a collision monitoring system 2, a ground communication system 3 and a digital flight companion system 4 arranged on the ground.

[0067] The passive wireless sensor array 1 comprises a plurality of passive wireless sensors installed laterally and vertically, with the lateral spacing between the passive wireless sensors being the same as the vertical spacing, and is used to generate a resonant wave signal carrying strain information when the spacecraft is hit by tiny debris;

[0068] The impact location system 2 is used to obtain the resonant wave signal carrying strain information through a laser reader, and use a convolutional neural network to process the resonant wave signal carrying strain information to determine the strain position and strain magnitude;

[0069] The ground communication system 3 is used to transmit a laser beam to the passive wireless sensor array through a laser reader and receive backscattered signals generated by each passive wireless sensor;

[0070] The digital companion flight system 4 is used to obtain the spacecraft's on-orbit data as the spacecraft's real-time health information. By fusing historical health information, real-time health information, and predicted health information, it can monitor and warn the spacecraft's on-orbit health status and formulate the spacecraft's on-orbit operation strategy.

[0071] In a possible implementation, the system further includes a passive wireless sensor with an integrated temperature sensor disposed on an inner layer of a spacecraft bulkhead, for monitoring the temperature of the spacecraft bulkhead.

[0072] In a possible implementation, the impact location system 2 is further configured to fit the strain distribution of the spacecraft bulkhead according to the strain position and strain magnitude of the spacecraft bulkhead, and draw a strain distribution diagram of the spacecraft bulkhead.

[0073] refer to Figure 2When the impact point in the figure is hit, the strain information carried by the echoes of the four nearby sensors (sensors 1 to 4) will change. The echo signals carrying the strain information are obtained by the laser reader and the convolutional neural network is used to process the echo signals carrying the strain information to determine the strain position and strain magnitude of the spacecraft cabin wall.

[0074] Along the X direction, the strain distribution along the X direction can be obtained by curve fitting based on the strain information of each point. Similarly, along the Y direction, the strain distribution along the Y direction can be obtained by curve fitting based on the strain information of each point. From this, the strain distribution map of the spacecraft cabin wall can be drawn.

[0075] refer to Figure 4 ,In a possible implementation manner, the system further includes a display component 5 for performing a three-dimensional visual display of the strain distribution diagram of the spacecraft cabin wall;

[0076] In this embodiment, in order to intuitively display the strain condition of the spacecraft bulkhead after being impacted, the three-dimensional strain distribution diagram is graphically displayed through the display component 5 based on the strain distribution diagram of the spacecraft bulkhead.

[0077] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the above embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the above embodiments, or replace some or all of the technical features therein with equivalents. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A design method for a spacecraft small debris impact monitoring system, characterized in that: include: Step 1: Passive Wireless Sensor Design Using multi-field coupling analysis software for calculations and simulations, a passive wireless sensor with a zigzag interdigital finger and four reinforcement grid structures was designed. The specific structure includes: a laser antenna, a reflective quad-mirror consisting of two primary and two secondary mirrors, four zigzag interdigital fingers connected to the laser antenna, and four reinforcement grids. Each zigzag interdigital finger has an angle of 120 degrees, and each reinforcement grid has two angles, each of which is 120 degrees. The sensor uses an ultra-high temperature resistant ceramic material with a melting point of over 1400°C as its substrate. The ceramic material is cleaned with deionized water and placed in a constant temperature oven for evaporation to remove the plasma water molecules on the substrate. Positive photoresist is sprayed on the surface of the ceramic material and placed in a constant temperature oven for evaporation again. The photoresist is then placed in a photolithography machine for double-sided exposure and development to obtain an etched pattern area. The ceramic material is then placed in a nitrogen temperature oven for preparation. In the etched pattern area, a layer of metal platinum strain sensitive layer is directly written by laser, followed by a layer of aluminum oxide. These two layers constitute a sensitive structure. A seven-layer sensitive structure is built on the etched pattern area, and then a high-temperature resistant material is added to form a passivation layer for bonding, so that the sensor's temperature sensitivity can reach 1300°C. In order to improve reliability, double-sided laser direct writing bonding is performed on the substrate. When one side of the structure fails due to the external environment, the entire sensor function does not fail. Finally, the laser mirror is adhered to the sensor body. Step 2: Passive Wireless Sensor Deployment Passive wireless sensors are fixed to the inner layer of the spacecraft bulkhead in the form of an array. The passive wireless sensor array is installed according to the shape of the measured bulkhead. The spacecraft bulkhead is rectangular. The horizontal and vertical spacing of the passive wireless sensors are the same. When arranging the array, the sensor positions are calibrated to determine the position of each sensor. Step 3: Design of impact monitoring system When cracks and deformations occur in the spacecraft cabin area where the passive wireless sensor is located, the zigzag fingers and reinforcement grids deform, generating resonant wave signals carrying strain information. The resonant wave signals carrying strain information are acquired by a laser reader and processed using a convolutional neural network to determine the strain position and magnitude of the spacecraft cabin wall. When strain acts on the sensor, it changes the speed S and wavelength L of the resonant wave. Strain causes the elastic constant, density and other parameters of the piezoelectric material to change. Changing the speed S will cause the width of the zigzag interdigitated electrode to be stretched or compressed, changing the wavelength L. In summary, according to the micromotion theory, the resonant frequency shift (RFS) is obtained: RFS = (1) = (2) = (3) = (4) ≈ (5) Where S is the speed of the resonant wave, L is the wavelength of the resonant wave, ΔS is the change in the speed of the resonant wave, and ΔL is the change in the wavelength of the resonant wave; According to the resonant frequency shift RFS of the laser signal received by the laser reader, ΔS is the speed change of the resonant wave, and ΔL is the wavelength change of the resonant wave. By inputting the parameters into the multi-field coupling analysis software, the strain can be obtained; Step 4: Ground communication system design Communication is achieved by emitting laser beams from a ground-based laser reader. The area and position of the laser emitted by the laser reader are identical to those of the passive wireless sensor array, and they correspond one-to-one. The laser emitted by the ground-based laser reader can penetrate the spacecraft cabin wall to reach the laser antenna, which then returns the backscattered signal generated by the passive wireless sensor. Step 5: Design of digital flight companion system The on-orbit data is transmitted to the spacecraft digital flight model of the ground flight control center. The digital flight system integrates historical health information, real-time health information, and predicted health information to monitor and warn of the spacecraft's on-orbit health status, formulate on-orbit operation strategies for the spacecraft, and achieve synchronous operation of the ground digital model and the actual spacecraft, conducting full-mission and full-time digital flight.

2. A space debris impact monitoring system for spacecraft, characterized in that: include: Passive wireless sensor arrays installed on the inner layer of the spacecraft bulkhead, impact monitoring systems, ground communication systems, and digital flight companion systems installed on the ground; The passive wireless sensor array includes a plurality of passive wireless sensors installed laterally and vertically, with the lateral spacing between the passive wireless sensors being the same as the vertical spacing, and is used to generate a resonant wave signal carrying strain information when the spacecraft is hit by small debris; The passive wireless sensor comprises a zigzag interdigital finger and a four-strengthened grid structure, specifically comprising: a laser antenna, the laser antenna comprising a reflective quad mirror consisting of a double primary mirror and a double secondary mirror, four zigzag interdigital fingers and four reinforcement grids connected to the laser antenna, each zigzag interdigital finger having an included angle of 120 degrees, and each reinforcement grid having two included angles, each of which is 120 degrees; The impact monitoring system is configured to acquire a resonant wave signal carrying strain information through a laser reader, process the resonant wave signal carrying strain information using a convolutional neural network to determine the strain position and magnitude, and fit the strain distribution of the spacecraft bulkhead based on the strain position and magnitude of the spacecraft bulkhead to draw a strain distribution map of the spacecraft bulkhead. The ground communication system is used to transmit a laser beam to the passive wireless sensor array through a laser reader and receive backscattered signals generated by each passive wireless sensor; The digital companion flight system is used to obtain on-orbit data of the spacecraft as real-time health information of the spacecraft. By fusing historical health information, real-time health information, and predicted health information, it can monitor and warn the health status of the spacecraft on-orbit and formulate the spacecraft on-orbit operation strategy.

3. The space debris impact monitoring system for spacecraft according to claim 2, characterized in that: The system further comprises a passive wireless sensor with an integrated temperature sensor arranged on an inner layer of the spacecraft bulkhead for monitoring the temperature of the spacecraft bulkhead.

4. The space debris impact monitoring system for spacecraft according to claim 2, characterized in that: The system further comprises a display component for performing three-dimensional visual display of the strain distribution diagram of the spacecraft cabin wall.

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