Spacecraft structure impact source positioning system

Abstract

The invention is suitable for the technical field of spacecrafts, and provides a spacecraft structure collision source positioning system, which comprises a detection system, a collision early warning system and a processing system, the detection system is used for obtaining and transmitting state information of the surface of the spacecraft and comprises a sensor module and a data acquisition module. The collision early warning system is used for preventing the spacecraft from colliding with other space objects, real-time simulation and trajectory prediction are carried out on the current operation state of the spacecraft, and whether the current operation state of the spacecraft needs to be changed or not is conveniently and visually known and judged; the processing system is used for processing collision caused by the fact that the collision early warning system cannot avoid collision or early warning fails, and the device solves the problems that a spacecraft cannot avoid space debris collision with the small size and the high speed, and an accurate collision position is difficult to obtain after collision. The purposes that as few PVDF sensors are adopted as possible, it is guaranteed that impact points are accurately positioned, and impact force is obtained are achieved.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft technology, and more particularly to a spacecraft structure impact source positioning system. Background Art

[0002] Spacecraft refers to various types of aircraft that operate in space outside the Earth's atmosphere, basically in accordance with the laws of celestial mechanics. Spacecraft can be divided into two categories: unmanned spacecraft and manned spacecraft. Unmanned spacecraft include artificial satellites, probes, etc., which are used for scientific research, communication, navigation, meteorological observation, earth resource survey and other tasks.

[0003] In the current aerospace field, artificial satellites are the most common spacecraft, and their structures usually include a shell, a power system (main power supply composed of solar panels, backup power supply) and a propulsion system. However, the micrometeoroids and space debris in the universe pose a major threat to satellites. The trajectories of these substances are difficult to predict, making satellites extremely vulnerable to impact damage or even direct damage during operation. Although spacecraft can use avoidance measures to deal with larger, slower-moving, and detectable debris, they can only rely on protective structures to reduce damage to smaller, faster-moving space debris that cannot be detected.

[0004] However, although the thick and solid shell can resist some impacts, it is difficult to cope with the impact of high-speed micrometeoroids and space debris, and may cause the satellite to increase in size and weight, which in turn greatly increases the cost. Therefore, it is very important to obtain the behavior and status information of the satellite after the impact, so as to timely assess the damage and take corresponding measures to ensure the normal operation of the satellite's key systems; first of all, it should be clear that after the impact occurs, the impact position needs to be accurately located to determine whether it has lost its main working ability. In the existing technology, redundant sensors and thin-film sensors are usually used to cover the surface of the artificial satellite as much as possible, but carrying too many sensors will also increase the cost of the artificial satellite, and because the impact position is unpredictable, there will be a phenomenon that the specific position is difficult to determine because it happens to hit between two sets of sensors.

[0005] Therefore, the present invention proposes a spacecraft structure impact source positioning system to solve the above-mentioned problem. Summary of the invention

[0006] In view of the deficiencies in the prior art, the purpose of the present invention is to provide a spacecraft structure impact source positioning system that can initially avoid collisions between spacecraft and meteoroids and space debris, and accurately locate the impact point after a collision occurs.

[0007] To achieve the above object, the present invention provides the following technical solutions: A spacecraft structure impact source positioning system, including a detection system, a collision warning system, and a processing system; The detection system is used to obtain and transmit state information of the spacecraft surface. The detection system includes a sensor module and a data acquisition module. The sensor module includes multiple groups of sensors, which are divided into multiple groups of shell sensors arranged on the outer shell of the spacecraft, multiple groups of power supply sensors arranged on the outer wall of the spacecraft power supply, and multiple groups of propulsion sensors arranged in the spacecraft propulsion system according to their functions. The sensor module detects stress wave information at the placement position and converts the stress wave into an electrical signal for transmission. The data acquisition module is used to receive the electrical signal transmitted by the sensor module, and pack the electrical signal and auxiliary information into a data group for output; The collision warning system is used to avoid collisions between spacecraft and other space objects. By simulating the current operating state of the spacecraft in real time and predicting its trajectory, it is easy to determine whether the spacecraft needs to change its current operating state. The collision warning system includes ground radar, optical telescope and space monitoring satellite, which monitor and record the position, speed and orbital parameters of satellites and space debris in the Earth orbit in real time. The processing system is used to handle collisions that cannot be avoided or caused by the failure of the collision warning system. The processing system receives the data group transmitted by the data acquisition module, and performs calculations on the received data group. The specific location and degree of the impact on the spacecraft are obtained through analysis of the processed information, and the current overall state of the spacecraft is determined.

[0008] By adopting the above technical scheme, the present invention obtains the state information of the spacecraft surface in real time through the detection system, and comprehensively monitors whether the spacecraft has been hit by a collision. Through the collision warning system, early warning and avoidance of potential collisions with other space objects can be given, minimizing the risk of damage to the spacecraft. For collisions that cannot be avoided or have already occurred, the processing system can respond quickly, and the stress wave analysis technology is used to determine the specific location of the collision through precise calculation. In the processing system, the present invention cleverly utilizes the propagation characteristics of stress waves, and realizes comprehensive monitoring of the entire surface of the spacecraft through multiple sets of hyperbolic positioning calculation methods, and only a small number of PVDF sensors are required to cover all external areas. In addition, through the attenuation characteristics of stress waves, the specific impact stress of the impact point is accurately calculated, and the impact position and stress data are combined to provide a scientific basis for the maintenance and repair of the spacecraft. Through this comprehensive monitoring and processing scheme, the present invention not only ensures the safe operation of the spacecraft in a complex space environment, but also can quickly make effective treatment when damage occurs, ensuring that the spacecraft efficiently completes various tasks.

[0009] The present invention is further configured as follows: multiple groups of the housing sensors, multiple groups of power sensors, and multiple groups of propulsion sensors use PVDF sensors, and the stress wave signals generated by the impact are received by the PVDF sensors; Among them, multiple groups of shell sensors, multiple groups of power sensors, and multiple groups of propulsion sensors are distributed in a grid and are arranged and divided according to the sensing range and stress wave receiving capability of the sensors.

[0010] The present invention is further configured as follows: a plurality of groups of the shell sensors are used to detect whether the shell of the spacecraft is impacted, a plurality of groups of the power supply sensors are used to detect whether the power supply of the spacecraft is impacted, and a plurality of groups of the propulsion sensors are used to detect whether the propulsion system of the spacecraft is impacted; Multiple groups of the housing sensors are set to type A; multiple groups of the power supply sensors are set to type B; and multiple groups of the propulsion sensors are set to type C.

[0011] The present invention is further configured as follows: the auxiliary information includes sensor type information and sensor position information, wherein the sensor position information is coordinate information of the sensor located at different parts of the spacecraft; Among them, the coordinate information of different parts of the spacecraft is divided into planes based on the artificially customized baseline, and all the planes are summarized on the overall plane. A rectangular coordinate system is established with one end point of the overall plane as the reference point, and each point on the plane and each group of sensor positions are represented by (x, y) coordinates; The sensor type information and the sensor location information are stored in the sensor type table in the format of a triple, recorded as (type, x, y).

[0012] The present invention is further configured as follows: the general plane is composed of different part planes, the different part planes are separated from different parts, and the different part planes and all (x, y) coordinates thereof are recorded into a plane position type table; The data recorded in the plane position type table include the part type, all (x, y) coordinates of the type, and impact resistance, where the impact resistance is the maximum impact stress that the part can withstand.

[0013] The present invention is further configured as follows: the processing system obtains the impact point information by analyzing the arrival time difference and position of the stress wave generated after the impact occurs. The specific processing flow is as follows: S1. After the impact occurs, at least four groups of sensors around the impact point receive stress wave signals; S2. After at least four groups of sensors receive stress wave signals, multiple sets of hyperbolic equations are established based on the time difference of receiving the stress waves: Among them, each set of coordinates obtained are the coordinates of the possible impact points, , are four sets of sensor position coordinates, , is the time difference of the group receiving the stress wave signal corresponding to each equation, is the propagation speed of stress waves, which takes different values ​​according to the material of the spacecraft; S3, through the optimization algorithm to obtain each group of possible collision point coordinates Perform error processing; S4. Perform position determination based on the obtained collision point coordinates (x, y).

[0014] The present invention is further configured as follows: the position determination in S4 is performed by retrieving a plane position type table to determine the position of the impact point (x, y) located in the spacecraft.

[0015] The present invention is further configured to calculate the impact stress of the impact point according to the impact point coordinates (x, y) and a random group of sensors receiving stress wave signals: A1. Obtain the straight-line distance between a random group of sensors receiving stress wave signals and the impact point. The calculation formula is: Where i is a random number between 1 and 4 (i=1, 2, 3, 4); A2. According to the attenuation phenomenon of stress wave during propagation, an attenuation model is set, and its calculation formula is: Among them, Fr is the stress data at the impact point, Ft is the stress data of a random group of sensors receiving stress wave signals, d is the straight-line distance between a random group of sensors receiving stress wave signals obtained in A1 and the impact point, and n is the attenuation index, which is set to 2; A3. Conduct stress comparison; If the Fr data exceeds the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a dangerous state at this time; If the Fr data does not exceed the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a safe state at this time.

[0016] The present invention is further configured as follows: the dangerous state is that the spacecraft is hit by a collision that damages its structure, and the processing system feeds back the impact time, impact coordinates, and impact stress to the operator, and the operator takes corresponding measures to handle the situation according to the impact coordinates and impact stress; The safe state is that the spacecraft continues to perform its mission along the original orbit even if it suffers a collision that does not affect its structure.

[0017] The present invention is further configured as follows: the collision warning system monitors satellites and space debris around the orbit that may collide with the spacecraft according to the orbit of the spacecraft, and does not take any measures when it is detected that smaller debris is approaching; If larger debris approaches, the collision warning system sends a signal to the propulsion system to fine-tune the spacecraft's orbit to avoid collision with the larger debris. After the collision warning system fine-tunes the spacecraft's orbit, the original orbital data is retained, and when there are no satellites or space debris that may collide with it in the original orbit, the propulsion system is controlled to return it to its original orbit.

[0018] In summary, the present application includes at least one of the following beneficial technical effects: 1. The present invention obtains the status information of the spacecraft surface in real time through the detection system, and comprehensively monitors whether the spacecraft has suffered a collision. Through the collision warning system, early warning and avoidance of potential collisions with other space objects can be given, thereby minimizing the risk of damage to the spacecraft. For collisions that cannot be avoided or have already occurred, the processing system can respond quickly, and the stress wave analysis technology is used to determine the specific location of the collision through precise calculation. In the processing system, the present invention cleverly utilizes the propagation characteristics of stress waves, and through multiple sets of hyperbolic positioning calculation methods, it realizes comprehensive monitoring of the entire surface of the spacecraft, and only a small number of PVDF sensors are required to cover all external areas. In addition, through the attenuation characteristics of stress waves, the specific impact stress of the impact point is accurately calculated, and the impact position and stress data are combined to provide a scientific basis for the maintenance and repair of the spacecraft. Through this comprehensive monitoring and processing scheme, the present invention not only ensures the safe operation of the spacecraft in a complex space environment, but also can quickly and effectively deal with damage when it occurs, ensuring that the spacecraft efficiently completes various tasks.

[0019] 2. The present invention adopts multiple sets of hyperbola positioning calculation methods to calculate the coordinates of the impact point, and performs positioning through the signals of four sets of sensors, which can reduce the influence of signal interference or error of a single transmitting station, and the intersection of multiple sets of hyperbolas can reduce measurement errors and improve positioning accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 The present invention is a module schematic diagram of a spacecraft structure impact source positioning system.

[0021] Figure 2 It is a calculation principle diagram of the multi-group hyperbolic positioning calculation method in the present invention. DETAILED DESCRIPTION

[0022] It should be noted that, in the absence of conflict, the embodiments and features in the embodiments of the present application may be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0023] It should be noted that, unless otherwise specified, all technical and scientific terms used in this application have the same meanings as commonly understood by ordinary technicians in the technical field to which this application belongs.

[0024] See also Figure 1-Figure 2 , the present invention provides the following technical solutions: For example, see Figure 1-Figure 2 , a spacecraft structure impact source positioning system, including a detection system, a collision warning system, and a processing system.

[0025] The detection system is used to obtain and transmit the status information of the spacecraft surface. The detection system includes a sensor module and a data acquisition module. The sensor module includes multiple groups of sensors, which are divided into multiple groups of shell sensors arranged on the outer shell of the spacecraft, multiple groups of power supply sensors arranged on the outer wall of the spacecraft power supply, and multiple groups of propulsion sensors arranged in the spacecraft propulsion system according to their functions; the sensor module detects the stress wave information at the placement position and converts the stress wave into electrical signals for transmission. The data acquisition module is used to receive the electrical signals transmitted by the sensor module, and package the electrical signals and auxiliary information into data groups for output.

[0026] Because PVDF (polyvinylidene fluoride) can effectively receive and detect stress waves through its piezoelectric effect. When the stress wave propagates to the PVDF sensor, the PVDF sensor is subjected to mechanical stress and generates charges, which form voltage signals between the electrodes. Through appropriate circuits and signal processing techniques, these voltage signals can be detected and analyzed, thereby realizing the detection and identification of stress waves. Therefore, multiple groups of shell sensors, multiple groups of power sensors, and multiple groups of propulsion sensors in this embodiment use PVDF sensors, and the stress wave signals generated by the impact are received by the PVDF sensors; wherein, multiple groups of shell sensors, multiple groups of power sensors, and multiple groups of propulsion sensors are in the form of grid distribution, and are arranged and divided according to the sensing range and the ability to receive stress waves of the sensors. Since in this embodiment, multiple groups of hyperbolic positioning calculation methods are used to calculate the impact position and impact stress, the spacecraft does not need to arrange as many sensors as possible to ensure that there are sensors at each position.

[0027] In this embodiment, multiple groups of shell sensors are used to detect whether the spacecraft shell is impacted, multiple groups of power supply sensors are used to detect whether the spacecraft power supply is impacted, and multiple groups of propulsion sensors are used to detect whether the spacecraft propulsion system is impacted; multiple groups of shell sensors are set to type A; multiple groups of power supply sensors are set to type B; multiple groups of propulsion sensors are set to type C.

[0028] The auxiliary information in the data group packaged by the sensor module includes sensor type information and sensor position information, wherein the sensor position information is the coordinate information of the sensor located at different parts of the spacecraft.

[0029] Among them, the sensor type information and sensor location information are stored in the sensor type table in the format of a triplet, recorded as (type, x, y). The coordinates can be retrieved according to the type, and the type of the current sensor can be retrieved according to the coordinates. This triplet format storage can provide convenience for operators to retrieve the type of sensor at a certain coordinate. In addition, due to the different materials of different parts of the spacecraft, different types of sensors can be selected according to the materials at different positions. For example, the propagation effect and distance of stress waves at the power supply are smaller than those of the outer shell, so a more precise sensor can be selected at the power supply. Therefore, the type here can also represent the specific type of sensor to a certain extent, which is convenient for the operator to perform subsequent maintenance and replacement.

[0030] Among them, since the impact of the spacecraft can only occur on the outside, the coordinate information of different parts of the spacecraft is divided into planes based on the artificially customized baseline, and all the planes are summarized on the overall plane in sequence, and a rectangular coordinate system is established with one end point of the overall plane as the reference point, and each point on the plane and each group of sensor positions are represented by (x, y) coordinates. It should be noted that all planes are summarized on the overall plane in sequence for the uniqueness of the coordinates.

[0031] It should be noted that when all planes are sequentially aggregated on the master plane, the length and shape of the boundary lines of the planes may be difficult to fit together. In this embodiment, the planes only need to be placed sequentially without changing the shape of the planes by cutting or splicing to facilitate splicing. Therefore, there will be wasted empty points in the coordinate system, and the existence of such empty points will not affect the present invention.

[0032] Furthermore, when the parts are split into planes, they are recorded through the three-dimensional model and the corresponding coordinates, and artificially customized baselines are marked in the three-dimensional model for easy observation by operators. Each point in the three-dimensional model has a one-to-one relationship with the corresponding coordinate point. It should be noted that the technology for splitting the three-dimensional model into planes is very mature in the prior art, so this embodiment does not make specific limitations on it. For example, after using MATLAB's plot3 function to draw a three-dimensional figure, use the view function to project it onto a two-dimensional plane.

[0033] If a set of rectangular coordinate systems are established for all planes, then all planes will have the same coordinates (x, y). The existence of such identical coordinates requires the introduction of another digit to determine which plane it is located on. This method increases the difficulty of storage; and this method also makes it difficult for operators to observe, because they need to determine which plane the (x, y) coordinates belong to, increasing the risk of error. Aggregating all planes into a total plane seems to increase the area of ​​the total plane. However, after this operation, each set of (x, y) coordinates corresponds to a unique point, ensuring the unique correspondence between coordinates and positions.

[0034] The general plan is composed of different part planes, and different part planes are composed of different parts. The different part planes and all their (x, y) coordinates are recorded in a plane position type table. Through the plane position type table, it is possible to retrieve which type of part each set of coordinates is located in, so as to determine its impact resistance.

[0035] The data recorded in the plane position type table include the part type, all (x, y) coordinates of the type, and impact resistance. The impact resistance is the maximum impact stress that the part can withstand. It should be noted that the impact resistance is usually determined by the material of the part and the function that the part bears. For example, usually because the shell needs to protect the working parts inside, the material of the shell will be relatively hard, and the shell does not have independent performance, so its impact resistance is usually strong; and the power supply is usually composed of solar panels, the strength of the solar panel is low, and because it must receive sunlight through the solar panel, it is impossible to install protective measures on the outside, so the power supply has a weak impact resistance. In this embodiment, the plane position type table is shown in Table 1:

[0036] Among them, since all (x, y) correspond to unique locations, Coordinates, that is, (x, y) coordinates are set as the primary key in this table. The entire table can be searched through Coordinates. This method can ensure that no errors occur when staff search for locations, thereby improving the robustness of the system.

[0037] In this embodiment, the collision warning system is used to avoid collision between the spacecraft and other space objects. By performing real-time simulation of the current operating state of the spacecraft and trajectory prediction, it is easy to intuitively understand and judge whether the state of the spacecraft needs to change the current operating state; the collision warning system includes ground radar, optical telescope and space monitoring satellite, which monitor and record in real time the position, speed and orbital parameters of satellites and space debris in the Earth orbit.

[0038] It should be noted that the collision warning system usually establishes a simulation model in advance when the spacecraft has not been launched. Before launch, a synchronous simulation is performed based on the size and trajectory of the spacecraft, as well as the trajectory and size of other artificial satellites in similar orbits, to avoid collisions caused by orbital conflicts.

[0039] The collision warning system monitors satellites and space debris around the spacecraft that may collide with it based on the spacecraft's orbit. When it detects that smaller debris is approaching, no measures are taken because it is difficult for smaller debris to have an impact on the spacecraft, and each maneuver consumes fuel, affecting the life of the spacecraft and mission execution. If larger debris approaches, in order to avoid larger debris hitting the spacecraft and causing damage to the spacecraft, the collision warning system sends a signal to the propulsion system to fine-tune the spacecraft's orbit to avoid collision with larger debris, while also avoiding conflicts with other satellites. After the collision warning system fine-tunes the spacecraft's orbit, it retains the original orbit data and monitors the original orbit in real time. When there are no satellites or space debris that may collide with it in the original orbit, the propulsion system is controlled to immediately return it to the original orbit to avoid the satellite's original operation being unable to be carried out due to orbit changes.

[0040] However, in actual operation, not all satellites and space debris can be captured by the collision warning system. Due to the rather complex situation in space, it is difficult for the collision warning system to detect when too fast micrometeorites and space debris approach. In addition, if a collision occurs between micrometeorites, the products caused by the collision may hit the spacecraft at an unpredictable angle. In this case, it is almost impossible for the collision warning system to respond. Therefore, the present invention also provides a processing system for processing collisions that cannot be avoided by the collision warning system or that fail to warn.

[0041] Specifically, the processing system receives the data groups transmitted by the data acquisition module, and performs calculations on the received data groups. Through analysis of the processed information, the specific location where the spacecraft was hit and the degree of the impact are obtained, and the current overall state of the spacecraft is judged, so that the operator can decide whether the spacecraft should continue to perform the mission, or be repaired or even recalled based on the overall state of the spacecraft.

[0042] Therefore, the processing system needs to obtain the precise location and degree of impact of the spacecraft. The processing system obtains the impact point information by comprehensively analyzing the arrival time difference and location of the stress wave generated after the impact. The specific processing flow is as follows: S1. After a collision occurs, at least four groups of sensors around the collision point receive stress wave signals. It should be noted that usually after a collision occurs, more than four groups of sensors can receive stress wave signals. In order to avoid stress interference caused by multiple collisions and the influence of stress wave attenuation, the four groups of sensors that detect stress waves earliest are used for calculation in this embodiment. S2. After at least four groups of sensors receive the stress wave signal, multiple groups of hyperbolic equations are established according to the time difference of receiving the stress wave. It should be noted that in order to ensure the accuracy of the coordinates of the impact point where the impact may occur, when establishing the hyperbolic equation, multiple groups of hyperbolic equations are established as much as possible. Generally speaking, at least four groups of hyperbolic equations are established: Among them, the coordinates The coordinates of the impact point where the impact may occur are: , are the position coordinates of the four groups of sensors sorted according to the time difference, It is the position information of the first group of sensors that received the stress wave, and so on; is the time difference of the group receiving the stress wave signal corresponding to each equation, is the propagation speed of stress waves, which takes different values ​​according to the material of the spacecraft (usually the speed of sound); It should be noted that a sensor may receive two or more stress waves in a short period of time. This may be due to two or more impacts occurring on the surface of the spacecraft. Since it is uncertain which impact caused the stress wave, it may interfere with the calculation of the coordinates of the impact point where the impact may occur. Therefore, in order to avoid this situation, the present invention solves this problem through conflict avoidance.

[0043] The specific conflict avoidance method is to monitor a sensor when it receives a stress wave. The monitoring time is the transmission time of the stress wave within the distance between the two groups of sensors. Since the distance between the two groups of sensors is fixed, the monitoring time equation is calculated according to the propagation speed of the stress wave. The calculation formula is:

[0044] Among them, Ti is the monitoring time, D is the distance between the two sets of sensors, is the propagation speed of stress waves, which takes different values ​​according to the material of the spacecraft (usually the speed of sound); When a sensor receives a stress wave, it immediately starts monitoring. If the sensor receives more than one set of stress wave information during the monitoring time, the sensor coordinates are discarded and this set of coordinates is not taken into the hyperbolic equation for calculation. Although the stress wave transmission direction and stress wave propagation speed can also be used to determine which impact the stress wave received by this sensor is, since more than one set of hyperbolic equations are established in this embodiment, discarding one set of hyperbolic equations will not have a substantial impact on the results. Therefore, discarding the stress wave information obtained by this sensor is a more time-saving and cost-effective option.

[0045] It should be noted that if the time and distance of the two impacts are close, it is easy for multiple groups of sensors to receive more than one group of stress waves within the monitoring time. At this time, although all these data and information are discarded, for the sake of accuracy, this embodiment requires two sets of sensor coordinates. Due to the different impact points, during the stress wave transmission process, there will definitely be sensors that only receive one set of stress waves within the monitoring time.

[0046] Correspondingly, if the sensor only receives one set of stress wave information during the monitoring time, the information received by the sensor is valid, indicating that no two impacts occurred within a short period of time, and the set of sensor coordinates can be brought into the hyperbola formula for calculation.

[0047] S3. Each group of coordinates of possible collision points is obtained through the optimization algorithm. Perform error processing; It should be noted that, usually, U sets of hyperbolic equations will obtain U sets of impact point coordinates where impact may occur. However, due to measurement errors and actual accuracy, the number of actual coordinates obtained may be reduced, or even only one set may be obtained. In this embodiment, in order to obtain the most accurate impact point coordinates, an optimization algorithm is used to minimize the error, wherein the formula of the optimization algorithm is: Among them, u is the number of coordinates of possible collision points obtained by multiple sets of hyperbolic equations. If u is 1, the set of coordinates is directly output; if u is greater than 1, the coordinates calculated by the optimization algorithm are output.

[0048] It should be noted that the coordinates of the impact point can be obtained to a certain extent by using ordinary hyperbolic point positioning method or single point positioning method. However, its positioning accuracy is difficult to reach the accuracy level required by the spacecraft. Positioning through signals from multiple sensors can reduce the impact of signal interference or errors from a single transmitting station, and the intersection of multiple sets of hyperbolas can reduce measurement errors and improve positioning accuracy.

[0049] S4. Perform position determination based on the obtained collision point coordinates (x, y).

[0050] Specifically, the position determination in S4 determines the position of the impact point (x, y) of the spacecraft by searching the plane position type table.

[0051] In actual operation, after a collision occurs, only obtaining the position of the impact point can only determine where the collision occurred, but cannot evaluate the overall state of the spacecraft. To obtain the overall state of the spacecraft, not only the coordinates of the impact point are needed, but also the impact force at the impact point.

[0052] Specifically, the impact stress at the impact point is calculated based on the impact point coordinates (x, y) and a random group of sensors that receive stress wave signals: A1. Obtain the straight-line distance between a random group of sensors receiving stress wave signals and the impact point. It should be noted that this straight-line distance is the propagation distance of the stress wave, and the calculation formula is: Where i is a random number within 1-u (i=1, 2, 3, 4...); A2. Since stress waves will attenuate during propagation due to factors such as material damping and nuclear diffusion, it is necessary to set up an attenuation model, and its calculation formula is: Among them, Fr is the stress data at the impact point, Ft is the stress data of a random group of sensors receiving stress wave signals, d is the straight-line distance between a random group of sensors receiving stress wave signals obtained in A1 and the impact point, and n is the attenuation index, which is set to 2; A3. Conduct stress comparison; If the Fr data exceeds the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a dangerous state at this time; If the Fr data does not exceed the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a safe state at this time.

[0053] The dangerous state is when the spacecraft is hit by a collision that damages its structure. The processing system will feed back the impact time, impact coordinates and impact stress to the operator, who will take appropriate measures to handle the situation based on the impact coordinates and impact stress.

[0054] The safe state is when the spacecraft suffers an impact that does not affect its structure and continues to move along its original orbit to perform its mission.

[0055] The present invention obtains the state information of the spacecraft surface in real time through the detection system, and comprehensively monitors whether the spacecraft has been hit by a collision. Through the collision warning system, early warning and avoidance of potential collisions with other space objects can be provided, thereby minimizing the risk of damage to the spacecraft. For collisions that cannot be avoided or have occurred, the processing system can respond quickly, and the stress wave analysis technology can be used to determine the specific location of the collision through precise calculation.

[0056] In the processing system, the propagation characteristics of stress waves are used to achieve comprehensive monitoring of the entire surface of the spacecraft through multiple sets of hyperbolic positioning calculation methods, and only a small number of PVDF sensors are needed to cover all external areas. In addition, the specific impact stress of the impact point is accurately calculated through the attenuation characteristics of stress waves, and the combination of impact position and stress data provides a scientific basis for the maintenance and repair of spacecraft.

[0057] Through this comprehensive monitoring and processing solution, the present invention not only ensures the safe operation of the spacecraft in a complex space environment, but also can quickly and effectively deal with damage when it occurs, ensuring that the spacecraft can efficiently complete various tasks.

[0058] The device embodiments described above are merely illustrative and are not all embodiments. For example, the division of the units is only a logical function division, and there may be other division methods in actual implementation. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed may be through some communication interfaces, indirect coupling or communication connection of devices or units, which may be electrical, mechanical or other forms. All other embodiments obtained by ordinary technicians in this field without making creative work should fall within the scope of protection of the present invention.

Claims

1. A spacecraft structure impact source positioning system, characterized by: Including detection system, collision warning system, and processing system; The detection system is used to obtain and transmit state information of the spacecraft surface. The detection system includes a sensor module and a data acquisition module. The sensor module includes multiple groups of sensors, which are divided into multiple groups of shell sensors arranged on the outer shell of the spacecraft, multiple groups of power supply sensors arranged on the outer wall of the spacecraft power supply, and multiple groups of propulsion sensors arranged in the spacecraft propulsion system according to their functions. The sensor module detects stress wave information at the placement position and converts the stress wave into an electrical signal for transmission. The data acquisition module is used to receive the electrical signal transmitted by the sensor module, and pack the electrical signal and auxiliary information into a data group for output; The collision warning system is used to avoid collisions between spacecraft and other space objects. By simulating the current operating state of the spacecraft in real time and predicting its trajectory, it is easy to determine whether the spacecraft needs to change its current operating state. The collision warning system includes ground radar, optical telescope and space monitoring satellite, which monitor and record the position, speed and orbital parameters of satellites and space debris in the Earth orbit in real time. The processing system is used to handle collisions that cannot be avoided or caused by the failure of the collision warning system. The processing system receives the data group transmitted by the data acquisition module, and performs calculations on the received data group. The specific location and degree of the impact on the spacecraft are obtained through analysis of the processed information, and the current overall state of the spacecraft is determined.

2. A spacecraft structure impact source positioning system according to claim 1, characterized in that: The multiple groups of housing sensors, multiple groups of power sensors, and multiple groups of propulsion sensors use PVDF sensors, and the stress wave signals generated by the impact are received by the PVDF sensors; Among them, multiple groups of shell sensors, multiple groups of power sensors, and multiple groups of propulsion sensors are distributed in a grid and are arranged and divided according to the sensing range and stress wave receiving capability of the sensors.

3. A spacecraft structure impact source positioning system according to claim 1, characterized in that: Multiple groups of the shell sensors are used to detect whether the shell of the spacecraft is impacted, multiple groups of the power supply sensors are used to detect whether the power supply of the spacecraft is impacted, and multiple groups of the propulsion sensors are used to detect whether the propulsion system of the spacecraft is impacted; Multiple groups of the housing sensors are set to type A; multiple groups of the power supply sensors are set to type B; and multiple groups of the propulsion sensors are set to type C.

4. A spacecraft structure impact source positioning system according to claim 3, characterized in that: The auxiliary information includes sensor type information and sensor location information, wherein the sensor location information is the coordinate information of the sensor located at different parts of the spacecraft; Among them, the coordinate information of different parts of the spacecraft is divided into planes based on the artificially customized baseline, and all the planes are summarized on the overall plane. A rectangular coordinate system is established with one end point of the overall plane as the reference point, and each point on the plane and each group of sensor positions are represented by (x, y) coordinates; The sensor type information and the sensor location information are stored in the sensor type table in the format of a triple, recorded as (type, x, y).

5. A spacecraft structure impact source positioning system according to claim 4, characterized in that: The general plane is composed of different part planes, and the different part planes are separated from different parts. The different part planes and all their (x, y) coordinates are recorded in a plane position type table; The data recorded in the plane position type table include the part type, all (x, y) coordinates of the type, and impact resistance, where the impact resistance is the maximum impact stress that the part can withstand.

6. A spacecraft structure impact source positioning system according to claim 5, characterized in that: The processing system obtains the impact point information by analyzing the arrival time difference and position of the stress wave generated after the impact. The specific processing flow is as follows: S1. After the impact occurs, at least four groups of sensors around the impact point receive stress wave signals; S2. After at least four groups of sensors receive stress wave signals, multiple groups of hyperbolic equations are established based on the time difference of receiving the stress waves: Among them, each set of coordinates obtained are the coordinates of the possible impact points, , are four sets of sensor position coordinates, , is the time difference of the group receiving the stress wave signal corresponding to each equation, is the propagation speed of stress waves, which takes different values ​​according to the material of the spacecraft; S3, through the optimization algorithm to obtain each group of possible collision point coordinates Perform error processing; S4. Perform position determination based on the obtained collision point coordinates (x, y).

7. A spacecraft structure impact source positioning system according to claim 6, characterized in that: The position determination in S4 determines the position of the impact point (x, y) of the spacecraft by searching a plane position type table.

8. A spacecraft structure impact source positioning system according to claim 6, characterized in that: The impact stress at the impact point is calculated based on the impact point coordinates (x, y) and a random group of sensors that receive stress wave signals: A1. Obtain the straight-line distance between a random group of sensors receiving stress wave signals and the impact point. The calculation formula is: Where i is a random number between 1 and 4 (i=1, 2, 3, 4); A2. According to the attenuation phenomenon of stress wave during propagation, an attenuation model is set, and its calculation formula is: Among them, Fr is the stress data at the impact point, Ft is the stress data of a random group of sensors receiving stress wave signals, d is the straight-line distance between a random group of sensors receiving stress wave signals obtained in A1 and the impact point, and n is the attenuation index, which is set to 2; A3. Conduct stress comparison; If the Fr data exceeds the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a dangerous state at this time; If the Fr data does not exceed the maximum impact stress that can be sustained at the impact point recorded in the plane position type table, it means that the spacecraft is in a safe state at this time.

9. A spacecraft structure impact source positioning system according to claim 8, characterized in that: The dangerous state is when the spacecraft is hit by a collision that damages its structure. The processing system feeds back the impact time, impact coordinates, and impact stress to the operator, who then takes appropriate measures to handle the situation based on the impact coordinates and impact stress. The safe state is that the spacecraft continues to perform its mission along the original orbit even if it suffers a collision that does not affect its structure.

10. The spacecraft structure impact source positioning system according to claim 1, characterized in that: The collision warning system monitors satellites and space debris around the orbit that may collide with the spacecraft according to its orbit, and does not take any measures when it detects that smaller debris is approaching. If larger debris approaches, the collision warning system sends a signal to the propulsion system to fine-tune the spacecraft's orbit to avoid collision with the larger debris. After the collision warning system fine-tunes the spacecraft's orbit, the original orbital data is retained, and when there are no satellites or space debris that may collide with it in the original orbit, the propulsion system is controlled to return it to its original orbit.

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