Satellite autonomous safety management method and system

Abstract

The application discloses a satellite autonomous safety management method and system, which is based on inertial characteristic information, predicts the satellite-rocket separation probability, controls the satellite main control system to be powered on when the satellite-rocket separation probability is equal to a first preset value, then establishes a collision early warning operation environment after the satellite main control system is powered on, loads a collision early warning program and safety strategy parameters, and directly starts collision risk detection based on the preloaded collision early warning program and safety strategy parameters after the satellite-rocket separation is completed. The method and system do not need telemetry information, discrete state signals or additional control interfaces provided by a carrier rocket, and can complete separation window prediction, autonomous power-on control and collision early warning strategy pre-starting only by using information obtained by a satellite inertial sensor, and the system has high universality and autonomous operation capability.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a satellite autonomous safety management method and system. Background Technology

[0002] With the rapid development of large-scale satellite networking projects such as low-Earth orbit (LEO) internet constellations, remote sensing constellations, and navigation enhancement constellations, mass production, dense stacking launches, and rapid on-orbit deployment of satellites have become important trends in spacecraft development. New-generation satellite systems, represented by LEO internet satellites, typically employ a multi-satellite launch or stacking launch method, deploying dozens or even hundreds of satellites in a single launch mission. To ensure safety and reliability during the launch phase, satellites are usually in a non-operational state during the launch vehicle's flight, only being autonomously powered on after separation via a change in the state of a limit switch. This method is simple in structure, highly reliable, and has been widely applied in various spacecraft missions.

[0003] However, as the scale of low-Earth orbit internet satellite constellations continues to expand and the requirements for satellite autonomous operation capabilities increase, the traditional power-on method based on separation event triggering has gradually revealed certain limitations. First, the satellite can only start its main control system after confirming successful separation from the launch vehicle. This necessitates initialization and self-checks for the satellite's computer, attitude control system, and navigation equipment after separation, resulting in a response delay. For satellite systems that require rapid establishment of autonomous control capabilities, rapid acquisition of navigation information, or rapid execution of post-separation safety strategies, this delay can impact the overall system efficiency. Furthermore, when this single triggering condition malfunctions or fails, the satellite may fail to power on as expected, necessitating further improvements in system robustness and autonomy.

[0004] On the other hand, during the terminal phase of a launch vehicle's flight and the satellite-rocket separation phase, the rocket system typically undergoes several typical flight maneuvers, including engine shutdown, attitude adjustment, inertial flight, final stage orientation, and separation preparation. During these processes, the inertial characteristics sensed by the satellite, such as acceleration, angular velocity, and vibration spectrum, undergo significant changes. These inertial characteristics can be acquired by the satellite's own sensors without relying on additional interface information from the launch vehicle, offering advantages such as strong autonomy, good versatility, and adaptability to various launch vehicle models. However, currently, there is a lack of a technical solution that utilizes multi-source inertial characteristic fusion analysis to predict the satellite-rocket separation window and, based on this, enable autonomous satellite wake-up and early deployment of separation collision warning strategies. Summary of the Invention

[0005] To address some or all of the problems of existing technologies and in order to improve the rapid response capability and autonomous operation capability of satellites after separation, the first aspect of this invention provides a satellite autonomous safety management method, comprising: Based on inertial feature information, predict the probability of separation of the satellite and rocket or the predicted separation time. When the probability of separation between the satellite and the launch vehicle is equal to the first preset value, or when a specified time elapses before the predicted separation time, the satellite's main control system is powered on and put into operation. Establish a collision warning operating environment and load the collision warning program and safety policy parameters; Once the spacecraft and rocket have separated, collision risk detection will begin directly based on the pre-loaded collision warning program and safety strategy parameters.

[0006] Furthermore, the inertial characteristic information includes the satellite's three-axis acceleration information, three-axis angular velocity information, structural vibration characteristic information, and inertial characteristic change rate information during the launch vehicle's flight.

[0007] Furthermore, the structural vibration characteristic information includes: The root mean square value of vibration, used to characterize the vibration intensity of the satellite structure, is calculated based on the following formula: , in, The acceleration value at the sampling time. This represents the number of data points within the sampling window. Vibration spectrum characteristic parameters, which characterize the distribution of vibration energy, are obtained based on vibration signals through spectrum analysis.

[0008] Furthermore, the prediction of the satellite-rocket separation probability or satellite-rocket separation time based on inertial feature information includes: By integrating the inertial feature information, the flight phases of the launch vehicle's final flight stage are determined, wherein the flight phases include engine shutdown state, inertial flight state, final stage attitude adjustment state, and separation preparation state. When the current state is determined to be engine off, a separation probability function is established based on the inertial feature information to determine the star-rocket separation probability.

[0009] Furthermore, when the average longitudinal acceleration decreases beyond the second preset value within N consecutive sampling periods, and the root mean square of vibration decreases beyond the third preset value, the current state is determined to be the engine shutdown state.

[0010] Furthermore, establishing the separation probability function based on the inertial feature information includes: Based on the average longitudinal acceleration during engine operation, determine the engine shutdown characteristic quantities; The rate of change of angular velocity is determined based on the triaxial angular velocity information; The triaxial acceleration information is subjected to a fast Fourier transform to obtain the current vibration spectrum and the spectrum stability factor is determined. After determining that the current state is engine off, start the timer and determine the time feature quantity based on the cumulative duration since the engine was off; The engine shutdown characteristic, vibration root mean square value change, angular velocity change rate, spectral stability factor, and time characteristic are normalized and then weighted to establish a separation probability function.

[0011] Furthermore, the first preset value ranges from 0.7 to 0.85.

[0012] Furthermore, by sending a wake-up command to the satellite's main power distribution system, the satellite's main control system is powered on and put into operation.

[0013] Furthermore, establishing a collision warning operating environment includes: Read the pre-stored satellite release sequence information, satellite separation direction information, and separation velocity parameters; Establish a safety monitoring state machine for the separation phase.

[0014] Based on the satellite autonomous safety management method described above, a second aspect of the present invention provides a satellite autonomous safety management system, comprising: The inertial information acquisition module is used to collect inertial characteristic information of the satellite in real time during the flight of the launch vehicle; The separation window prediction module is used to predict the star-rocket separation probability or star-rocket separation time based on the inertial feature information. The autonomous power-on control module is used to control the satellite main control system to power on and run when the satellite-rocket separation probability is equal to a specified value or when a specified time before the satellite-rocket separation time is reached. The collision warning strategy pre-start module is used to establish a collision warning operating environment and load the collision warning program and safety strategy parameters before the star-rocket separation.

[0015] This invention discloses a satellite autonomous safety management method and system. By collecting inertial information such as acceleration, angular velocity, and vibration characteristics of the satellite during the launch phase, it identifies flight states such as launch vehicle engine shutdown, inertial flight, attitude adjustment, and separation preparation. Furthermore, it constructs a separation window prediction model to achieve early prediction of satellite-launcher separation events. Compared to existing methods that rely on separation switches or mechanical release events to trigger power-on, this method enables the satellite to autonomously perceive and predict the separation process, improving the satellite's autonomous operation level.

[0016] The described method and system, when predicting the approaching satellite-launch separation window, control the satellite's main control system to enter operational status in advance. After separation, there is no need to execute the complete system power-on procedure, significantly shortening the response time after separation and increasing the speed at which the satellite enters normal operating status. Simultaneously, after the main control system enters operational status, collision warning procedures, separation safety management procedures, and preset avoidance strategy parameters are pre-loaded, enabling the collision warning system to complete the establishment of the operating environment and parameter initialization before satellite-launch separation. After separation, the satellite can directly enter the collision risk monitoring and safety decision-making mode, avoiding the safety response lag caused by system startup delays in traditional methods and improving separation safety in multi-satellite continuous release scenarios.

[0017] The described method and system utilize feature information acquired by the satellite's own inertial sensors to predict the separation window. This eliminates the need for additional telemetry information, discrete state signals, or control interfaces from the launch vehicle, and also avoids modifications to the existing satellite-launcher separation mechanism design. Compared to pre-power-on schemes that rely on rocket state information, this effectively reduces the coupling between the satellite and launch vehicle systems, improving compatibility and versatility between different satellite and launch vehicle models. Furthermore, the method and system employ a two-stage control mechanism of separation window prediction and separation confirmation. The inertial feature prediction results serve as the basis for autonomous wake-up, while the separation confirmation signal serves as the final execution basis. This ensures system safety while improving the reliability of wake-up decisions and reducing the risk of the satellite failing to enter operational status in a timely manner due to the failure of a single trigger mechanism.

[0018] Based on the described method, no new mechanical components or electrical interfaces need to be added to the satellite-launch interface, separation mechanism, or stacking structure. Separation window prediction and autonomous power-on control can be achieved simply by adding a low-power autonomous wake-up management unit. This does not change the existing satellite structural design and launch process, resulting in low engineering implementation risk and easy promotion and application. Furthermore, for low-Earth orbit internet satellite systems employing a single launcher for multiple satellites in a stacked configuration, the method enables the main control system to be hot-standby activated and collision warning strategies to be pre-activated before the continuous release of multiple satellites. This improves the satellite's autonomous safety assurance capabilities during the separation phase, making it particularly suitable for the batch launch and rapid deployment requirements of large-scale low-Earth orbit satellite constellations. Attached Figure Description

[0019] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the various embodiments of the present invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope. In the drawings, identical or corresponding parts will be indicated by identical or similar reference numerals for clarity.

[0020] Figure 1 A flowchart illustrating a satellite autonomous safety management method according to an embodiment of the present invention is shown. Figure 2 The diagram shows a structural schematic of a satellite autonomous safety management system according to an embodiment of the present invention. Detailed Implementation

[0021] In the following description, the invention is described with reference to various embodiments. However, those skilled in the art will recognize that the embodiments may be practiced without one or more specific details or in conjunction with other alternatives and / or additional methods or components. In other instances, well-known structures or operations are not shown or described in detail so as not to obscure the inventive points of the invention. Similarly, for illustrative purposes, specific numbers and configurations are set forth to provide a comprehensive understanding of embodiments of the invention. However, the invention is not limited to these specific details. Furthermore, it should be understood that the embodiments shown in the drawings are illustrative representations and are not necessarily drawn to scale.

[0022] In this specification, references to "an embodiment" or "this embodiment" mean that a particular feature, structure, or characteristic described in connection with that embodiment is included in at least one embodiment of the invention. The phrase "in one embodiment" appearing throughout this specification does not necessarily refer to the same embodiment in all instances.

[0023] It should be noted that the embodiments of the present invention describe the method steps in a specific order; however, this is only for illustrating the specific embodiment and not for limiting the order of the steps. On the contrary, in different embodiments of the present invention, the order of the steps can be adjusted according to actual needs.

[0024] Existing satellites typically employ a power-on method after separation from the launch vehicle, triggered by limit switches. The satellite's main control system must complete power-on, self-test, and initialization processes after separation, resulting in response lag. Some solutions utilize the launch vehicle interface signals to assist in separation status detection, but these are all external event-driven response mechanisms and do not yet achieve separation window prediction based on the satellite's own inertial characteristics or autonomous pre-start of the main control system. Therefore, to address the issue of existing satellites' inability to establish autonomous perception and safety decision-making capabilities in advance during the critical separation phase, this invention aims to achieve separation window prediction based on inertial feature fusion, autonomous power-on of the main control system, and pre-start of the separation collision warning strategy, thereby improving the safety and autonomy of satellite separation in multi-satellite stacked launch scenarios. This invention provides a satellite autonomous safety management method and system, comprising an autonomous wake-up management module and a collision warning strategy pre-start module. The autonomous wake-up management module operates continuously during the satellite launch phase, features ultra-low power consumption, and operates independently of the satellite's main power distribution system. The autonomous wake-up management module collects inertial characteristic information of the satellite in real time during the launch vehicle's flight and analyzes this information to identify key state changes in the final stage of the launch vehicle's flight. Based on multiple inertial characteristic parameters, it constructs a separation window prediction model and outputs the satellite-launcher separation probability or separation time prediction. When preset conditions are met, the satellite's main control system is controlled to enter power-on operation mode ahead of schedule. After the main control system starts up, the collision warning strategy pre-start unit begins executing the pre-start process to pre-configure and prepare for potential collision risks after separation. This allows the main control system to directly enter collision risk monitoring mode after satellite-launcher separation without reloading the collision warning program and safety strategy parameters.

[0025] The technical solution of the present invention will be further described below with reference to the accompanying drawings of the embodiments.

[0026] Figure 1 This diagram illustrates a flow chart of a satellite autonomous safety management method according to an embodiment of the present invention. Figure 1 As shown, a satellite autonomous safety management method includes: First, in step 101, the probability or time of satellite-rocket separation is predicted. Based on inertial feature information, the predicted value of satellite-rocket separation probability or separation time is calculated. In one embodiment of the present invention, the inertial feature information includes, but is not limited to, triaxial acceleration information Ax, Ay, Az, triaxial angular velocity information ωx, ωy, ωz, structural vibration feature information, and inertial feature rate of change information. These are acquired through inertial information acquisition sensors, etc. In one embodiment of the present invention, the structural vibration feature information includes the root mean square (RMS) value of vibration and vibration spectrum feature parameter S, wherein the RMS value is the square root of the average square of the acceleration data collected by the inertial measurement unit within a preset time window, used to characterize the vibration intensity of the satellite structure, and its calculation formula is: , in, The acceleration value at the sampling time. This represents the number of data points within the sampling window. The vibration spectrum characteristic parameter S is a characteristic parameter obtained by spectrum analysis of the vibration signal collected by the inertial measurement unit, and is used to characterize the vibration energy distribution state.

[0027] The triaxial acceleration information Ax, Ay, Az, the triaxial angular velocity information ωx, ωy, ωz, the root mean square value of vibration (RMS), and the vibration spectrum characteristic parameter S are used to construct an inertial feature vector: F=[Ax,Ay,Az,ωx,ωy,ωz,RMS,S], The flight phases of the launch vehicle's final flight are determined based on the inertial feature vectors. These phases include engine shutdown, inertial flight, final stage attitude adjustment, and separation preparation. When the current state is determined to be engine shutdown, a separation probability function is established based on the inertial feature information to determine the probability of satellite-rocket separation. In one embodiment of the invention, an engine shutdown event is determined to have occurred when the following conditions are met: the average longitudinal acceleration decreases by more than a preset threshold A1 over N consecutive sampling periods, and the root mean square value of vibration decreases by more than a preset threshold A2. Upon detecting an engine shutdown event, a separation prediction mode is entered. Specifically, this involves further analyzing the rate of change of angular velocity, the stability of the vibration spectrum, the duration of inertial flight, and the completion status of attitude maneuvers. A separation probability function Psep is established based on these parameters, where Psep represents the predicted probability of satellite-rocket separation occurring at the current moment.

[0028] In one embodiment of the present invention, engine shutdown characteristic quantities, attitude adjustment characteristic quantities, vibration stability characteristic quantities, and time characteristic quantities are first constructed. Then, these characteristic quantities are normalized, and a separation probability function is established. Let the mean longitudinal acceleration during engine operation be Aref, and the mean longitudinal acceleration at the current moment be Acurr. Define the engine shutdown characteristic quantity F1 = (Aref - Acurr) / Aref. When F1 is close to 1, it indicates that the engine thrust has significantly decreased or disappeared. Simultaneously, calculate the change in the root mean square value of vibration F2 = 1 - RMScurr / RMSref, where RMSref is the root mean square value of vibration during engine operation, and RMScurr is the current root mean square value of vibration. When the engine is operating, F2 ≈ 0; when the engine is shut down, F2 ≈ 1. Based on this, the occurrence of a final-stage engine shutdown event can be determined by the value of F2. For example, when F2 is greater than 0.9, a final-stage engine shutdown event is considered to have occurred. The attitude adjustment features include the rate of change of angular velocity, and the angular velocity magnitude is determined based on the acquired three-axis angular velocity information ωx (roll axis), ωy (pitch axis), and ωz (yaw axis). The rate of change of angular velocity is calculated as F3 = dΩ / dt. F3 shows a significant change when the launch vehicle performs final-stage attitude adjustment. After the attitude adjustment ends, F3 gradually approaches zero. Therefore, F3 can be used to identify the start and end times of final-stage attitude adjustment, as well as the start time of the separation preparation phase. The vibration stability characteristic includes a spectral stability factor. A Fast Fourier Transform (FFT) analysis is performed on the triaxial acceleration information signal to obtain the current vibration spectrum S(f), and the main frequency components are extracted: f1, f2, f3… When the engine is working, fp corresponds to the maximum peak frequency of the spectrum. Based on this, a spectral stability factor F4 = 1 - ΔS / Sref is constructed, where ΔS represents the difference between the current spectrum and the reference spectrum, and Sref represents the stable spectral characteristics during the inertial flight phase. When F4 approaches 1, it indicates that the rocket has entered a stable inertial flight state. The time feature is determined based on the accumulated time after engine shutdown. A timer is started after engine shutdown is detected. Let T be the accumulated time after engine shutdown, and Tref be the typical separation time preset in the mission configuration file. Then, the time feature F5 can be defined as F5 = T / Tref. When F5 gradually approaches 1, it indicates that the system is gradually approaching the expected separation window. Tref can be configured according to different launch vehicle models and launch missions. Based on this, the separation probability function is expressed as follows: Psep=W1·F1'+W2·F2'+W3·F3'+W4·F4'+W5·F5' Where F1', F2', F3', F4', and F5' are the normalized values ​​of F1, F2, F3, F4, and F5, and W1, W2, W3, W4, and W5 are weighting coefficients, satisfying: W1 + W2 + W3 + W4 + W5 = 1. In one embodiment of the present invention, the weighting parameters can be dynamically adjusted according to different launch missions, for example, they can be set to W1 = 0.25, W2 = 0.20, W3 = 0.20, W4 = 0.15, and W5 = 0.20.

[0029] Next, in step 102, the main control system is powered on. When the satellite-rocket separation probability equals a first preset value, or a specified time before the predicted separation time, the satellite main control system is powered on. In one embodiment of the invention, the first preset value ranges from 0.70 to 0.85. The satellite main control system is powered by the satellite main power distribution system (PCDU), and is in a power-off or dormant state during the launch phase. When conditions are met, a wake-up command is sent to the satellite main power distribution system, thereby controlling the satellite main control system to power on.

[0030] Next, in step 103, collision warning pre-activation. Before satellite-rocket separation, a collision warning operating environment is established, and the collision warning program and safety strategy parameters are loaded. Specifically, this involves reading pre-stored satellite release sequence information, satellite separation direction information, and separation velocity parameters; establishing a safety monitoring state machine for the separation phase; loading the collision risk assessment program; loading the collision avoidance strategy parameter set; and establishing a rapid response task queue after separation. During this process, only program and parameter preparation is completed; collision risk calculation is not performed.

[0031] Finally, in step 104, collision risk detection is performed. Immediately after satellite-rocket separation, collision risk detection begins directly based on the pre-loaded collision warning program and safety strategy parameters. In one embodiment of the invention, when the main control system detects a satellite-rocket separation confirmation event based on limit switch action, it quickly establishes a relative motion model of adjacent satellites based on the pre-loaded release sequence information and separation parameters, combined with real-time navigation data after separation. Since the collision warning program and related parameters have been loaded before separation, the satellite does not need to undergo program initialization again after separation and can directly enter the collision risk assessment and safety decision-making process. Compared with the existing method of power-on triggered by limit switches, this significantly shortens the safety response time after separation and improves the autonomous safety assurance capability in multi-satellite continuous release missions.

[0032] Based on the satellite autonomous safety management method described above Figure 2 This diagram illustrates the structure of a satellite autonomous safety management system according to an embodiment of the present invention. Figure 2As shown, a satellite autonomous safety management system includes an inertial information acquisition module 201, a separation window prediction module 202, an autonomous power-on control module 203, and a collision warning strategy pre-start module 204. The inertial information acquisition module 201 is used to acquire inertial characteristic information of the satellite during the launch vehicle's flight in real time, including, for example, a three-axis accelerometer, a three-axis angular velocity meter, and corresponding computing units. The separation window prediction module 202 is used to predict the satellite-launch vehicle separation probability or separation time based on the inertial characteristic information. The autonomous power-on control module 203 is used to control the satellite's main control system to power on and run when the satellite-launch vehicle separation probability equals a specified value or a specified time before the separation time. The collision warning strategy pre-start module 204 is used to establish a collision warning operating environment and load the collision warning program and safety strategy parameters before satellite-launch vehicle separation.

[0033] The inertial information acquisition module 201, separation window prediction module 202, and autonomous power-on control module 203 are powered by the satellite battery 001 through the primary bus and are in a normally open operating state. The collision warning strategy pre-start module 204 is located in the satellite main control system 003 and is powered by the satellite main power distribution system 002. During the launch phase, it is in a power-off or dormant state. When preset conditions are met, the autonomous power-on control module 203 sends an autonomous power-on control signal to the satellite main power distribution system 002, causing it to start powering on the satellite main control system 003. After receiving the corresponding signal from the satellite-rocket separation limit switch 004, the satellite main control system 003 begins collision detection.

[0034] In addition, the present invention also provides a satellite that includes the satellite autonomous safety management system as described above.

[0035] Although various embodiments of the invention have been described above, it should be understood that they are presented by way of example only and not as limitations. It will be apparent to those skilled in the art that various combinations, modifications, and alterations can be made without departing from the spirit and scope of the invention. Therefore, the breadth and scope of the invention disclosed herein should not be limited by the exemplary embodiments disclosed above, but should be defined solely by the appended claims and their equivalents.

Claims

1. A satellite autonomous safety management method, characterized in that, include: Based on inertial feature information, predict the probability of separation of the satellite and rocket or the predicted separation time. When the probability of separation between the satellite and the launch vehicle is equal to the first preset value, or when a specified time elapses before the predicted separation time, the satellite's main control system is powered on and put into operation. After the satellite main control system is powered on, a collision warning operating environment is established, and the collision warning program and safety strategy parameters are loaded. Once the spacecraft and rocket have separated, collision risk detection will begin directly based on the pre-loaded collision warning program and safety strategy parameters.

2. The satellite autonomous safety management method as described in claim 1, characterized in that, The inertial characteristic information includes the satellite's three-axis acceleration information, three-axis angular velocity information, structural vibration characteristic information, and inertial characteristic change rate information during the launch vehicle's flight.

3. The satellite autonomous safety management method as described in claim 2, characterized in that, The structural vibration characteristic information includes: The root mean square value of vibration, which characterizes the vibration intensity of the satellite structure, is calculated based on the following formula: ， in, The acceleration value at the sampling time. This represents the number of data points within the sampling window. Vibration spectrum characteristic parameters, which characterize the distribution of vibration energy, are obtained based on vibration signals through spectrum analysis.

4. The satellite autonomous safety management method as described in claim 2, characterized in that, The prediction of the satellite-rocket separation probability or satellite-rocket separation time based on inertial feature information includes: By integrating the inertial feature information, the flight phases of the launch vehicle's final flight stage are determined, wherein the flight phases include engine shutdown state, inertial flight state, final stage attitude adjustment state, and separation preparation state. When the current state is determined to be engine off, a separation probability function is established based on the inertial feature information to determine the star-rocket separation probability.

5. The satellite autonomous safety management method as described in claim 4, characterized in that, When the average longitudinal acceleration decreases beyond the second preset value within N consecutive sampling periods, and the root mean square of vibration decreases beyond the third preset value, the current state is determined to be the engine shutdown state.

6. The satellite autonomous security management method as described in claim 4, characterized in that, The step of establishing a separation probability function based on the inertial feature information includes: Based on the average longitudinal acceleration during engine operation, determine the engine shutdown characteristic quantities; The rate of change of angular velocity is determined based on the triaxial angular velocity information; The triaxial acceleration information is subjected to a fast Fourier transform to obtain the current vibration spectrum and the spectrum stability factor is determined. After determining that the current state is engine off, start the timer and determine the time feature quantity based on the cumulative duration since the engine was off; The engine shutdown characteristic, vibration root mean square value change, angular velocity change rate, spectral stability factor, and time characteristic are normalized and then weighted to establish a separation probability function.

7. The satellite autonomous safety management method as described in claim 1, characterized in that, The first preset value ranges from 0.7 to 0.

85.

8. The satellite autonomous safety management method as described in claim 1, characterized in that, The satellite's main control system is powered on and put into operation by sending a wake-up command to the satellite's main power distribution system.

9. The satellite autonomous security management method as described in claim 1, characterized in that, Establishing a collision warning operating environment includes: Read the pre-stored satellite release sequence information, satellite separation direction information, and separation velocity parameters; Establish a safety monitoring state machine for the separation phase.

10. A satellite autonomous safety management system, characterized in that, include: The inertial information acquisition module is configured to acquire inertial characteristic information of the satellite in real time during the flight of the launch vehicle; The separation window prediction module is configured to predict the star-rocket separation probability or star-rocket separation time based on the inertial feature information. The autonomous power-on control module is configured to control the satellite main control system to power on and run when the satellite-rocket separation probability equals a specified value or a specified time before the satellite-rocket separation time. The collision warning strategy pre-start module is configured to establish a collision warning operating environment and load the collision warning program and safety strategy parameters before the star-rocket separation.

Images