A method for calculating satellite inertia matrix under abnormal deployment of satellite solar panels

Abstract

The present application provides a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. The method comprises: monitoring the deployment status of the satellite's solar panels in real time based on a preset sensor, obtaining deployment status data of the solar panels, and determining the deployment status of the solar panels based on the deployment status data; in response to the abnormality in deployment of the solar panels, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panels, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. The present application achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment by obtaining the first inertia matrix of the satellite body and the second inertia matrix of the solar panels when an abnormality in deployment of the solar panels is detected in real time, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. This improves the stability of the satellite's attitude control.

Description

Technical Field

[0001] The present application relates to the field of satellite technology, and in particular to a method for calculating a satellite inertia matrix under abnormal deployment of a satellite solar panel. Background Art

[0002] When a satellite's solar panels are deployed, they change the satellite's mass distribution and inertial properties, which in turn affects the satellite's inertia matrix. Therefore, accurate and real-time acquisition of the satellite's inertia matrix is ​​essential for satellite design, orbit control, and attitude control systems.

[0003] Currently, in the traditional inertia calculation method, the satellite's inertia matrix is ​​calculated based on the mass distribution and geometric shape under static conditions.

[0004] However, during the deployment of solar panels, especially when the deployment of solar panels is abnormal, the inertia distribution of the satellite will change, affecting the satellite's attitude control and orbit adjustment, resulting in a decrease in the accuracy of the calculation of the satellite's inertia matrix, thereby reducing the stability of the satellite's attitude control. Summary of the Invention

[0005] In view of this, the purpose of the present application is to provide a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. When the abnormality of solar panel deployment is monitored in real time, the first inertia matrix of the satellite body and the second inertia matrix of the solar panel are obtained, and the total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix. This achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment, thereby improving the stability of satellite attitude control.

[0006] In a first aspect, an embodiment of the present application provides a method for calculating a satellite inertia matrix when a satellite solar panel is abnormally deployed, the method comprising:

[0007] Based on a preset sensor, the deployment state of the solar panel of the satellite is monitored in real time to obtain deployment state data of the solar panel, and the deployment state of the solar panel is determined based on the deployment state data; wherein the deployment state includes deployment abnormality and complete deployment; and the deployment abnormality includes at least incomplete deployment, asymmetric deployment, and deployment jam;

[0008] In response to the abnormal deployment of the solar panel, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panel; wherein the second inertia matrix represents a rotational inertia matrix of the solar panel;

[0009] The total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix; wherein the total inertia matrix is ​​the inertia matrix of the satellite when the satellite solar panels are abnormally deployed.

[0010] In a possible implementation, when the solar panel is a dual-axis controlled panel, the solar panel includes a first controlled rotation axis and a second controlled rotation axis; and obtaining the second inertia matrix of the solar panel includes:

[0011] Determining a satellite body coordinate system of the satellite and a sailboard fixed coordinate system of the solar sailboard;

[0012] Based on the satellite body coordinate system and the sail panel fixed coordinate system, obtaining a first position of the second controlled rotation axis of the solar sail panel in the satellite body coordinate system and a second position of the second controlled rotation axis of the solar sail panel in the sail panel fixed coordinate system;

[0013] A first rotation angle of a first control rotation axis and a second rotation angle of a second control rotation axis of the solar panel are determined, and a second inertia matrix of the solar panel is obtained based on the first position, the second position, the first rotation angle, and the second rotation angle.

[0014] In a possible implementation, acquiring the second inertia matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle includes:

[0015] Acquire a posture conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle;

[0016] An initial second inertia matrix of the solar panel in the panel-fixed coordinate system is determined, and a second inertia matrix of the solar panel is obtained by calculation based on the initial second inertia matrix.

[0017] In a possible implementation, in the satellite body coordinate system, the rotation direction of the first controlled rotation axis of the solar sail panel is consistent with the y-axis of the satellite body coordinate system; in the satellite body coordinate system, the rotation direction of the second controlled rotation axis of the solar sail panel is consistent with the x-axis of the sail panel fixed coordinate system;

[0018] The first position does not change with the rotation angle of the first control rotation axis, and the second position does not change with the rotation angle of the second control rotation axis.

[0019] In a possible implementation, acquiring a position conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle includes:

[0020] Overlapping the sailboard fixed coordinate system with the satellite body coordinate system to obtain a current sailboard fixed coordinate system, and translating the solar sailboard along the current sailboard fixed coordinate system by a displacement of a first position based on the first position;

[0021] Rotating the solar panel around the y-axis of the sailboard fixed coordinate system by a first rotation angle, and rotating the solar panel around the x-axis of the sailboard fixed coordinate system by a second rotation angle;

[0022] Based on the second position, the solar panel is translated along the negative direction of the current panel fixed coordinate system by the displacement of the second position to obtain a position and posture transformation matrix of the solar panel.

[0023] In a possible implementation, the posture conversion matrix of the solar panel is obtained according to the following formula:

[0024]

[0025] in, represents the position conversion matrix of the solar panel, α represents the first rotation angle, β represents the second rotation angle, (x1, y1, z1) represents the first position, and (x2, y2, z2) represents the second position; represents the displacement of the first position, Indicates the first rotation angle, Indicates the second rotation angle, Indicates the displacement of the second position in the negative direction.

[0026] In a possible implementation, the second inertia matrix of the solar panel is calculated based on the initial second inertia matrix, and is implemented by the following formula:

[0027]

[0028] in, represents the second moment of inertia matrix, represents the initial second inertia matrix, represents the posture conversion matrix of the solar panel, and m represents the mass of a single solar panel.

[0029] In a second aspect, an embodiment of the present application further provides a device for calculating a satellite inertia matrix when a satellite solar panel is abnormally deployed, the device comprising:

[0030] a first acquisition module, configured to monitor the deployment state of the satellite's solar panels in real time based on a preset sensor, obtain deployment state data of the solar panels, and determine the deployment state of the solar panels based on the deployment state data; wherein the deployment state includes deployment abnormality and complete deployment; and the deployment abnormality includes at least incomplete deployment, asymmetric deployment, and deployment jam;

[0031] A second acquisition module is configured to acquire a first inertia matrix of the satellite body and a second inertia matrix of the solar panel in response to an abnormal deployment of the solar panel; wherein the second inertia matrix represents a rotational inertia matrix of the solar panel;

[0032] The third acquisition module is configured to calculate a total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix; wherein the total inertia matrix is ​​the inertia matrix of the satellite when the satellite solar panels are abnormally deployed.

[0033] In a possible implementation, when the solar panel is a dual-axis controlled panel, the solar panel includes a first controlled rotation axis and a second controlled rotation axis; the second acquisition module is specifically configured to:

[0034] Determining a satellite body coordinate system of the satellite and a sailboard fixed coordinate system of the solar sailboard;

[0035] Based on the satellite body coordinate system and the sail panel fixed coordinate system, obtaining a first position of the second controlled rotation axis of the solar sail panel in the satellite body coordinate system and a second position of the second controlled rotation axis of the solar sail panel in the sail panel fixed coordinate system;

[0036] A first rotation angle of a first control rotation axis and a second rotation angle of a second control rotation axis of the solar panel are determined, and a second inertia matrix of the solar panel is obtained based on the first position, the second position, the first rotation angle, and the second rotation angle.

[0037] In a possible implementation manner, the second acquisition module is specifically configured to:

[0038] Acquire a posture conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle;

[0039] An initial second inertia matrix of the solar panel in the panel-fixed coordinate system is determined, and a second inertia matrix of the solar panel is obtained by calculation based on the initial second inertia matrix.

[0040] In a possible implementation, in the satellite body coordinate system, the rotation direction of the first controlled rotation axis of the solar sail panel is consistent with the y-axis of the satellite body coordinate system; in the satellite body coordinate system, the rotation direction of the second controlled rotation axis of the solar sail panel is consistent with the x-axis of the sail panel fixed coordinate system;

[0041] The first position does not change with the rotation angle of the first control rotation axis, and the second position does not change with the rotation angle of the second control rotation axis.

[0042] In a possible implementation manner, the second acquisition module is specifically configured to:

[0043] Overlapping the sailboard fixed coordinate system with the satellite body coordinate system to obtain a current sailboard fixed coordinate system, and translating the solar sailboard along the current sailboard fixed coordinate system by a displacement of a first position based on the first position;

[0044] Rotating the solar panel around the y-axis of the sailboard fixed coordinate system by a first rotation angle, and rotating the solar panel around the x-axis of the sailboard fixed coordinate system by a second rotation angle;

[0045] Based on the second position, the solar panel is translated along the negative direction of the current panel fixed coordinate system by the displacement of the second position to obtain a position and posture transformation matrix of the solar panel.

[0046] In a possible implementation, the posture conversion matrix of the solar panel is obtained according to the following formula:

[0047]

[0048] in, represents the position conversion matrix of the solar panel, α represents the first rotation angle, β represents the second rotation angle, (x1, y1, z1) represents the first position, and (x2, y2, z2) represents the second position; represents the displacement of the first position, Indicates the first rotation angle, Indicates the second rotation angle, Indicates the displacement of the second position in the negative direction.

[0049] In a possible implementation, the second inertia matrix of the solar panel is calculated based on the initial second inertia matrix, and is implemented by the following formula:

[0050]

[0051] in, represents the second moment of inertia matrix, represents the initial second inertia matrix, represents the posture conversion matrix of the solar panel, and m represents the mass of a single solar panel.

[0052] In a third aspect, an embodiment of the present application provides an electronic device comprising: a processor, a storage medium and a bus, wherein the storage medium stores machine-readable instructions executable by the processor. When the electronic device is running, the processor communicates with the storage medium through the bus, and the processor executes the machine-readable instructions to perform the steps of the satellite inertia matrix calculation method under abnormal satellite solar panel deployment as described in any one of the first aspects.

[0053] In a fourth aspect, an embodiment of the present application provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the steps of the method for calculating the satellite inertia matrix under abnormal satellite solar panel deployment as described in any one of the first aspects are executed.

[0054] An embodiment of the present application provides a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. The method monitors the deployment status of the satellite's solar panels in real time based on a preset sensor, obtains deployment status data of the solar panels, and determines the deployment status of the solar panels based on the deployment status data. In response to the abnormality of the solar panel deployment, the method obtains the first inertia matrix of the satellite body and the second inertia matrix of the solar panels, and calculates the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. The present application achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment by obtaining the first inertia matrix of the satellite body and the second inertia matrix of the solar panels when an abnormality of the solar panel deployment is detected in real time, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. This improves the stability of the satellite's attitude control.

[0055] In order to make the above-mentioned objects, features and advantages of the present application more obvious and easy to understand, preferred embodiments are given below and described in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following is a brief introduction to the drawings required for use in the embodiments. It should be understood that the following drawings only show certain embodiments of the present application and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other relevant drawings can be obtained based on these drawings without creative work.

[0057] Figure 1 This is a flow chart of a method for calculating a satellite inertia matrix under abnormal deployment of a satellite solar panel according to an embodiment of the present application;

[0058] Figure 2 This is a flow chart of a method for calculating a satellite inertia matrix under abnormal satellite solar panel deployment according to another embodiment of the present application;

[0059] Figure 3 It is a schematic diagram of the satellite body coordinate system;

[0060] Figure 4 It is a schematic diagram of the sailboard fixed coordinate system;

[0061] Figure 5 It is a schematic diagram of obtaining the posture transformation matrix of the solar panel;

[0062] Figure 6 2 is a schematic structural diagram of a device for calculating a satellite inertia matrix under abnormal deployment of a satellite solar panel according to an embodiment of the present application;

[0063] Figure 7 This is a schematic structural diagram of an electronic device provided according to an embodiment of the present application. DETAILED DESCRIPTION

[0064] In order to make the purpose, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. It should be understood that the drawings in the present application only serve the purpose of illustration and description and are not used to limit the scope of protection of the present application. In addition, it should be understood that the schematic drawings are not drawn to scale. The flowcharts used in this application illustrate the operations implemented according to some embodiments of the present application. It should be understood that the operations of the flowcharts can be implemented out of sequence, and steps without logical context can be reversed or implemented simultaneously. In addition, those skilled in the art, under the guidance of the contents of this application, can add one or more other operations to the flowchart, or remove one or more operations from the flowchart.

[0065] In addition, the described embodiments are only a part of the embodiments of the present application, rather than all of the embodiments. The components of the embodiments of the present application generally described and shown in the drawings here can be arranged and designed in various configurations. Therefore, the following detailed description of the embodiments of the present application provided in the drawings is not intended to limit the scope of the claimed application, but merely represents selected embodiments of the present application. Based on the embodiments of the present application, all other embodiments obtained by those skilled in the art without making creative work are within the scope of protection of the present application.

[0066] It should be noted that the term "comprising" will be used in the embodiments of the present application to indicate the existence of the features declared thereafter, but does not exclude the addition of other features.

[0067] Considering that the deployment of a satellite's solar panels changes the satellite's mass distribution and inertial characteristics, thus affecting the satellite's inertia matrix, it is necessary to accurately obtain the satellite's inertia matrix in real time during satellite design, orbit control, and attitude control systems.

[0068] Currently, traditional inertia calculation methods use the mass distribution and geometry of a satellite under static conditions to calculate its inertia matrix. However, during solar panel deployment, especially when the panels deploy abnormally, the satellite's inertia distribution can change, affecting the satellite's attitude control and orbit adjustment. This can reduce the accuracy of the satellite's inertia matrix calculations and, consequently, the stability of satellite attitude control.

[0069] To address this problem, the present application provides a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. When the abnormality of solar panel deployment is monitored in real time, the first inertia matrix of the satellite body and the second inertia matrix of the solar panel are obtained, and the total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix. This achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment, thereby improving the stability of satellite attitude control.

[0070] Figure 1 This is a flow chart of a method for calculating the satellite inertia matrix under abnormal satellite solar panel deployment according to an embodiment of the present application. Figure 1 As shown, the method for calculating the satellite inertia matrix when the satellite solar panels are deployed abnormally in the embodiment of the present application may specifically include:

[0071] S101 : monitoring the deployment state of a solar panel of a satellite in real time based on a preset sensor, obtaining deployment state data of the solar panel, and determining the deployment state of the solar panel based on the deployment state data.

[0072] S102 : In response to an abnormal deployment of the solar panel, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panel.

[0073] S103 . Calculate the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix.

[0074] In the above-mentioned method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment, when the abnormality of solar panel deployment is monitored in real time, the first inertia matrix of the satellite body and the second inertia matrix of the solar panel are obtained, and the total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix. This achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment, thereby improving the stability of satellite attitude control.

[0075] The above exemplary steps of the embodiment of the present application are described below with reference to specific examples:

[0076] S101 , monitoring the deployment state of the solar panel of the satellite in real time based on a preset sensor, obtaining deployment state data of the solar panel, and determining the deployment state of the solar panel based on the deployment state data.

[0077] In the embodiments of the present application, the sensor includes at least a position sensor, an optical sensor, or a current sensor. The deployment status includes abnormal deployment (less than 180 degrees) and fully deployed (180 degrees). Abnormal deployment includes at least incomplete deployment, asymmetric deployment, and deployment jamming. The deployment status data is data representing the deployment status, such as the deployment angle. The deployment status of the satellite's solar panels is monitored in real time using real-time sensors to obtain the deployment status data. The deployment status of the solar panels is then determined based on the deployment status data for subsequent processing. For example, the deployment status of the solar panels is determined based on the deployment angle. When the deployment angle is 180 degrees, it indicates full deployment, i.e., a normal deployment state. When the deployment angle is less than 180 degrees, it indicates an abnormal deployment state.

[0078] S102 , in response to an abnormal deployment of the solar panel, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panel.

[0079] In an embodiment of the present application, the second inertia matrix represents the rotational inertia matrix of the solar panel. When it is determined that the solar panel is deployed abnormally according to the deployment status data obtained in step S101, the first inertia matrix of the satellite body and the second inertia matrix of the solar panel are obtained for subsequent processing.

[0080] S103 , calculating a total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix.

[0081] In an embodiment of the present application, the total inertia matrix is ​​the inertia matrix, or mass matrix, of the satellite when the satellite solar panels are abnormally deployed. The total inertia matrix of the satellite is calculated based on the first inertia matrix of the satellite body and the second inertia matrix of the solar panels obtained in step S102.

[0082] The total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix using the following formula:

[0083]

[0084] in, is the total inertia matrix of the satellite, is the first inertia matrix of the satellite body, is the second inertia matrix of the solar panel.

[0085] It should be noted that It is a preset value and does not change with the state of the solar panel, that is, a fixed value. It will change when the solar sail is deployed abnormally. At this time, in order to ensure the accuracy of the calculated total inertia matrix of the satellite, this application needs to determine a new accurate Therefore, the core of this application is to calculate , calculate Then, the new satellite total inertia matrix can be obtained .

[0086] In this way, the real-time and accurate satellite inertia matrix under abnormal deployment of the satellite solar panels is obtained, ensuring that the satellite inertia matrix calculated is accurate regardless of whether the solar panels are deployed abnormally, thereby improving the stability of satellite attitude control.

[0087] The embodiment of the present application provides a method for calculating the satellite inertia matrix under abnormal deployment of a satellite solar panel. The method monitors the deployment status of the satellite's solar panels in real time based on a preset sensor, obtains deployment status data of the solar panels, and determines the deployment status of the solar panels based on the deployment status data. In response to the abnormal deployment of the solar panels, the first inertia matrix of the satellite body and the second inertia matrix of the solar panels are obtained, and the total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix. The method for calculating the satellite inertia matrix under abnormal deployment of a satellite solar panel provided by the present application, when the abnormal deployment of the solar panel is detected in real time, obtains the first inertia matrix of the satellite body and the second inertia matrix of the solar panels, and calculates the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix, thereby achieving a more real-time and accurate calculation of the satellite's inertia matrix under abnormal deployment of the solar panels, thereby improving the stability of the satellite attitude control.

[0088] Furthermore, when the solar sail panel is a dual-axis controlled sail panel, the solar sail panel includes a first controlled rotation axis and a second controlled rotation axis. Figure 2 As shown, the "obtaining the second inertia matrix of the solar panel" in step S102 in the above embodiment may include the following steps:

[0089] S201, determining a satellite body coordinate system of the satellite and a solar panel fixed coordinate system of the solar panel.

[0090] In the embodiment of the present application, the satellite body coordinate system of the satellite and the sailboard fixed coordinate system of the solar sailboard are determined for subsequent processing. Figure 3 As shown, it represents the satellite body coordinate system, such as Figure 4 As shown, it represents the sail panel fixed coordinate system (of one of the two solar sail panels).

[0091] S202 , based on the satellite body coordinate system and the sailboard fixed coordinate system, obtain a first position of the second controlled rotation axis of the solar sailboard in the satellite body coordinate system and a second position of the second controlled rotation axis of the solar sailboard in the sailboard fixed coordinate system.

[0092] In an embodiment of the present application, based on the satellite body coordinate system and the sail panel fixed coordinate system in step S201, the first position (x1, y1, z1) of the second control rotation axis of the solar sail panel in the satellite body coordinate system and the second position (x2, y2, z2) of the second control rotation axis of the solar sail panel in the sail panel fixed coordinate system are obtained for subsequent processing.

[0093] S203 , determining a first rotation angle of a first control rotation axis and a second rotation angle of a second control rotation axis of the solar panel, and acquiring a second inertia matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle.

[0094] In an embodiment of the present application, a first rotation angle (denoted by α) of a first control rotation axis and a second rotation angle (denoted by β) of a second control rotation axis of a solar panel are determined, and a second inertia matrix of the solar panel is obtained based on the first position (x1, y1, z1), the second position (x2, y2, z2), the first rotation angle α and the second rotation angle β.

[0095] Among them, in the satellite body coordinate system, the rotation direction of the first control rotation axis of the solar sail panel is consistent with the y-axis of the satellite body coordinate system; in the satellite body coordinate system, the rotation direction of the second control rotation axis of the solar sail panel is consistent with the x-axis of the sail panel fixed coordinate system; the first position does not change with the rotation angle of the first control rotation axis, and the second position does not change with the rotation angle of the second control rotation axis.

[0096] In some embodiments, a posture transformation matrix of the solar panel is obtained based on the first position, the second position, the first rotation angle, and the second rotation angle; an initial second inertia matrix of the solar panel in the panel fixed coordinate system is determined, and a second inertia matrix of the solar panel is calculated based on the initial second inertia matrix.

[0097] Optionally, when obtaining the posture transformation matrix of the solar panel based on the first position, the second position, the first rotation angle and the second rotation angle, the panel fixed coordinate system is overlapped with the satellite body coordinate system to obtain the current panel fixed coordinate system, and based on the first position, the solar panel is translated along the current panel fixed coordinate system by the displacement of the first position; the solar panel is rotated around the y-axis of the panel fixed coordinate system by the first rotation angle, and the solar panel is rotated around the x-axis of the panel fixed coordinate system by the second rotation angle; based on the second position, the solar panel is translated along the negative direction of the current panel fixed coordinate system by the displacement of the second position to obtain the posture transformation matrix of the solar panel. For example, Figure 5 As shown, through Figure 5 The displacement and rotation shown above are used to obtain the position transformation matrix of the solar panel.

[0098] It should be noted that the posture conversion matrix of the solar panel is obtained according to the following formula, that is, the posture conversion matrix of the solar panel based on the first position, the second position, the first rotation angle and the second rotation angle is obtained by the following formula:

[0099]

[0100] in, represents the position transformation matrix of the solar panel, α represents the first rotation angle, β represents the second rotation angle, (x1, y1, z1) represents the first position, and (x2, y2, z2) represents the second position; represents the displacement of the first position, Indicates the first rotation angle, Indicates the second rotation angle, Indicates the displacement of the second position in the negative direction.

[0101] It should be noted that the second inertia matrix of the solar panel is obtained according to the following formula, that is, the second inertia matrix of the solar panel is obtained by calculation based on the initial second inertia matrix, which is implemented by the following formula:

[0102]

[0103] in, represents the second moment of inertia matrix, represents the initial second inertia matrix, represents the posture conversion matrix of the solar panel, and m represents the mass of a single solar panel.

[0104] Therefore, the present application dynamically calculates the satellite's inertia matrix according to the real-time deployment status of the solar panels, ensuring the accuracy of the inertia matrix calculation under abnormal conditions. At the same time, regardless of whether the solar panels are fully deployed or what kind of abnormality occurs, the satellite's inertia matrix can be accurately calculated. The inertia matrix can accurately reflect the inertial characteristics of the satellite, ensuring the stability of the satellite's attitude control.

[0105] It should be noted that the present application provides a method for calculating the satellite inertia matrix when the satellite solar panels are abnormally deployed, that is, a method for calculating the satellite inertia matrix, that is, a method for calculating the satellite inertia when the satellite solar panels are abnormally deployed.

[0106] Figure 6 Schematic diagram of the structure of the satellite inertia matrix calculation device under abnormal satellite solar panel deployment according to the embodiment of the present application. Figure 6 As shown, the satellite inertia matrix calculation device 600 under abnormal satellite solar panel deployment in an embodiment of the present application may specifically include:

[0107] The first acquisition module 601 is used to monitor the deployment status of the satellite's solar panels in real time based on preset sensors, obtain deployment status data of the solar panels, and determine the deployment status of the solar panels based on the deployment status data; wherein the deployment status includes deployment abnormality and complete deployment; the deployment abnormality at least includes incomplete deployment, asymmetric deployment, and deployment jamming.

[0108] The second acquisition module 602 is configured to acquire a first inertia matrix of the satellite body and a second inertia matrix of the solar panel in response to an abnormal deployment of the solar panel; wherein the second inertia matrix represents a rotational inertia matrix of the solar panel.

[0109] The third acquisition module 603 is configured to calculate a total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix; wherein the total inertia matrix is ​​the inertia matrix of the satellite when the satellite solar panels are abnormally deployed.

[0110] In a possible embodiment, when the solar panel is a dual-axis controlled panel, the solar panel includes a first controlled rotation axis and a second controlled rotation axis; the second acquisition module is specifically configured to:

[0111] Determine the satellite body coordinate system of the satellite and the sailboard fixed coordinate system of the solar sailboard;

[0112] Based on the satellite body coordinate system and the sail panel fixed coordinate system, a first position of the second control rotation axis of the solar sail panel in the satellite body coordinate system and a second position of the second control rotation axis of the solar sail panel in the sail panel fixed coordinate system are obtained;

[0113] A first rotation angle of a first control rotation axis and a second rotation angle of a second control rotation axis of the solar panel are determined, and a second inertia matrix of the solar panel is obtained based on the first position, the second position, the first rotation angle and the second rotation angle.

[0114] In a possible implementation, the second acquisition module is specifically configured to:

[0115] Acquire a position conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle;

[0116] An initial second inertia matrix of the solar panel in the panel fixed coordinate system is determined, and a second inertia matrix of the solar panel is obtained by calculation based on the initial second inertia matrix.

[0117] In one possible embodiment, in the satellite body coordinate system, the rotation direction of the first control rotation axis of the solar sail panel is consistent with the y-axis of the satellite body coordinate system; in the satellite body coordinate system, the rotation direction of the second control rotation axis of the solar sail panel is consistent with the x-axis of the sail panel fixed coordinate system;

[0118] The first position does not change with the rotation angle of the first control rotation axis, and the second position does not change with the rotation angle of the second control rotation axis.

[0119] In a possible implementation, the second acquisition module is specifically configured to:

[0120] Overlapping the sailboard fixed coordinate system with the satellite body coordinate system to obtain a current sailboard fixed coordinate system, and translating the solar sailboard along the current sailboard fixed coordinate system by the displacement of the first position based on the first position;

[0121] Rotating the solar panel around the y-axis of the sail panel fixed coordinate system by a first rotation angle, and rotating the solar panel around the x-axis of the sail panel fixed coordinate system by a second rotation angle;

[0122] Based on the second position, the solar panel is translated along the negative direction of the current panel fixed coordinate system by the displacement of the second position to obtain a position and attitude transformation matrix of the solar panel.

[0123] In a possible implementation, the posture transformation matrix of the solar panel is obtained according to the following formula:

[0124]

[0125] in, represents the position transformation matrix of the solar panel, α represents the first rotation angle, β represents the second rotation angle, (x1, y1, z1) represents the first position, and (x2, y2, z2) represents the second position; represents the displacement of the first position, Indicates the first rotation angle, Indicates the second rotation angle, Indicates the displacement of the second position in the negative direction.

[0126] In a possible implementation, the second inertia matrix of the solar panel is calculated based on the initial second inertia matrix, which is implemented by the following formula:

[0127]

[0128] in, represents the second moment of inertia matrix, represents the initial second inertia matrix, represents the position transformation matrix of the solar panel, and m represents the mass of a single solar panel.

[0129] The satellite inertia matrix calculation device under abnormal satellite solar panel deployment provided in an embodiment of the present application monitors the deployment status of the satellite's solar panels in real time based on a preset sensor, obtains deployment status data of the solar panels, and determines the deployment status of the solar panels based on the deployment status data. In response to the abnormal solar panel deployment, the first inertia matrix of the satellite body and the second inertia matrix of the solar panels are obtained, and the total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix. The satellite inertia matrix calculation device under abnormal satellite solar panel deployment of the present application obtains the first inertia matrix of the satellite body and the second inertia matrix of the solar panels when the abnormal solar panel deployment is detected in real time, and calculates the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. This achieves more real-time and accurate calculation of the satellite's inertia matrix under abnormal solar panel deployment conditions, thereby improving the stability of satellite attitude control.

[0130] like Figure 7 As shown, an electronic device 700 provided in an embodiment of the present application includes: a processor 701, a memory 702 and a bus, wherein the memory 702 stores machine-readable instructions executable by the processor 701. When the electronic device is running, the processor 701 communicates with the memory 702 through the bus, and the processor 701 executes the machine-readable instructions to perform the steps of the satellite inertia matrix calculation method under the above-mentioned abnormal deployment of the satellite solar panel.

[0131] Specifically, the above-mentioned memory 702 and processor 701 can be general-purpose memory and processor, which are not specifically limited here. When the processor 701 runs the computer program stored in the memory 702, it can execute the satellite inertia matrix calculation method under the above-mentioned satellite solar panel deployment abnormality.

[0132] Corresponding to the above-mentioned method for calculating the satellite inertia matrix under abnormal satellite solar panel deployment, an embodiment of the present application also provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the steps of the above-mentioned method for calculating the satellite inertia matrix under abnormal satellite solar panel deployment are executed.

[0133] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working process of the system and device described above can refer to the corresponding process in the method embodiment, and will not be repeated in this application. In the several embodiments provided in this application, it should be understood that the disclosed system, device and method can be implemented in other ways. The device embodiments described above are merely schematic. For example, the division of the modules is only a logical function division. There may be other division methods in actual implementation. For example, multiple modules or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some communication interfaces, indirect coupling or communication connection of devices or modules, which can be electrical, mechanical or other forms.

[0134] The modules described as separate components may or may not be physically separate, and the components shown as modules may or may not be physical units, that is, they may be located in one place or distributed across multiple network elements. Some or all of these units may be selected to achieve the purpose of this embodiment according to actual needs.

[0135] In addition, each functional unit in each embodiment of the present application may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit.

[0136] If the functions are implemented as software functional units and sold or used as independent products, they can be stored in a non-volatile computer-readable storage medium that is executable by a processor. Based on this understanding, the technical solution of this application, or the portion that contributes to the prior art, or the portion of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions for enabling a computer device (which can be a personal computer, server, or network device, etc.) to execute all or part of the steps of the deployment method described in each embodiment of this application. The aforementioned storage medium includes various media that can store program code, such as a USB flash drive, a mobile hard drive, ROM, RAM, a magnetic disk, or an optical disk.

[0137] The above are only specific embodiments of the present application, but the scope of protection of this application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of this application. Therefore, the scope of protection of this application should be based on the scope of protection of the claims.

Claims

1. A method for calculating the satellite inertia matrix under abnormal deployment of satellite solar panels, characterized in that: The method comprises: Based on a preset sensor, the deployment state of the solar panel of the satellite is monitored in real time to obtain deployment state data of the solar panel, and the deployment state of the solar panel is determined based on the deployment state data; wherein the deployment state includes deployment abnormality and complete deployment; and the deployment abnormality includes at least incomplete deployment, asymmetric deployment, and deployment jam; In response to the abnormal deployment of the solar sail panel, a first inertia matrix of the satellite body and a second inertia matrix of the solar sail panel are obtained; wherein the second inertia matrix represents the rotational inertia matrix of the solar sail panel; when the solar sail panel is a dual-axis controlled sail panel, the solar sail panel includes a first control rotation axis and a second control rotation axis; obtaining the second inertia matrix of the solar sail panel includes: determining a satellite body coordinate system of the satellite and a sail panel fixed coordinate system of the solar sail panel; obtaining a first position of the second control rotation axis of the solar sail panel in the satellite body coordinate system and a second position of the second control rotation axis of the solar sail panel in the sail panel fixed coordinate system based on the satellite body coordinate system and the sail panel fixed coordinate system; determining a first rotation angle of the first control rotation axis and a second rotation angle of the second control rotation axis of the solar sail panel, and obtaining the second inertia matrix of the solar sail panel based on the first position, the second position, the first rotation angle and the second rotation angle; The total inertia matrix of the satellite is calculated based on the first inertia matrix and the second inertia matrix; wherein the total inertia matrix is ​​the inertia matrix of the satellite when the satellite solar panels are abnormally deployed.

2. The method according to claim 1, characterized in that The acquiring a second inertia matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle includes: Acquire a posture conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle; An initial second inertia matrix of the solar panel in the panel-fixed coordinate system is determined, and a second inertia matrix of the solar panel is obtained by calculation based on the initial second inertia matrix.

3. The method according to claim 2, characterized in that In the satellite body coordinate system, the rotation direction of the first controlled rotation axis of the solar sail panel is consistent with the y-axis of the satellite body coordinate system; in the satellite body coordinate system, the rotation direction of the second controlled rotation axis of the solar sail panel is consistent with the x-axis of the sail panel fixed coordinate system; The first position does not change with the rotation angle of the first control rotation axis, and the second position does not change with the rotation angle of the second control rotation axis.

4. The method according to claim 3, characterized in that The acquiring the posture conversion matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle includes: Overlapping the sailboard fixed coordinate system with the satellite body coordinate system to obtain a current sailboard fixed coordinate system, and translating the solar sailboard along the current sailboard fixed coordinate system by a displacement of a first position based on the first position; Rotating the solar panel around the y-axis of the sailboard fixed coordinate system by a first rotation angle, and rotating the solar panel around the x-axis of the sailboard fixed coordinate system by a second rotation angle; Based on the second position, the solar panel is translated along the negative direction of the current panel fixed coordinate system by the displacement of the second position to obtain a position and posture transformation matrix of the solar panel.

5. The method according to claim 4, characterized in that The posture conversion matrix of the solar panel is obtained according to the following formula: in, represents the position conversion matrix of the solar panel, α represents the first rotation angle, β represents the second rotation angle, (x1, y1, z1) represents the first position, and (x2, y2, z2) represents the second position; represents the displacement of the first position, Indicates the first rotation angle, Indicates the second rotation angle, Indicates the displacement of the second position in the negative direction.

6. The method according to claim 5, characterized in that The second inertia matrix of the solar panel is calculated based on the initial second inertia matrix and is achieved by the following formula: in, represents the second moment of inertia matrix, represents the initial second inertia matrix, represents the posture conversion matrix of the solar panel, and m represents the mass of a single solar panel.

7. A satellite inertia matrix calculation device under abnormal satellite solar panel deployment, characterized in that: The device comprises: a first acquisition module, configured to monitor the deployment state of the satellite's solar panels in real time based on a preset sensor, obtain deployment state data of the solar panels, and determine the deployment state of the solar panels based on the deployment state data; wherein the deployment state includes deployment abnormality and complete deployment; and the deployment abnormality includes at least incomplete deployment, asymmetric deployment, and deployment jam; a second acquisition module, configured to acquire a first inertia matrix of a satellite body and a second inertia matrix of the solar panel in response to an abnormal deployment of the solar panel; wherein the second inertia matrix represents a rotational inertia matrix of the solar panel; when the solar panel is a dual-axis controlled panel, the solar panel includes a first control rotation axis and a second control rotation axis; the second acquisition module is specifically configured to determine a satellite body coordinate system of the satellite and a panel fixed coordinate system of the solar panel; based on the satellite body coordinate system and the panel fixed coordinate system, obtain a first position of the second control rotation axis of the solar panel in the satellite body coordinate system and a second position of the second control rotation axis of the solar panel in the panel fixed coordinate system; determine a first rotation angle of the first control rotation axis and a second rotation angle of the second control rotation axis of the solar panel, and acquire the second inertia matrix of the solar panel based on the first position, the second position, the first rotation angle, and the second rotation angle; The third acquisition module is configured to calculate a total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix; wherein the total inertia matrix is ​​the inertia matrix of the satellite when the satellite solar panels are abnormally deployed.

8. An electronic device, characterized in that: include: A processor, a memory, and a bus, wherein the memory stores machine-readable instructions executable by the processor. When the electronic device is running, the processor and the memory communicate via the bus. When the machine-readable instructions are executed by the processor, the steps of the method for calculating the satellite inertia matrix under abnormal satellite solar panel deployment as described in any one of claims 1 to 6 are performed.

9. A computer-readable storage medium, characterized in that The computer-readable storage medium stores a computer program, which, when executed by a processor, executes the steps of the method for calculating the satellite inertia matrix under abnormal deployment of a satellite solar panel as described in any one of claims 1 to 6.

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