A method and system for controlling the attitude of a low earth orbit satellite, a computer device and a storage medium

Abstract

The application discloses a kind of orbit satellite attitude control method and system, computer equipment, storage medium, belong to satellite control technical field, in the control method of the present application, when the satellite needing to carry out attitude maneuvering of orbit control, PID closed-loop control theory is used to carry out attitude control;It also provides the control system, computer equipment and readable storage medium used;The method of the present application can realize fast, small overshoot attitude maneuvering to orbit satellite, engineering implementation operability is strong, especially suitable for low-orbit communication satellite in orbit control time's attitude maneuvering.

Description

Technical Field

[0001] This invention belongs to the field of satellite control technology, specifically relating to an attitude control method and system for orbiting satellites, computer equipment, and storage medium. Background Technology

[0002] Satellite attitude control refers to the process of adjusting and maintaining a satellite's spatial position and orientation (attitude) in orbit through a series of control methods and technologies. Its purpose is to ensure that the satellite maintains the correct pointing during mission execution, thereby ensuring data accuracy, signal stability, and satellite safety. For communication satellites, attitude control is necessary to ensure that their antennas are always pointed towards the ground receiving station for stable signal transmission. Remote sensing satellites require precise attitude control to acquire high-quality ground images and data for environmental monitoring, agricultural management, and urban planning. In recent years, with the increasing intensity of space exploration and research activities, the method of launching large spacecraft with a single rocket, limited by rocket carrying capacity and envelope size, is gradually failing to meet diverse operational objectives and performance requirements. Technologies for on-orbit assembly of intelligent swarm spacecraft and collaborative mission completion are gradually developing. This requires precise attitude control of the spacecraft's orbital system to enable inter-satellite information exchange and mission collaboration.

[0003] When a spacecraft is in orbit, its attitude deviates from its expected value due to internal and external disturbance torques. Attitude stabilization control is required to maintain this attitude. Simultaneously, as the spacecraft's position relative to its target changes during orbital motion, communication antennas need to be aligned with signal receiving devices to facilitate communication between the spacecraft and the ground, as well as between spacecraft themselves. This necessitates attitude maneuvering control. The task of attitude stabilization control is to design reasonable control laws and select appropriate actuators based on the determined attitude information, so as to stabilize the attitude near the desired value and meet the attitude performance requirements. Attitude maneuvering refers to the process of reorienting the spacecraft from a known attitude to another desired attitude.

[0004] For short-duration spacecraft, jet control is the sole attitude control actuator, relying on the reaction force generated by the propellant expelled by the jet propulsion system to form control torque. For long-life spacecraft, jet control instruments are used for short periods, generally only for attitude control during periods of high disturbance (such as during orbital control or depressurization), attitude maneuvers, and momentum device unloading. Long-term flight relies on momentum exchange devices. Attitude control is achieved by exchanging angular momentum between the spacecraft's internal momentum exchange device and the spacecraft's own momentum. For various long-life, high-precision three-axis stabilized spacecraft, momentum exchange devices are primarily used for attitude control. Common momentum exchange devices include momentum wheels and control moment gyroscopes. Compared to momentum wheels, control moment gyroscopes provide larger attitude control torque and angular momentum capacity. For large spacecraft with long-term orbital flight, which have large mass and inertia, the requirements for attitude control torque and angular momentum capacity are correspondingly large, and control moment gyroscopes are mainly used for attitude control.

[0005] Some scientific exploration satellites only require knowledge of the time when space or atmospheric physics parameters were acquired, the satellite's orbital position, and instantaneous attitude for post-processing of the data. These spacecraft do not require attitude control but do require attitude determination, typically with an accuracy of a few degrees to a tenth of a degree. Communication and broadcast satellites require antenna pointing accuracy of approximately one-tenth of the beamwidth. Earth observation satellites (reconnaissance satellites, Earth resource satellites, and meteorological satellites, etc.) need to distinguish, identify, and locate targets, requiring high attitude accuracy (a tenth of a degree) and attitude stability (a few arcseconds per second). Astronomical satellites require extremely high attitude accuracy (a few arcseconds) and attitude stability (on the order of 10⁻³ arcseconds per second).

[0006] The prerequisite for spacecraft attitude control is accurate attitude description. Common attitude description methods include rotation matrices, Euler angles, Euler quaternions, and modified Rodrigues parameters (MRPs). Euler angles and MRPs have singularities, and Euler quaternions have the drawback of non-uniqueness. Rotation matrices can globally and uniquely describe the attitude of a rigid body. In early research, the three-dimensional special orthogonal matrix group SO(3) formed by rotation matrices was a non-Euclidean manifold. Due to the imperfect development of its model stability theory, other parameterized models were usually chosen for control design. Currently, research on Lie groups is gradually developing, and control methods based on SO(3) exhibit high control accuracy and fast control performance.

[0007] Low-Earth orbit (LEO) communication satellites often carry large-area solar panels, which typically exhibit low fundamental frequency, weak damping, and dense modal characteristics. To avoid the impact of attitude maneuvers on these panels while simultaneously achieving rapid maneuvers, the ideal requirement for attitude maneuvers is zero overshoot and rapid speed. Currently, large-angle attitude maneuvers for satellites are often described using quaternions and controlled using PID closed-loop control theory. However, research on simple and efficient control methods for large-angle attitude maneuvers during periodic orbit control of LEO communication satellites is limited. Summary of the Invention

[0008] This invention provides a method and system for attitude control of orbiting satellites, a computer device, and a storage medium, which overcomes the shortcomings of the prior art.

[0009] To achieve the above objectives, the present invention adopts the following technical solution:

[0010] A method for attitude control of an orbiting satellite includes the following steps:

[0011] S1: When a satellite undergoing orbit control needs to perform attitude maneuvers, calculate the error angle between the satellite's current attitude and the target attitude; determine whether the error angle is greater than the designed error threshold; if the error angle is greater than the error threshold, perform attitude maneuver path planning for the satellite. The attitude maneuver path planning results include: the time period occupied by the process of adjusting the satellite's attitude from the initial moment of the attitude maneuver to the final desired attitude, the attitude angular acceleration planning results during this time period, the attitude angular velocity planning results during this time period, and the attitude angular position planning results during this time period.

[0012] S2: Based on the attitude maneuver path planning results, the first control torque is calculated using the SO(3) controller; if the error angle is less than the error threshold, the first control torque is directly calculated using the SO(3) controller; based on the first control torque, it is determined whether the error angle is greater than the integral threshold; if the error angle is greater than the integral threshold, the first control torque is used as the control torque; if the error angle is less than the integral threshold, an error angle integral term is introduced into the first control torque as the second control torque, wherein the satellite is equipped with a control torque gyroscope;

[0013] S3: Compare the satellite's attitude angle with the attitude error threshold and the integral threshold to obtain the calculation method for the control torque; based on the calculation method for the control torque, control the attitude control actuator on the satellite to generate the second desired torque;

[0014] S4: Based on the configuration position of the control moment gyroscope on the satellite and the target parameters for orbit control of the satellite, adjust the torque direction of the control moment gyroscope to the desired torque direction, which corresponds to the final desired attitude.

[0015] An attitude control system for an orbiting satellite includes: an attitude maneuver path planning module, a control torque direction adjustment module, an attitude maneuver control quantity calculation module, an attitude maneuver control method switching module, and a control module;

[0016] The attitude maneuver path planning module is used to: when a satellite undergoing orbit control needs to perform attitude maneuvers, based on the trapezoidal attitude maneuver path planning method, combined with the satellite's attitude at the initial moment of the attitude maneuver and the final desired attitude, to plan the attitude maneuver path for the satellite and obtain the attitude maneuver path planning result. The satellite is equipped with a control moment gyroscope.

[0017] The control torque direction adjustment module is used to: adjust the control torque direction of the control torque gyroscope to the desired control torque direction according to the configuration position of the control torque gyroscope on the satellite and the target parameters for orbit control of the satellite. The desired control torque direction corresponds to the final desired attitude.

[0018] The attitude maneuver control quantity calculation module is used to: calculate the first desired torque of the satellite during attitude control based on the attitude maneuver path planning results or attitude error; and calculate the second desired torque of the satellite during attitude control.

[0019] The attitude maneuver control method switching module is used to: compare the satellite's attitude error angle with the attitude error threshold and the integral threshold, switch the control method, and obtain the switching result;

[0020] The control module is used to: control the attitude control actuators on the satellite to generate the second desired torque based on the calculated second desired torque result.

[0021] A computer device includes a processor coupled to a memory storing at least one computer program, which is loaded and executed by the processor to enable the computer device to implement any of the above-mentioned satellite attitude maneuver control methods.

[0022] A computer-readable storage medium storing at least one computer program, which is loaded and executed by a processor to enable a computer to implement any of the above-mentioned satellite attitude maneuver control methods.

[0023] Beneficial effects: This invention provides a method and system for attitude control of orbital satellites, computer equipment, and storage medium. When a satellite undergoing orbital control needs to perform attitude maneuvers, the PID closed-loop control theory is used for attitude control, which enables the satellite to perform fast and low-overshoot attitude maneuvers. It is highly operable in engineering implementation and is especially suitable for attitude maneuvers of low-Earth orbit communication satellites during attitude control. Attached Figure Description

[0024] Figure 1 This is a flowchart illustrating the attitude maneuver control method for a satellite in an embodiment of the present invention.

[0025] Figure 2 This is a schematic diagram of the attitude maneuver path planning results in an embodiment of the present invention;

[0026] Figure 3 This is a schematic diagram of the attitude angular acceleration at each moment in the attitude angular acceleration planning results in an embodiment of the present invention;

[0027] Figure 4 This is a schematic diagram of the attitude angular velocity at each moment in the attitude angular velocity planning results in an embodiment of the present invention;

[0028] Figure 5 This is a schematic diagram of the attitude angle position at each moment in the attitude angle position planning result of the embodiment of the present invention;

[0029] Figure 6 This is a schematic diagram of the layout of the electric propulsion thruster in an embodiment of the present invention;

[0030] Figure 7 This is a schematic diagram of the attitude maneuver control system of a satellite in an embodiment of the present invention;

[0031] Figure 8 This is a schematic diagram of the structure of a computer device according to an embodiment of the present invention. Detailed Implementation

[0032] The embodiments of the present invention will now be described in further detail with reference to the accompanying drawings and specific examples.

[0033] like Figure 1 As shown, a method for satellite orientation and attitude control in its orbital system includes the following steps:

[0034] S1. When a satellite undergoing attitude control needs to perform attitude maneuvers, the attitude error angle is calculated based on the initial attitude of the satellite's attitude maneuver and the final desired attitude. The magnitude of the error angle and the attitude error threshold are judged, and it is selected whether to perform attitude maneuver planning. The SO(3) controller is used to calculate the first desired torque, and the trapezoidal attitude maneuver path planning method is used to perform attitude maneuver planning. The trapezoidal attitude maneuver path planning method combines the initial attitude of the satellite's attitude maneuver and the final desired attitude to perform attitude maneuver path planning for the satellite and obtain the attitude maneuver path planning result. The satellite is equipped with a control torque gyroscope.

[0035] The attitude maneuver path planning results include: the time period occupied by the process of adjusting the satellite's attitude from the initial moment of the attitude maneuver to the final desired attitude; the attitude angular acceleration planning results, attitude angular velocity planning results, and attitude angular position planning results within this time period; this time period is divided into three sub-time periods, in chronological order: the constant acceleration motion period. Uniform motion for a period of time Heheng accelerated movement period , , and It was also planned using the trapezoidal attitude maneuver path planning method.

[0036] The attitude angular acceleration planning result includes: the attitude angular acceleration at each moment within the time period; the attitude angular velocity planning result includes: the attitude angular velocity at each moment within the time period; the attitude angular position planning result includes: the attitude angular position at each moment within the time period. The attitude angular acceleration is denoted as... Let the attitude angular velocity be denoted as The attitude angle position is denoted as The changes in attitude angular acceleration, attitude angular velocity, and attitude angular position within the three sub-time periods are as follows: Figure 2 As shown, the details are as follows:

[0037] 1) The attitude angle error is:

[0038] ,

[0039] in This describes the rotation matrix corresponding to the spacecraft's current attitude. For matrix transpose, This describes the rotation matrix corresponding to the spacecraft target attitude. This is the inverse operation for antisymmetric matrices.

[0040] 2) The angular velocity error is:

[0041] ,

[0042] in The angular velocity of the spacecraft, This represents the target angular velocity of the spacecraft.

[0043] 3) The attitude angular acceleration planning result is:

[0044] ,

[0045] in For the angular acceleration planning parameters of the design, To control the three time periods, in At time t, the angular acceleration is planned as ,exist At time t, the angular acceleration is planned to be 0. At time t, the angular acceleration is planned as .

[0046] 4) The attitude angular velocity planning result is:

[0047] ;

[0048] 5) The attitude angle position planning result is:

[0049] ,

[0050] In the attitude angular velocity planning results and attitude angular position planning results, The initial attitude angular velocity is the initial attitude angular velocity at the start of the satellite's attitude maneuver. To determine the initial attitude angle position at the start of the planning, i.e., the attitude angle position at the initial moment of the satellite's attitude maneuver, when time is... The attitude angle planning result is the angular position of the satellite's final desired attitude. , The maximum value of attitude angular acceleration during the time interval from the initial attitude adjustment of the satellite to the final desired attitude. Let be the minimum attitude angular velocity during the time interval taken from the initial attitude adjustment of the satellite to the final desired attitude, where , The settings are based on the satellite's attitude maneuverability.

[0051] It should be noted that in the attitude angle acceleration planning results It is a holistic representation. "in "This cannot be interpreted as 'the maximum value of attitude angular acceleration during the time period from the initial attitude adjustment of the satellite to the final desired attitude during the attitude maneuver'." ".

[0052] Figure 2 middle The maximum angular velocity of the attitude during the time interval from the initial attitude adjustment of the satellite to the final desired attitude. It was obtained by the trapezoidal attitude maneuver path planning method.

[0053] Based on the attitude maneuver path planning results, calculate the first desired torque of the satellite during attitude control:

[0054] ,

[0055] in, This represents the first desired torque, which can be understood as the torque generated by the actuators required for open-loop attitude control of the satellite. For planning Axial attitude angular acceleration, For planning Axial attitude angular acceleration, For planning Axial attitude angular acceleration, , , These are the satellite's principal inertia parameters.

[0056] When the attitude angle error is less than the attitude error threshold, based on PID control theory, the first desired torque for satellite attitude control is calculated. The first desired torque is: , The differential coefficients are... This is the proportionality coefficient;

[0057] S2. Calculate the attitude error angle based on the attitude during satellite attitude maneuver and the final desired attitude, determine the magnitude of the error angle and the integration threshold, select whether to introduce the error angle integration term in the controller, and use the SO(3) controller to calculate the second desired torque.

[0058] The attitude error integral term and the second control torque are as follows:

[0059] 1) The attitude error integral term is:

[0060] ,

[0061] 2) Introducing the error integral term to calculate the second desired control torque is:

[0062] ,

[0063] ,

[0064] In the above formula, R is the rotation matrix of the i-th spacecraft. i The rotation matrix R of the j-th spacecraft j The error between them, k I It is the constant coefficient of the integral term.

[0065] 3) If the attitude error angle is less than the integral threshold, the second desired control torque is:

[0066] ,

[0067] 4) Attitude control requires the generation of coordinates in the satellite's body coordinate system by actuators. axis, axis, The control torques of the shafts are denoted as follows: , , .

[0068] S3. Based on the results of the switching control method, control the attitude control actuator on the satellite to generate a second desired torque to adjust the satellite's attitude at the initial moment of attitude maneuvering to the final desired attitude; specifically including the following steps:

[0069] Determine the layout of the control moment gyroscope in the satellite

[0070] The layout principles for control moment gyroscopes on satellites include:

[0071] 1) The control moment gyroscope is equipped with an electric propulsion thruster, which is a Hall thruster, used to unload the control moment gyroscope;

[0072] 2) The principle for the layout of gas cylinders in the propulsion system used in conjunction with electric propulsion thrusters is to select the position with the smallest distance between the gas cylinder's center of mass and the satellite's center of mass;

[0073] 3) The electric propulsion thrusters are symmetrically positioned on the satellite, and the thrust is generated past the satellite's center of mass without generating additional thrust components. They are arranged along the XYZ axes of the satellite's body coordinate system. When the control torque of the control torque gyroscope is unloaded by the electric propulsion thrusters, the control torque gyroscope is usually at 90% of its limit speed.

[0074] 4) The layout of other equipment on the satellite must avoid the influence range of the plume angle of the electric propulsion thruster.

[0075] Optionally, in the above technical solution, the satellite is a low-orbit communication satellite.

[0076] In the above embodiments, although the steps are numbered S1, S2, etc., they are only specific embodiments given by the present invention. Those skilled in the art can adjust the execution order of S1, S2, etc. according to the actual situation, which is also within the protection scope of the present invention. It can be understood that in some embodiments, some or all of the above embodiments may be included.

[0077] like Figure 2 As shown, the angular acceleration, angular velocity, and angular position results generated by the path planning in this embodiment exhibit a smooth transition trend, which meets the requirements of high-precision control.

[0078] For specific cases, simulations were conducted to verify the designed attitude control method.

[0079] Figure 3 The figure shows the planning result of the angular acceleration curve under the attitude control method of this invention.

[0080] Figure 4 The figure shows the planning result of the angular velocity curve under the attitude control method of this invention.

[0081] Figure 5 The figure shows the planning result of the angular position curve under the attitude control method of this invention.

[0082] like Figure 7 As shown, a satellite attitude maneuver control system 200 of this embodiment includes an attitude maneuver path planning module 201, a control torque direction adjustment module 202, an attitude maneuver control quantity calculation module 203, an attitude maneuver control method switching module 204, and a control module 205.

[0083] The attitude maneuver path planning module 201 is used to: when a satellite undergoing orbit control needs to perform attitude maneuvers, based on the trapezoidal attitude maneuver path planning method, combine the satellite's attitude at the initial moment of the attitude maneuver and the final desired attitude, to perform attitude maneuver path planning for the satellite and obtain the attitude maneuver path planning result, wherein the satellite is equipped with a control moment gyroscope.

[0084] The control torque direction adjustment module 202 is used to: adjust the control torque direction of the control torque gyroscope to the desired control torque direction according to the configuration position of the control torque gyroscope on the satellite and the target parameters for orbit control of the satellite, wherein the desired control torque direction corresponds to the final desired attitude;

[0085] The attitude maneuver control quantity calculation module 203 is used to: calculate the first desired torque of the satellite during attitude control based on the attitude maneuver path planning results or attitude error; and calculate the second desired torque of the satellite during attitude control.

[0086] The attitude maneuver control method switching module 204 is used to: compare the satellite's attitude error angle with the attitude error threshold and the integral threshold, switch the control method, and obtain the switching result;

[0087] The control module 205 is used to: control the attitude control actuator on the satellite to generate a second desired torque based on the calculated second desired torque result.

[0088] Furthermore, the attitude maneuver path planning module 201 first performs path planning, and then transmits the path planning result to the attitude maneuver control method switching module 204 to select the control method. After that, the attitude maneuver control quantity calculation module 203 calculates the desired control torque, and then, in combination with the torque gyroscope configuration, the control torque direction adjustment module 202 obtains the control signal of the control torque gyroscope. Finally, the control module 205 drives the control actuator to generate control torque according to the control signal.

[0089] Furthermore, it also includes a control moment gyroscope satellite layout design module, which is used to determine the layout of the control moment gyroscope in the satellite.

[0090] The attitude maneuver control quantity calculation module is also specifically used to: calculate the second desired torque of the satellite during attitude control based on PID control theory.

[0091] Furthermore, the system provided in the above embodiments is only illustrated by the division of the above functional modules. In practical applications, the above functions can be assigned to different functional modules as needed, that is, the system can be divided into different functional modules according to the actual situation to complete all or part of the functions described above. In addition, the system and method embodiments provided in the above embodiments belong to the same concept, and their specific implementation process can be found in the method embodiments, which will not be repeated here.

[0092] like Figure 8 As shown, a computer device 300 of this embodiment includes a processor 320 coupled to a memory 310. The memory 310 stores at least one computer program 330, which is loaded and executed by the processor 320 to enable the computer device 300 to implement the above-mentioned satellite attitude maneuver control method. Specifically:

[0093] The computer device 300 can vary considerably due to differences in configuration or performance. It may include one or more processors 320 (Central Processing Units, CPUs) and one or more memories 310. The one or more memories 310 store at least one computer program 330, which is loaded and executed by the one or more processors 320 to enable the computer device 300 to implement any of the satellite attitude maneuver control methods provided in the above embodiments. Of course, the computer device 300 may also have wired or wireless network interfaces, a keyboard, and input / output interfaces for input and output. The computer device 300 may also include other components for implementing device functions, which will not be elaborated upon here.

[0094] An embodiment of the present invention provides a computer-readable storage medium storing at least one computer program, which is loaded and executed by a processor to enable the computer to implement any of the above-described satellite attitude maneuver control methods.

[0095] Alternatively, the computer-readable storage medium may be a read-only memory (ROM), a random access memory (RAM), a compact disc read-only memory (CD-ROM), magnetic tape, a floppy disk, and an optical data storage device, etc.

[0096] In this embodiment, a computer program product or computer program is also provided, which includes computer instructions stored in a computer-readable storage medium. The processor of a computer device reads the computer instructions from the computer-readable storage medium and executes the computer instructions, causing the computer device to perform any of the above-described satellite attitude maneuver control methods.

[0097] It should be noted that the terms "first," "second," etc., used in the specification and claims of this invention are used to distinguish similar objects and represent a limitation on a specific order or sequence. Where appropriate, the order of use for similar objects can be interchanged so that the embodiments of this application described herein can be implemented in an order other than that shown or described.

[0098] Those skilled in the art will recognize that this invention can be implemented as a system, method, or computer program product. Therefore, this invention can be specifically implemented in the following forms: it can be entirely hardware, entirely software (including firmware, resident software, microcode, etc.), or a combination of hardware and software, generally referred to herein as a "circuit," "module," or "system." Furthermore, in some embodiments, this invention can also be implemented as a computer program product contained in one or more computer-readable media, which includes computer-readable program code.

[0099] Any combination of one or more computer-readable media can be used. A computer-readable medium can be a computer-readable signal medium or a computer-readable storage medium. A computer-readable storage medium can be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples (a non-exhaustive list) of computer-readable storage media include: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this invention, a computer-readable storage medium can be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system, apparatus, or device.

[0100] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A method for attitude control of an orbiting satellite, characterized in that, Includes the following steps: When a satellite undergoing orbit control needs to perform attitude maneuvers, the error between the satellite's current attitude and the target attitude angle is calculated, and it is determined whether the error is greater than the designed error threshold. If the error exceeds the threshold, attitude maneuver path planning is performed on the satellite. Based on the attitude maneuver path planning results, the first control torque is calculated using the SO(3) controller; Based on the first control torque, determine whether the error is greater than the integral threshold; If the error is greater than the integral threshold, the first control torque will be used as the control torque. If the error is less than the integral threshold, then the error angle integral term is introduced into the first control torque as the second control torque; The attitude maneuver path planning results include: the time period occupied by the process of adjusting the satellite's attitude from the initial moment of the attitude maneuver to the final desired attitude; the attitude angular acceleration planning results, attitude angular velocity planning results, and attitude angular position planning results within this time period; this time period is divided into three sub-time periods, in chronological order: the constant acceleration motion period. Uniform motion for a period of time Heheng accelerated movement period , , and It was also planned using the trapezoidal attitude maneuver path planning method; The attitude angular acceleration planning result is as follows: ， in For the angular acceleration planning parameters of the design, To control the three time periods, in At time t, the angular acceleration is planned as ,exist At time t, the angular acceleration is planned to be 0. At time t, the angular acceleration is planned as ; The attitude angular velocity planning result is as follows: ； The attitude angle position planning result is as follows: ； In the attitude angular velocity planning results and attitude angular position planning results, The initial attitude angular velocity is the initial attitude angular velocity at the start of the satellite's attitude maneuver. To determine the initial attitude angle position at the start of the planning, i.e., the attitude angle position at the initial moment of the satellite's attitude maneuver, when time is... The attitude angle planning result is the angular position of the satellite's final desired attitude. , The maximum value of attitude angular acceleration during the time interval from the initial attitude adjustment of the satellite to the final desired attitude. Let be the minimum attitude angular velocity during the time interval taken from the initial attitude adjustment of the satellite to the final desired attitude, where , The settings are based on the satellite's attitude maneuverability.

2. The orbital satellite attitude control method according to claim 1, characterized in that, The calculation of the first control torque is based on PID control theory.

3. The orbital satellite attitude control method according to claim 1 or 2, characterized in that, The satellite in question is a low-Earth orbit communication satellite.

4. The orbital satellite attitude control method according to claim 2, characterized in that, The control method also includes determining the layout of the electric propulsion thrusters in the satellite.

5. The orbital satellite attitude control system used in the method according to any one of claims 1-4, characterized in that, include: The system comprises an attitude maneuver path planning module, a control torque direction adjustment module, an attitude maneuver control quantity calculation module, an attitude maneuver control switching module, and a control module. The attitude maneuver path planning module performs path planning and then transmits the path planning result to the attitude maneuver control switching module to select a control method. The attitude maneuver control quantity calculation module then calculates the desired control torque and, in conjunction with the torque gyroscope configuration, the control torque direction adjustment module obtains the control signal from the control torque gyroscope. Finally, the control module drives the control actuator to generate control torque based on the control signal.

6. The orbital satellite attitude control system according to claim 5, characterized in that, The attitude maneuver control quantity calculation module includes a first attitude maneuver control quantity calculation module and a second attitude maneuver control quantity calculation module.

7. The orbital satellite attitude control system according to claim 6, characterized in that, The first attitude maneuver control quantity calculation module is used to calculate the first desired torque based on PID control theory.

8. The orbital satellite attitude control system according to claim 6, characterized in that, The second attitude maneuver control quantity calculation module is used to calculate the second desired torque.

9. A computer device, characterized in that, The computer device includes a processor coupled to a memory storing at least one computer program, which is loaded and executed by the processor to enable the computer device to implement the orbital satellite attitude control method as described in any one of claims 1 to 4.

10. A computer storage medium, characterized in that, The computer storage medium stores at least one computer program, which is loaded and executed by a processor to enable the computer to implement the orbital satellite attitude control method according to any one of claims 1 to 4.

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