Asymmetric configuration satellite attitude maneuver process coupling compensation method

Abstract

The application discloses a kind of asymmetric configuration satellite attitude maneuvering process coupling compensation method, comprising: establishing the satellite attitude dynamics model including the non-diagonal item of moment of inertia matrix;Read current attitude sensor data, and calculate deviation quaternion and deviation angular velocity;According to the attitude dynamics model and current angular velocity data, calculate coupling compensation feedforward control torque;Based on the coupling compensation feedforward control torque, deviation quaternion and deviation angular velocity, the control torque after compensation is calculated;The control torque after compensation is limited and priority torque distribution, and the final control torque is obtained;Final control torque is sent to actuator drive satellite and carries out attitude maneuvering.Solve the problem that coupling effect caused by inertia product cannot be effectively compensated when asymmetric configuration satellite is in large-angle attitude maneuvering in prior art, and then affect control precision and stability.

Description

Technical Field

[0001] This invention relates to the technical field of spacecraft attitude control, and in particular to a coupling compensation method for the attitude maneuvering process of an asymmetric configuration satellite. Background Technology

[0002] During their on-orbit operation, satellites need to perform various attitude maneuvers, with typical applications including Earth-staring imaging and Earth-scanning imaging. These attitude maneuvers typically place dual technical demands on satellites: on the one hand, they need to achieve large-angle attitude rotations within a short time; on the other hand, they need to maintain high attitude stability and pointing accuracy simultaneously. However, during attitude maneuvers, the satellite's angular velocity and angular acceleration induce gyroscopic coupling torques. These coupling torques interact between various attitude axes, directly affecting the control accuracy and stability during attitude maneuvers, becoming a key technical bottleneck restricting mission performance.

[0003] In existing technologies, a feedback linearization control scheme based on a rigid body dynamics model has been proposed to address the problem of large-angle attitude maneuver control for satellites. This scheme involves designing a large-angle attitude maneuver controller and utilizing a quaternion feedback mechanism to achieve attitude tracking control. This control method achieves good control performance on symmetrically configured satellites with accurate nominal inertia parameters. However, in applications involving asymmetrically configured satellites, the control model neglects the off-diagonal term of the inertia tensor (i.e., the inertia product), leading to ineffective compensation of gyro coupling torque during large-angle attitude maneuvers. This results in significantly increased attitude overshoot and prolonged settling time. Practical experience shows that this method can only meet the preset control accuracy requirements when the satellite's inertia distribution is approximately symmetrical and the maneuver angle is small.

[0004] Another existing technology is the attitude decoupling control scheme based on sliding mode variable structure control theory: an attitude decoupling controller is constructed by introducing a switching function to achieve decoupling control of the satellite's three-axis attitude channels. This method has a certain robustness to model uncertainties, but it has significant technical drawbacks: firstly, the selection of the switching gain relies on engineering experience and lacks a systematic design basis; secondly, no feedforward compensation mechanism is designed for the inertial coupling terms generated by asymmetric satellite configurations. Therefore, on satellites with large inertial products, inter-axis coupling oscillations still occur when performing high-speed attitude maneuvers, and the high-frequency chattering problem of the control torque is prominent. This method can only maintain stable control performance in scenarios where the satellite is maneuvering at a small angle relative to the inertial frame.

[0005] In summary, existing attitude control methods generally suffer from insufficient coupling compensation, inadequate robustness, and significant high-frequency jitter when dealing with large-angle attitude maneuvers of asymmetric satellites, making it difficult to simultaneously meet the dual requirements of high dynamic response speed and high-precision control. Therefore, there is an urgent need to propose an attitude control method that can achieve effective coupling compensation tailored to the structural characteristics of asymmetric satellites, in order to solve the aforementioned problems of existing technologies. Summary of the Invention

[0006] The present invention aims to solve the problem in the prior art that the coupling effect caused by the inertial product cannot be effectively compensated when asymmetric configuration satellites perform large-angle attitude maneuvers, thus affecting control accuracy and stability.

[0007] This invention provides a coupling compensation method for attitude maneuvering processes of asymmetric configuration satellites, comprising: Establish a satellite attitude dynamics model that includes off-diagonal terms of the rotational inertia matrix; Read the current attitude sensor data and calculate the deviation quaternion and deviation angular velocity; Based on the attitude dynamics model and the current angular acceleration data, calculate the coupling compensation feedforward control torque; The compensated control torque is calculated based on the coupled compensation feedforward control torque, the deviation quaternion, and the deviation angular velocity. The compensated control torque is limited and prioritized for torque allocation to obtain the final control torque; The final control torque is sent to the actuators to drive the satellite to perform attitude maneuvers.

[0008] In one embodiment of the present invention, the attitude dynamics model includes: ; in The output torque of the flywheel assembly. The angular velocity of the satellite, For the satellite's angular acceleration, Let ω be the angular momentum of the flywheel assembly. Here is the rotational inertia matrix; ; , , Let xyz be the moment of inertia of the principal axes. , , It is the inertial product.

[0009] In one embodiment of the present invention, the attitude dynamics model includes: ; in , , For the three-axis output torque of the flywheel assembly, , , The three-axis angular velocities of the satellite, , , The satellite's three-axis angular acceleration, , , The angular momentum of the flywheel assembly is angular momentum along its three axes.

[0010] In one embodiment of the present invention, calculating the coupling compensation feedforward control torque based on the attitude dynamics model and the current angular velocity data includes: Angular acceleration is calculated using a differential tracker; Calculate the coupling compensation feedforward control torque based on the moment of inertia matrix and angular acceleration.

[0011] In one embodiment of the present invention, the transfer function of the differential tracker includes: ; Where s is the complex frequency and T is the filtering time constant.

[0012] In one embodiment of the present invention, the coupling compensation feedforward control torque is: ; in , , For coupling compensation of feedforward control torque The three-axis components; It is a matrix composed of satellite inertial product parameters.

[0013] In one embodiment of the present invention, the compensated control torque is: ; in It is a biased quaternion. The deviation angular velocity, This is the proportional control coefficient. The differential control coefficient, This refers to the satellite's angular momentum.

[0014] In one embodiment of the present invention, the compensated control torque is limited by the following formula: ; in To the maximum output torque of the actuator, The three-axis components of the control torque after compensation.

[0015] In one embodiment of the present invention, the priority torque allocation strategy is as follows: ; in, This is due to the unrealized coupling compensation torque residual caused by saturation. To compensate for the weighting coefficients, they are dynamically adjusted according to the current maneuver phase.

[0016] The present invention has the following beneficial effects: (1) By explicitly modeling the off-diagonal terms of the moment of inertia matrix, feedforward compensation is directly performed for the coupling effect caused by the asymmetric configuration, which significantly reduces the inter-axis coupling oscillation. (2) Combining robust control strategies improves the adaptability to model parameter uncertainties and external disturbances, and avoids high-frequency chattering; (3) It is suitable for large-angle and fast maneuvering scenarios, and can ensure the attitude tracking accuracy and system stability under complex configurations. Attached Figure Description

[0017] Figure 1 A satellite attitude control method according to an embodiment of the present invention is shown; Figure 2 The diagram shows the attitude maneuver angular velocity without coupling compensation in one embodiment of the present invention; Figure 3 A partial view of the attitude maneuver angular velocity without coupling compensation is shown in one embodiment of the present invention; Figure 4 The diagram shows the attitude maneuver angular velocity under coupled compensation in one embodiment of the present invention; Figure 5 A partial diagram of the attitude maneuver angular velocity under coupled compensation in one embodiment of the present invention is shown. Detailed Implementation

[0018] In the following description, the invention is described with reference to various embodiments. However, those skilled in the art will recognize that the embodiments may be practiced without one or more specific details or with other alternatives and / or additional methods, materials, or components. In other instances, well-known structures, materials, or operations are not shown or described in detail so as not to obscure the inventive points of the invention. Similarly, for illustrative purposes, specific quantities, materials, and configurations are set forth to provide a comprehensive understanding of embodiments of the invention. However, the invention is not limited to these specific details.

[0019] In this invention, the various embodiments are merely intended to illustrate the solutions of the invention and should not be construed as limiting.

[0020] In this specification, references to "an embodiment" or "this embodiment" mean that a particular feature, structure, or characteristic described in connection with that embodiment is included in at least one embodiment of the invention. The phrase "in one embodiment" appearing throughout this specification does not necessarily refer to the same embodiment in all instances.

[0021] Furthermore, the numbering of the steps in the methods of the present invention does not limit the execution order of the method steps. Unless otherwise specified, the method steps may be executed in different orders.

[0022] The present invention will be further described below with reference to the accompanying drawings and specific embodiments.

[0023] Figure 1 A satellite attitude control method according to an embodiment of the present invention is shown.

[0024] like Figure 1 As shown, the satellite attitude control method in this embodiment includes: S1. Set the desired satellite attitude quaternion and angular velocity.

[0025] For asymmetric satellite configurations, the rotational inertia matrix has a significant off-diagonal term (product of inertia), expressed as: ; , , Let xyz be the moment of inertia of the principal axes. , , It is the inertial product.

[0026] The satellite's attitude dynamics equations are as follows: ; in The output torque of the flywheel assembly. The angular velocity of the satellite, For the satellite's angular acceleration, Let ω be the angular momentum of the flywheel assembly. Here is the rotational inertia matrix.

[0027] Written in triaxial form, as follows: ; in , , For the three-axis output torque of the flywheel assembly, , , The three-axis angular velocities of the satellite, , , The satellite's three-axis angular acceleration, , , The angular momentum of the flywheel assembly is angular momentum along its three axes.

[0028] For satellite attitude control systems, the desired attitude trajectory and desired angular velocity trajectory are first obtained based on the attitude maneuver mission planning. If only traditional feedback control is used without compensation for inertial coupling terms, the control accuracy will significantly decrease due to the asymmetric configuration.

[0029] S2. Obtain the current attitude quaternion and angular velocity.

[0030] Read the current attitude sensor data to obtain the current attitude quaternion and angular velocity.

[0031] S3. Calculate the deviation quaternion and angular velocity.

[0032] Calculate the deviation quaternion based on the desired attitude trajectory and desired angular velocity trajectory, as well as the current attitude quaternion and angular velocity. Deviation angular velocity .

[0033] S4, Differential tracker calculates angular acceleration.

[0034] Since satellite attitude sensors (such as gyroscopes and star sensors) can typically only directly output attitude quaternions and angular velocities, they cannot directly measure angular acceleration. Directly differentiating the angular velocity signal (e.g., using backward differential) would significantly amplify the high-frequency components of the measurement noise, causing the calculated angular acceleration to oscillate violently, which could lead to the failure of the coupling compensation torque or even trigger high-frequency chattering in the actuator. This invention introduces a differential tracker with filtering characteristics to obtain a smooth and real-time estimate of angular acceleration. The transfer function is shown below: ; Where s is the complex frequency and T is the filtering time constant.

[0035] S5. Calculate the coupling compensation torque.

[0036] If only traditional feedback control is used without compensating for inertial coupling terms, the control accuracy will decrease significantly due to the asymmetric configuration.

[0037] This invention proposes a coupled compensation feedforward control law to correct the desired control torque. Desired control torque Designed as follows: ; in , , For coupling compensation of feedforward control torque The three-axis components; It is a matrix composed of satellite inertial product parameters.

[0038] S6, the PD controller calculates the feedback control torque.

[0039] By combining the feedback torque of the satellite PD control with the coupled compensation feedforward control torque, the compensated control torque is obtained as follows: ; in It is a biased quaternion. The deviation angular velocity, This is the proportional control coefficient. The differential control coefficient, This refers to the satellite's angular momentum.

[0040] S7, Output Torque Limiting and Priority Assignment.

[0041] The satellite actuator has an output torque saturation limit. Let the maximum output torque of the actuator be... Unlike the three-axis joint limiting during a fixed-path maneuver, independent limiting of the three axes is required to fully compensate for the coupling torque.

[0042] Therefore, when the calculated control torque exceeds the saturation limit, torque distribution optimization is required, as follows: ; in To the maximum output torque of the actuator, The three-axis components of the control torque after compensation.

[0043] Because the coupling strength of each axis differs under asymmetric configurations, using proportional limiting may lead to coupling compensation failure. This invention proposes a priority torque allocation strategy: ; in, This is due to the unrealized coupling compensation torque residual caused by saturation. To compensate for the weighting coefficients, they are dynamically adjusted according to the current maneuver phase.

[0044] As can be seen from the above equation, prioritizing the implementation of the coupling compensation term under saturation conditions can effectively suppress attitude oscillations caused by actuator saturation.

[0045] Determine whether the attitude maneuver is complete. If not, return to step S2 to continue closed-loop control; if complete, switch to attitude hold mode.

[0046] Those skilled in the art should understand that the above embodiments are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Without departing from the core principles of the present invention, the following improvements and modifications can be made to the technical solution, all of which fall within the scope of protection of the present invention: (1) Variation of the method for obtaining angular acceleration: In this embodiment, the s / (1+Ts) transfer function is used for filtering differentiation. In practical applications, it can be replaced with other differentiation algorithms with noise suppression functions, such as a tracking differentiator (TD), a Kalman filter, or a higher-order sliding mode differentiator, while retaining the same core idea.

[0047] (2) Optimization of the control law architecture: The basic "feedforward compensation + feedback control" architecture can be further combined with sliding mode control to enhance anti-interference capability, or model predictive control (MPC) can be used to explicitly handle actuator saturation constraints. Simultaneously, a gain scheduling mechanism can be introduced to dynamically adjust control parameters based on the maneuver amplitude.

[0048] (3) Transformation of mathematical description form: Attitude description can be achieved using quaternions, modified Rodrigues parameters (MRP), and Euler angles; the coordinate system can be either an orbital frame or an inertial frame, and equivalent coupling compensation can be achieved through coordinate transformation, while the core idea remains unchanged.

[0049] In another embodiment of the present invention, the moment of inertia matrix of a satellite is shown in the following table: Table 1 Satellite rotational inertia matrix The satellite performs attitude maneuvers in a three-axis Earth-to-ground stable attitude. The parameters of the attitude maneuvers are as follows: Table 2 Satellite initial attitude and target attitude Without coupling compensation, the angular velocity of the satellite during attitude maneuvers is as follows: Figure 2 and Figure 3 As shown, the maximum angular velocity in the Z direction generated by the coupling during the satellite's attitude maneuver is 0.11° / s.

[0050] like Figure 4 and Figure 5 As shown, after coupling compensation torque, the angular velocity of the satellite's attitude maneuver is as follows: the maximum angular velocity in the Z direction generated by coupling during the satellite's attitude maneuver is 0.0317° / s.

[0051] Therefore, through coupling compensation, the coupling angular velocity was reduced from 0.11° / s to 0.0317° / s, a reduction of 71.18%.

[0052] Although various embodiments of the invention have been described above, it should be understood that they are presented by way of example only and not as limitations. It will be apparent to those skilled in the art that various combinations, modifications, and alterations can be made without departing from the spirit and scope of the invention. Therefore, the breadth and scope of the invention disclosed herein should not be limited by the exemplary embodiments disclosed above, but should be defined solely by the appended claims and their equivalents.

Claims

1. A coupling compensation method for attitude maneuvering processes of an asymmetric configuration satellite, characterized in that, include: Establish a satellite attitude dynamics model that includes off-diagonal terms of rotational inertia; Read the current attitude sensor data and calculate the deviation quaternion and deviation angular velocity; Based on the attitude dynamics model and the current angular acceleration data, calculate the coupling compensation feedforward control torque; The compensated control torque is calculated based on the coupled compensation feedforward control torque, the deviation quaternion, and the deviation angular velocity. The compensated control torque is limited and prioritized for torque allocation to obtain the final control torque; The final control torque is sent to the actuators to drive the satellite to perform attitude maneuvers.

2. The method according to claim 1, characterized in that, The attitude dynamics model includes: ； in The output torque of the flywheel assembly. The angular velocity of the satellite, For the satellite's angular acceleration, Let ω be the angular momentum of the flywheel assembly. Here is the rotational inertia matrix; ； , , Let xyz be the moment of inertia of the principal axes. , , It is the inertial product.

3. The method according to claim 2, characterized in that, The attitude dynamics model includes: ； in , , For the three-axis output torque of the flywheel assembly, , , The three-axis angular velocities of the satellite, , , The satellite's three-axis angular acceleration, , , The angular momentum of the flywheel assembly is angular momentum along its three axes.

4. The method according to claim 1, characterized in that, The calculation of the coupling compensation feedforward control torque based on the attitude dynamics model and the current angular velocity data includes: Angular acceleration is calculated using a differential tracker; The coupling compensation feedforward control torque is calculated based on the satellite's rotational inertia matrix and angular acceleration.

5. The method according to claim 4, characterized in that, The differential tracker transfer function includes: ； Where s is the complex frequency and T is the filtering time constant.

6. The method according to claim 4, characterized in that, The coupling compensation feedforward control torque is: ； in , , For coupling compensation of feedforward control torque The three-axis components; It is a matrix composed of satellite inertial product parameters.

7. The method according to claim 1, characterized in that, The compensated control torque is: ； in It is a biased quaternion. The deviation angular velocity, This is the proportional control coefficient. The differential control coefficient, This refers to the satellite's angular momentum.

8. The method according to claim 1, characterized in that, The compensated control torque is limited using the following formula: ； in To the maximum output torque of the actuator, The three-axis components of the control torque after compensation.

9. The method according to claim 1, characterized in that, The priority torque allocation strategy is as follows: ； in, This is due to the unrealized coupling compensation torque residual caused by saturation. To compensate for the weighting coefficients, they are dynamically adjusted according to the current maneuver phase.

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