An attitude control method for a spacecraft with asymmetric structure

Abstract

This invention is an attitude control method for asymmetric spacecraft, comprising three steps: First, quantifying the disturbance torque of the asymmetric structure and establishing a multi-rigid-body dynamics model including a pyramid-configured control moment gyroscope group and the asymmetric spacecraft; second, designing the control law of the control moment gyroscope group, adjusting the rotation angle and angular velocity of the four frames to output control torque while avoiding singularities, for controlling the attitude of the asymmetric spacecraft; finally, designing the control law for the spacecraft base attitude, updating the inertial torque in real time by feeding back the current attitude error, and simultaneously providing nonlinear torque and compensation torque. The three torques are combined and provided as the desired control torque to the control law of the control moment gyroscope group.

Description

Technical Field

[0001] This invention relates to the field of spacecraft attitude control, specifically a method for attitude control of a spacecraft with an asymmetric structure based on a control moment gyroscope. Background Technology

[0002] With the rapid development of the aerospace industry, modern spacecraft are facing increasingly complex and sophisticated missions, placing higher demands on their attitude control and maneuverability. Traditional actuators, such as reaction wheels, jet thrusters, and magnetic torque converters, each have performance limitations and cannot meet the requirements for high maneuverability and rapid, stable attitude control. Control moment gyroscopes, as an effective angular momentum exchange mechanism, offer advantages such as large output torque, fast response, and zero pollution, and are considered the ideal actuator for spacecraft attitude control.

[0003] For spacecraft performing on-orbit docking missions or those with multi-modal features, the emergence of new composite structures after docking or differences in the structure of functional modules can lead to structural asymmetry and inconsistent component dynamic parameters. Currently, existing technologies have many shortcomings in addressing the attitude control requirements of asymmetric spacecraft. Due to the complex and variable structures of asymmetric spacecraft, existing control methods can only meet the attitude control needs of structurally symmetrical spacecraft to a certain extent and are difficult to apply to the attitude control of spacecraft with asymmetric structural characteristics. Therefore, designing an effective attitude control method specifically for the characteristics of asymmetric spacecraft has become an urgent problem to be solved. Summary of the Invention

[0004] To address the aforementioned problems, this invention proposes an attitude control method for asymmetric spacecraft based on a control moment gyroscope.

[0005] The technical solution of the present invention is as follows:

[0006] An attitude control method for an asymmetric spacecraft specifically includes the following steps:

[0007] First, the disturbance torque of the asymmetric structure is quantified, and a multi-rigid-body dynamics model is established, including the control torque gyroscope group of the pyramid configuration and the asymmetric structure spacecraft.

[0008] Secondly, the control law of the control moment gyroscope group is designed, and the rotation angle and angular velocity of the four frames are adjusted so that the control moment gyroscope group can output control torque while avoiding singularities, which is used to control the attitude of asymmetric spacecraft.

[0009] Finally, the control law for the attitude of the spacecraft base is designed. The inertial torque is updated in real time by feeding back the current attitude error, while nonlinear torque and compensation torque are provided. The three torques are combined and used as the desired control torque to provide the control torque gyroscope group with the manipulation law.

[0010] The beneficial effects of this invention are:

[0011] This invention presents an attitude control method for asymmetric spacecraft, using a control moment gyroscope group as the actuator. It fully considers the asymmetric structural characteristics and mission requirements of the spacecraft, achieving high-precision attitude control. This method is particularly suitable for the development of future modular and multifunctional spacecraft, providing new ideas and solutions for the design and control of asymmetric spacecraft. Attached Figure Description

[0012] Figure 1 This is a schematic diagram of the overall appearance structure of an asymmetric structure spacecraft according to a specific embodiment of the present invention;

[0013] Figure 2 This is a flowchart of the control system according to a specific embodiment of the present invention;

[0014] Figure 3 This is the attitude angle trajectory obtained by the present invention;

[0015] Figure 4 This refers to the angular velocity error.

[0016] Figure 5 For attitude quaternion error;

[0017] Figure 6 To control the output torque of the torque gyroscope group.

[0018] The labels in the attached diagram are explained as follows:

[0019] Left auxiliary spacecraft (101), spacecraft base (102), right auxiliary spacecraft (103), left joint (104), right joint (105). Detailed Implementation Plan

[0020] The present invention will be further described below with reference to the accompanying drawings and specific embodiments. The illustrative embodiments and descriptions herein are used to explain the present invention, but are not intended to limit the present invention.

[0021] This invention includes a dynamic model of an asymmetric spacecraft, a control law for a control moment gyroscope group, and a control law for the attitude of the spacecraft base.

[0022] The asymmetric spacecraft model described is a three-rigid-body structure, consisting of a spacecraft base, a left sub-spacecraft, and a right sub-spacecraft. The geometric center of the spacecraft base is its center of mass. The left sub-spacecraft is connected to the left side of the spacecraft base via a left joint, and the right sub-spacecraft is connected to the right side of the spacecraft base via a right joint. During attitude adjustments, the spacecraft base is affected by the torques of the left and right sub-spacecraft. Due to the different mass distributions of the left and right sub-spacecraft, there is a significant asymmetry in the moments of inertia, which disturbs the attitude control of the spacecraft base. A control torque gyroscope group is installed at the geometric center of the spacecraft base to provide control torque for its attitude control.

[0023] The rotational motion of the spacecraft base is controlled by the control torque M generated by the control torque gyroscope group. ctrl and the reaction torque M from the left and right auxiliary spacecraft. 12 and M 13 It has been decided to establish an attitude model for the spacecraft's base:

[0024]

[0025] In the formula, I1 is the moment of inertia of the spacecraft base, and ω1 is the angular velocity of the spacecraft base.

[0026] The attitude adjustment of an asymmetric spacecraft is a rotational system, and the time derivative of its attitude quaternion follows the formula:

[0027]

[0028] In the formula, q1 is the attitude quaternion of the spacecraft base. This is quaternion multiplication.

[0029] The attitude dynamic equations for the left and right sub-spacecraft are as follows:

[0030]

[0031]

[0032] In the formula, I2 is the moment of inertia of the left sub-spacecraft, ω2 is the angular velocity of the left sub-spacecraft, and M... e2 It is the joint driving torque of the left auxiliary spacecraft.

[0033] In the formula, I3 is the moment of inertia of the right auxiliary spacecraft, ω3 is the angular velocity of the right auxiliary spacecraft, and M... e3 It is the joint driving torque of the right auxiliary spacecraft.

[0034] According to the appendix Figure 1The coordinate system specifies that the rotation axes of the left and right joints are parallel to the y-axis of the spacecraft base, and the left and right joints respectively constrain the left and right auxiliary spacecraft to rotate only around the y-axis.

[0035] Under the action of the driving torque of the left joint and the driving torque of the right joint, both the left and right auxiliary spacecraft rotate about the y-axis relative to the spacecraft base.

[0036] For the left auxiliary spacecraft, the left joint reaction torque M acting on the spacecraft base is... 12 for:

[0037] M 12 =-M e2

[0038] For the right auxiliary spacecraft, the right joint reaction torque M acting on the spacecraft base is... 13 for:

[0039] M 13 =-M e3

[0040] According to the appendix Figure 1 The coordinate system specifies that a PD controller is used to control the left and right sub-spacecraft to reach the desired angle.

[0041] For the left sub-spacecraft:

[0042] For the right auxiliary spacecraft:

[0043] K p2 and K d2 These are the proportional and differential gains of the left auxiliary spacecraft controller, respectively.

[0044] K p3 and K d3 These are the proportional and differential gains of the right auxiliary spacecraft controller, respectively.

[0045] θ 2d and θ 3d These are the desired angles for the left and right auxiliary spacecraft, respectively.

[0046] and These are the expected angular velocities of the left and right auxiliary spacecraft, respectively.

[0047] θ2 and θ3 are the actual rotation angles of the left and right auxiliary spacecraft relative to the spacecraft base, respectively.

[0048] The moments of the left and right auxiliary spacecraft on the spacecraft base can be obtained as follows:

[0049]

[0050]

[0051] The disturbance torques exerted by the left and right sub-spacecraft on the spacecraft base are: T d =M 12 +M 13

[0052]

[0053] The dynamics of the left and right auxiliary spacecraft follow the basic equations of rigid body rotation. The angular accelerations of the left and right auxiliary spacecraft are:

[0054]

[0055]

[0056] The moments of inertia I2 of the left sub-spacecraft and I3 of the right sub-spacecraft indirectly affect the dynamic changes of θ2 and θ3 through their angular accelerations, thus affecting M. 12 and M 13 The value of .

[0057] When I2≠I3, i.e., when the inertia of the left and right sub-spacecraft are different, the torque contribution M exerted by the left and right sub-spacecraft on the spacecraft base is... 12 and M 13 Under the same control input, they exhibit different dynamic characteristics, which means that the torques applied by the left and right auxiliary spacecraft to the spacecraft base cannot cancel each other out, resulting in a disturbance torque on the spacecraft base and thus affecting the attitude control of the spacecraft base.

[0058] The attitude control input torque of an asymmetric spacecraft is the control torque M provided by the control torque gyroscope group. ctrl The state output is the angular velocity ω1 and angular acceleration of the spacecraft base. And the attitude quaternion q1.

[0059] The control torque gyroscope group's manipulation law model is a pseudo-inverse manipulation law, which can distribute the target control torque to multiple control torque gyroscopes.

[0060] The pseudo-inverse control of the control moment gyroscope group requires solving the following matrix equations:

[0061]

[0062] In the formula, T0 is the desired control torque vector, and J is the inertia tensor matrix of the control torque gyroscope group, which is related to the frame angle. To control the rate of change of angular momentum of the torque gyroscope group.

[0063] Solve using the pseudo-inverse method.

[0064]

[0065] In the formula, W is used to adjust the priority or sensitivity in different directions, which is dynamically adjusted by optimizing the parameters.

[0066] Angular momentum is obtained by integrating the rate of change of angular momentum.

[0067]

[0068] In the formula, h0 is the known initial value of angular momentum.

[0069] Further integration of the angular momentum yields the control moment gyroscope frame angle θ.

[0070] The actual output control torque of the control torque gyroscope group is T, and the dynamic model of the control torque gyroscope group is:

[0071]

[0072] Calculate the attitude quaternion error q of the spacecraft base. e and angular velocity error ω e The calculation formula is as follows:

[0073]

[0074]

[0075] ω e =ω s -ω g

[0076] In the formula, q s Let be the desired attitude quaternion of the spacecraft base. For q s The conjugate quaternion, q g The actual attitude quaternion of the spacecraft base. For q g The conjugate quaternion, ω s Let ω be the desired angular velocity of the spacecraft base. g This represents the actual angular velocity of the spacecraft base.

[0077] The control law for the spacecraft base attitude aims to generate the desired control torque for the control torque gyroscope group manipulation law. The main input to this control law is the attitude quaternion error q. e Angular velocity error ω e Actual angular velocity ω gand disturbance torque T d Attitude quaternion error q e This represents the attitude error between the current attitude and the target attitude, providing information on the deviation of the attitude angles. Angular velocity error ω e It represents the difference between the actual angular velocity and the desired angular velocity, providing information on the deviation of the attitude angular velocity.

[0078] The control law generation process includes error limiting processing, proportional-derivative controller generating inertial torque, disturbance torque generating compensation torque, angular momentum compensation generating nonlinear torque, and signal superposition.

[0079] Error limiting processing of the input quaternion error q e and angular velocity error ω e Amplitude limiting is implemented to avoid instability or nonlinear effects caused by excessive errors.

[0080] Proportional-derivative control uses proportional gain K p and differential gain K d Amplified attitude quaternion errors after amplitude limiting and attitude angular velocity error Generating inertial torque T pd The formula is:

[0081]

[0082] Based on the obtained disturbance torque, a compensation torque is added to the controller to counteract its influence, thereby reducing the attitude error of the spacecraft base. The compensation torque T b The formula is:

[0083]

[0084] To compensate for the nonlinear effect of angular momentum caused by rapid attitude changes, a vector cross product is calculated using the actual angular velocity and angular momentum to generate a nonlinear torque T. c :

[0085] T c =H×ω g

[0086] In the formula, H represents the angular momentum of the spacecraft base. Finally, the inertial torque, compensation torque, and nonlinear torque are superimposed to generate the desired control torque T0 of the control law:

[0087] T0 = ​​T pd +T b +T c

Claims

1. A method of attitude control of a spacecraft of asymmetric structure, characterized in that: First, the disturbance torque of the asymmetric structure is quantified, and a multi-rigid-body dynamics model is established, including a pyramid-configured control moment gyroscope group and an asymmetric structure spacecraft. Second, the control law of the control moment gyroscope group is designed, adjusting the rotation angle and angular velocity of the four frames to output control torque while avoiding singularities, which is used to control the attitude of the asymmetric structure spacecraft. Finally, the control law of the spacecraft base attitude is designed, updating the inertial torque in real time by feeding back the current attitude error, while providing nonlinear torque and compensation torque. The three torques are combined and used as the desired control torque to provide to the control law of the control moment gyroscope group.

2. The attitude control method of an asymmetrically structured spacecraft according to claim 1, characterized by: The asymmetric spacecraft includes a spacecraft base, a left sub-spacecraft, and a right sub-spacecraft. The left sub-spacecraft is connected to the left side of the spacecraft base via a left joint, and the right sub-spacecraft is connected to the right side of the spacecraft base via a right joint. The left and right sub-spacecrafts have asymmetrical structures and inconsistent dynamic parameters.

3. The asymmetric spacecraft attitude control method of claim 1, wherein: The inertial torque is generated by a proportional-differential controller, the compensation torque is generated by a disturbance torque, and the nonlinear torque is generated by angular momentum compensation.

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