A satellite attitude control method for suppressing flexible vibration

Abstract

The present invention provides a satellite flexible vibration suppression attitude control method. By adopting the idea of cascading input-to-state stability design, an attitude angle closed-loop control and an angular velocity tracking control form a cascaded system. The attitude controller is designed in a hierarchical manner. A second-order command filter is used to assign saturation constraints for tracking the virtual angular velocity and the virtual control quantity. The control method has the characteristics of small computational load and easy implementation in engineering. The flexible vibration suppression attitude controller realizes the smoothing processing of the angular velocity tracking command and the actuator control command by configuring the constraint amplitude, damping coefficient, and natural oscillation frequency of the command filter, so that the maximum angular velocity and angular acceleration of the satellite are constrained, and the angular acceleration of the excitation source of the satellite flexible attachment is limited. During the large-angle attitude maneuver, the satellite is constrained by the second-order saturation command filter, and the angular acceleration becomes zero after reaching the maximum angular velocity, and the vibration of the flexible attachment is suppressed.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft attitude control, and particularly to a satellite flexible vibration suppression attitude control method. Background Art

[0002] With the development of space application payloads and platform technologies, the requirements of satellite missions are constantly increasing. It is necessary to perform high-precision attitude control on the combination of rigid bodies and flexible appendages. Compared with rigid-body satellites, flexible-attached satellites have problems such as low test accuracy of satellite mass characteristic parameters, great difficulty in experimental operations, and easy oscillation during attitude maneuvering. Affected by external environmental disturbance torques, they are not only likely to excite flexible vibrations, but also seriously affect the pointing accuracy and stability after attitude maneuvering.

[0003] In terms of improving satellite attitude control accuracy, currently, mainstream algorithms such as on-orbit identification of satellite mass characteristic parameters, attitude maneuver path planning control, attitude maneuver adaptive robust control, sliding mode variable structure control, and active vibration control using piezoelectric materials are adopted, and fruitful research results have been achieved. However, further improvement is still needed in terms of high-precision attitude control, fast attitude maneuver control, and ease of engineering implementation.

[0004] In view of the problems existing in the prior art, the present invention designs a satellite flexible vibration suppression attitude control method, which can effectively balance satellite fast attitude maneuver control, high-precision attitude control, and is easy to be realized in satellite attitude control engineering. Summary of the Invention

[0005] In order to overcome the disadvantages and deficiencies existing in the prior art, the present invention provides a satellite flexible vibration suppression attitude control method, which takes into account satellite on-orbit fast attitude maneuver control, flexible vibration suppression, and high-precision attitude control, and realizes precise attitude maneuver control of on-orbit flexible-attached satellites.

[0006] The present invention provides a satellite flexible vibration suppression attitude control method, which is applied to on-orbit flexible satellites and includes the following steps:

[0007] Step 1: Establish a flexible satellite attitude dynamics model, and construct a flexible satellite attitude dynamics equation according to the flexible satellite attitude dynamics model:

[0008]

[0009] Wherein, represents the identity matrix, × represents the cross product, represents the attitude quaternion measured by the satellite, represents the angular velocity measured by the satellite, represents the moment of inertia calibrated on the ground of the satellite, represents The momentum turns to the installation matrix of the satellite system, represents the speed of the momentum wheel, represents the flexible mode coordinate vector of the satellite system, and denote the flexible modal damping and stiffness coefficient matrices respectively, and denote the modal damping coefficient and modal frequency respectively, represents the coupling matrix between the flexible mode and the rigid body, represents the spatial interference torque, where and The definition is as follows:

[0010]

[0011] Step 2: Define and calculate the attitude tracking error e1 and angular velocity error e2 of the flexible satellite based on the attitude dynamics equation of the flexible satellite:

[0012]

[0013] in, is the quaternion attitude error vector part, and the quaternion attitude error is defined as represents the desired attitude quaternion of the attitude control system during the control process, Represents the quaternion cross multiplication calculation, It represents the angular velocity command output by the second-order saturation command filter of the angular velocity controller. The second-order saturation command filter of the angular velocity controller is defined as:

[0014]

[0015] Among them, ζ1 and ω n1 They represent the damping coefficient and natural oscillation frequency of the second-order saturation command filter of the angular velocity controller, represents the virtual angular velocity control law, sat R (·) represents the amplitude saturation constraint function, which is defined as follows:

[0016]

[0017] in, express The maximum absolute value of the vector elements, Express Maximum value constraint of vector elements;

[0018] Step 3: Use the virtual angular velocity control law of the attitude control system in turn Process the attitude tracking error e1 and output the angular velocity command and the virtual actuator control law using the attitude control system Process the angular velocity error e2 and output the actuator control command u, the virtual angular velocity control law and the virtual actuator control law are respectively defined as follows:

[0019]

[0020] where and are the proportional coefficient and differential coefficient of the designed attitude controller, and are the desired angular velocity and the differential of the desired angular velocity, is the intermediate state quantity of the second-order saturation command filter for the angular velocity command, and are respectively and the inverse matrices of, represents the transformation matrix from the satellite body coordinate system to the desired reference frame:

[0021]

[0022] u is the actuator control command output by the second-order saturation command filter for the actuator, and the second-order saturation command filter for the actuator is defined as:

[0023]

[0024] where, ζ2 and ω n2 respectively represent the damping coefficient and natural oscillation frequency of the second-order command filter for the actuator.

[0025] After adopting the above technical solution, the present invention has at least the following beneficial effects:

[0026] 1. The present invention adopts the cascaded input-to-state stability design idea, making the attitude angle closed-loop control and angular velocity tracking control form a cascaded system. The attitude controller is designed in a hierarchical manner, and the second-order command filter is used to assign saturation constraints for tracking the virtual angular velocity and virtual control quantity. The control method has the characteristics of small computational amount and easy implementation in engineering.

[0027] 2. The flexible vibration suppression attitude controller of the present invention can achieve smooth processing of the angular velocity tracking command and the actuator control command by configuring the constraint amplitude, damping coefficient and natural oscillation frequency of the command filter, so as to constrain the maximum angular velocity and angular acceleration of the satellite, limit the angular acceleration of the excitation source of the satellite's flexible appendage. During the large-angle attitude maneuver, the satellite is constrained by the second-order saturation command filter, and the angular acceleration becomes zero after reaching the maximum angular velocity, and the vibration of the flexible appendage will be suppressed. Brief Description of the Drawings

[0028] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0029] Figure 1 It is a schematic flow diagram of the attitude control method for satellite flexible vibration suppression;

[0030] Figure 2 It is a schematic diagram of the structure of a second-order saturation command filter;

[0031] Figure 3 Curves of attitude angle and angular velocity during large-angle maneuvering;

[0032] Figure 4 Curves of momentum wheel speed and solar vector during large-angle maneuvering;

[0033] Figure 5 Curves of flexible mode during large-angle maneuvering. Detailed Embodiments

[0034] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0035] At present, most of the spacecraft launched by our country carry flexible appendages. The flexible appendages are developing towards a more complex structure and volume trend, and the control requirements for such spacecraft will become more stringent. Based on this trend, the control algorithms we use should further meet the requirements of fast maneuvering and fast stabilization of flexible spacecraft.

[0036] Refer to Figure 1 and Figure 2 , this embodiment provides an attitude control method for satellite flexible vibration suppression, which is applied to an on-orbit flexible satellite to perform fast attitude maneuvering control and high-precision attitude control for the flexible satellite, including:

[0037] Step 1: Establish a flexible satellite attitude dynamics model, and construct a flexible satellite attitude dynamics equation according to the flexible satellite attitude dynamics model:

[0038]

[0039] Wherein, represents the identity matrix, and × represents the cross product, represents the attitude quaternion obtained by satellite measurement, represents the angular velocity obtained by satellite measurement, represents the moment of inertia calibrated on the satellite ground, represents the installation matrix of the number of momentum wheels to the satellite body system, represents the flexible mode coordinate vector of the satellite body system, and respectively represent the flexible mode damping and stiffness coefficient matrices, and respectively represent the modal damping coefficient and modal frequency, represents the coupling matrix between the flexible mode and the rigid body, represents the space disturbance torque, where and are defined as follows:

[0040]

[0041] Step 2: Define and calculate the attitude tracking error e1 and angular velocity error e2 of the flexible satellite based on the flexible satellite attitude dynamics equation:

[0042]

[0043] Wherein, is the vector part of the quaternion attitude error, and the quaternion attitude error is defined as represents the desired attitude quaternion during the control process of the attitude control system, represents the quaternion cross product calculation, represents the angular velocity command output by the second-order saturation command filter of the angular velocity controller. The second-order saturation command filter of the angular velocity controller is defined as:

[0044]

[0045] Wherein, ζ1 and ω n1 respectively represent the damping coefficient and natural oscillation frequency of the second-order saturation command filter of the angular velocity controller, represents the virtual angular velocity control law, and sat R (·) represents the amplitude saturation constraint function, which is defined as follows:

[0046]

[0047] Wherein, Indicate The maximum value among the absolute values of the vector elements, Indicates for The maximum value constraint of the vector elements;

[0048] Step 3. Successively use the virtual angular velocity control law of the attitude control system Process the attitude tracking error e1 and output the angular velocity command And use the virtual actuator control law of the attitude control system Process the angular velocity error e2 and output the actuator control command u. The virtual angular velocity control law And the virtual actuator control law Are respectively defined as follows:

[0049]

[0050] Wherein, And Are the designed proportional coefficient and differential coefficient of the attitude controller, And Are the desired angular velocity and the differential of the desired angular velocity, Is the intermediate state quantity of the second-order saturation command filter of the angular velocity command, And Are respectively And The inverse matrices of, Represents the transformation matrix from the satellite body coordinate system to the desired reference frame:

[0051]

[0052] u is the actuator control command output by the second-order saturation command filter of the actuator. The second-order saturation command filter of the actuator is defined as:

[0053]

[0054] Wherein, ζ2 and ω n2 Respectively represent the damping coefficient and the natural oscillation frequency of the second-order command filter of the actuator.

[0055] Combined with the above embodiments, the following specific examples are proposed. It can be understood that the following specific examples only exemplarily elaborate on the specific implementation of the above embodiments, and do not limit the technical solutions of the above embodiments.

[0056] A simulation experiment is carried out on a certain micro-nano satellite. This micro-nano satellite is sun-pointing, operating on a sun-synchronous orbit with an orbital altitude of 500 km, the descending node is 6:00 pm, and the orbital inclination is 97.62 degrees.

[0057] (1) Satellite initial attitude angle setting:

[0058] The attitude angles of the satellite in the x, y, and z axes relative to the orbital system are set to [0.03, -0.05, 0.12] degrees;

[0059] (2) Satellite initial attitude angular velocity setting:

[0060] The attitude angular velocities of the satellite in the x, y, and z axes relative to the orbital system are set to [-0.05, 0.02, 0.1] degrees / s;

[0061] (3) Satellite configuration parameters:

[0062] The nominal angular momentum of the three-axis momentum wheels of the satellite is 80 mNms, the maximum output torque is 2.5 mNm, and the constraint of the satellite attitude maneuver angular velocity is 0.4° / s.

[0063] The moment of inertia of the satellite is set as follows:

[0064]

[0065] The modal damping coefficient of the satellite's flexible appendage The flexible modal frequency of the satellite is The flexible coupling matrix is as follows:

[0066]

[0067] (4) Satellite attitude control coefficient setting:

[0068] The virtual angular velocity control law and the virtual actuator control law The control coefficients in are ζ1 = 0.707, ζ2 = 0.707, ω n1 = 20, ω n2 = 20.

[0069] Figures 3 to 5 is the simulation result of the satellite flexible vibration suppression attitude control method under fast attitude maneuver control. The simulation results show that currently the satellite can achieve fast attitude maneuver control. After the satellite reaches the maximum angular velocity, it maneuvers at the maximum angular velocity, and the rotational speed of the momentum wheel no longer increases or decreases rapidly. During the process of the satellite's uniform maneuver (the angular acceleration tends to zero), the input of the excitation signal to the flexible mode is reduced, and the flexible mode vibration of the satellite is quickly suppressed. At the same time, the satellite does not cause attitude oscillation due to excessive initial attitude expectation error, does not require attitude maneuver planning, and can ensure a low steady-state error.

[0070] Although the present invention has been disclosed above by way of embodiments, it is not intended to limit the present invention. Any person skilled in the art within the scope of the present invention may make some modifications and refinements without departing from the spirit and scope of the present invention. Therefore, the protection scope of the present invention shall be subject to that defined by the claims.

Claims

1. A satellite flexible vibration suppression attitude control method, applied to a flexible satellite, characterized in that: include: Step 1: Establish a flexible satellite attitude dynamics model and construct the flexible satellite attitude dynamics equation based on the flexible satellite attitude dynamics model: in, represents the identity matrix, × represents the cross product, Represents the attitude quaternion measured by the satellite, represents the angular velocity measured by the satellite, represents the satellite ground-calibrated moment of inertia, express The momentum turns to the installation matrix of the satellite system, represents the speed of the momentum wheel, represents the flexible mode coordinate vector of the satellite system, and denote the flexible modal damping and stiffness coefficient matrices respectively, and denote the modal damping coefficient and modal frequency respectively, represents the coupling matrix between the flexible mode and the rigid body, represents the spatial interference torque, where and The definition is as follows: Step 2: Define and calculate the attitude tracking error e1 and angular velocity error e2 of the flexible satellite based on the attitude dynamics equation of the flexible satellite: in, is the quaternion attitude error vector part, and the quaternion attitude error is defined as represents the desired attitude quaternion of the attitude control system during the control process, Represents the quaternion cross multiplication calculation, It represents the angular velocity command output by the second-order saturation command filter of the angular velocity controller. The second-order saturation command filter of the angular velocity controller is defined as: Among them, ζ1 and ω n1 They represent the damping coefficient and natural oscillation frequency of the second-order saturation command filter of the angular velocity controller, represents the virtual angular velocity control law, sat R (·) represents the amplitude saturation constraint function, which is defined as follows: in, express The maximum absolute value of the vector elements, Express Maximum value constraint of vector elements; Step 3: Use the virtual angular velocity control law of the attitude control system in turn Process the attitude tracking error e1 and output the angular velocity command and the virtual actuator control law using the attitude control system Process the angular velocity error e2 and output the actuator control instruction u, the virtual angular velocity control law and virtual actuator control law They are defined as follows: in, and are the proportional coefficient and differential coefficient of the designed attitude controller, and are the desired angular velocity and the desired angular velocity differential, is the intermediate state quantity of the second-order saturation command filter of the angular velocity instruction, and They are and The inverse matrix of Represents the transformation matrix from the satellite body coordinate system to the desired reference system: u is the actuator control command output by the actuator second-order saturation command filter. The actuator second-order saturation command filter is defined as: Among them, ζ2 and ω n2 They represent the damping coefficient and natural oscillation frequency of the actuator's second-order command filter, respectively.

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