CN107491082A - Spacecraft Attitude Control mixing executing agency optimal control method - Google Patents

Abstract

The present invention discloses a kind of Spacecraft Attitude Control mixing executing agency optimal control method, for single-gimbal control momentum gyro geometry is unusual and the problems such as counteraction flyback dead band, saturation, proposes to make up the optimal control method of deficiency each other using the mixing executing agency that both are formed.Using above-mentioned optimal control method, realize to spacecraft momentum management and stability contorting.The control method can be applied to following quick maneuverable spacecraft and full electricity pushes away Spacecraft During Attitude Maneuver and keep control, realize quick high-precision control, and can improve the life-span of spacecraft.

Description

Optimal Control Method of Hybrid Actuator for Spacecraft Attitude Control

technical field

The invention relates to the field of spacecraft attitude control, in particular to an optimal control method for a spacecraft attitude control hybrid actuator.

Background technique

With the gradual increase in the requirements of space missions, spacecraft with agile maneuverability has become the focus of research since the last century. Especially for next-generation imaging satellites, space missions such as large-angle agile maneuvering, multi-target acquisition and reorientation have become necessary capabilities for obtaining high-resolution images.

Compared with the small torque output capability of the flywheel, the Single Gimbal Control Moment Gyro (SGCMG) has become the main actuator of the agile satellite because of its powerful torque amplification capability. The single-frame control moment gyroscope consists of a frame, a frame motor, a rotor and a rotor motor. When the system is working, the rotor with constant speed is driven by the frame motor to change the direction of the angular momentum of the system, thereby generating output torque. The single-frame control moment gyro has excellent performances such as large output torque, long life, energy saving and high efficiency. Especially because of its powerful torque output capability, it has been widely used, such as the International Space Station and the WorldView series of high-resolution earth imaging satellites. However, a major disadvantage of the control moment gyro system is its inherent geometric singularity. Once the system falls into a singular state, it cannot output the desired torque, which may cause the system to go out of control, which is not allowed in practical engineering applications. Therefore, the singularity analysis of the single-frame control moment gyro system and the research on the corresponding maneuvering strategy have become a research hotspot.

The research shows that the hybrid actuator system with single frame control moment gyro has a certain potential in dealing with singular problems.

Contents of the invention

In order to solve the problems in the prior art, the present invention provides an optimal control method for a spacecraft attitude control hybrid executive mechanism, aiming at agile spacecraft using a control moment gyroscope as an actuator, there is a problem that the control moment gyroscope is singular, and the spacecraft attitude control The hybrid actuator optimization control method enables the two to play their respective characteristics, ensuring that the hybrid actuator system can be free of singularity/saturation for a long time, and complete the high-precision agile attitude maneuver and control of the spacecraft.

The hybrid actuator optimization control method for spacecraft attitude control of the present invention mainly involves the following existing systems: (1) spacecraft angular momentum management system, (2) control moment gyro system, (3) reaction flywheel system, (4) spacecraft attitude Control system, (5) spacecraft attitude measurement and feedback system. It is characterized in that the hybrid actuator optimization control method includes the following steps:

Step 1: In each space mission, according to the target attitude requirements, the onboard control computer of the spacecraft calculates the corresponding control torque _τc sequence, which is used as the torque to be generated by the actuator to control the moment gyro system and the reaction flywheel system.

Step 2: The control torque gyro and reaction flywheel perform a self-test to determine the current degree of singularity (control torque gyro) and saturation (flywheel) of the system.

Step ₃ : The spaceborne control computer calculates the frame angle trajectory δ when only the torque is output by the control torque gyro according to the torque command sequence τc, and at the same time obtains the value sequence S of the singular metric function of the control torque gyro.

Step 3-1: Determine the current frame angle combination δ of the control moment gyro system by the attitude measurement and feedback system;

Step 3-2: Execute the singularity metric function to obtain the singularity degree of the control moment gyro system of the current pyramid configuration:

S=det(J ^T J), where J∈R ^3×4 is the Jacobian matrix of the control moment gyro system, and the current frame angle of the control moment gyro δ=(δ ₁ ,δ ₂ ,δ ₃ ,δ ₄ ) Sure.

Step 4: According to the singularity metric sequence S, judge the degree of singularity of the control moment gyroscope at any moment in the entire task period. If it exceeds the initial singularity threshold Then it is considered that the system is in a singular state, record the first singular moment as t _s , and go to step 5, otherwise go to step 9.

Step 5: The system enters the singularity correction stage of the control torque gyro, and adds a small singularity correction torque to the torque command at time Δt before time _ts Get a new sequence of torque commands Among them, the singular correction torque will be generated by the flywheel system. Execute step 3, step 4 and step 5 until the control moment gyro system is far away from singularity in the whole control cycle. Then, go to step 6.

Step 6: According to the correction torque sequence obtained in step 5 and the current state of the flywheel system, calculate the angular velocity sequence Ω when the flywheel system outputs the correction torque _TM . If the initial angular velocity of the flywheel meets the output requirement of _TM , go to step 8, otherwise go to step 7.

Step 7: According to the requirements of the initial state of the flywheel angular velocity sequence Ω, when the control torque gyro system does not fall into a singular state, add the flywheel initial state correction torque N _M to the spacecraft control torque sequence _τc , and adjust the flywheel initial angular velocity to meet the control torque Gyro singularity correction needs. Then, go to step 8.

Step 8: Determine the new spacecraft control torque sequence τ _c according to the final control torque gyro singular correction torque and the flywheel initial state correction torque N _M , and perform step 9.

Step 9: The frame angular velocity command sequence determined according to the control moment gyro steering law and the flywheel acceleration command determined by the flywheel control law In sequence, the control moment gyro system and the flywheel system are activated to generate output torques N _CMG and N _RW , which act on the spacecraft for attitude control and maneuvering.

Step 10: At the end of the spacecraft attitude maneuver mission cycle, turn off the control moment gyro system. And the moderate output torque of the flywheel system is used for attitude correction and fine alignment.

The beneficial effect of the invention is that: similar to the single-frame control moment gyro system, although the reaction flywheel system has no singularity, it has dead zone and saturation characteristics. The invention adopts a 4-SGCMG system with a pyramid configuration and an orthogonal configuration 3-RW to form a spacecraft attitude control hybrid actuator. The purpose is to use the flywheel system to weaken and overcome the singularity problem of the control moment gyro system, and at the same time use the control moment gyro system to adjust the flywheel, so that the two can play their respective characteristics to ensure that the hybrid actuator system can have no singularity for a long time, and complete the high-speed spacecraft. Accurate, agile, and maneuverable.

Description of drawings

Figure 1 is a block diagram of the hybrid actuator control program;

Figure 2 is a control logic diagram of a spacecraft containing a hybrid mechanism;

Figure 3 is a schematic diagram of the installation of the 4-SGCMG system and the 3-RW system in a pyramid configuration.

detailed description

The present invention will be further described below in conjunction with accompanying drawing.

The present invention proposes an optimal control method for a spacecraft attitude control hybrid executive mechanism, the control program block diagram is shown in Figure 1, and the control logic schematic diagram is shown in Figure 2, mainly related to the following existing systems: (1) spacecraft angular momentum management system, ( 2) Control moment gyro system, (3) reaction flywheel system, (4) spacecraft attitude control system, (5) spacecraft attitude measurement and feedback system. It is characterized in that the hybrid actuator optimization control method includes the following steps:

Step 1: In each space mission, according to the requirements of the target attitude, the onboard control computer of the spacecraft calculates the corresponding control torque _τc sequence, which is used as the torque to be generated by the actuator to control the moment gyro system and the reaction flywheel system;

Step 2: The control torque gyro and reaction flywheel perform a self-test to determine the current degree of singularity (control torque gyro) and saturation (flywheel) of the system. The self-test procedure is specifically:

(1) Calculate the degree to which the current control moment gyro system is away from the singular state

S＝det(J ^T J)

Where J is the Jacobian matrix of the SGCMG system, and J is the frame angle function that is J=J(δ);

(2) Calculate the distance saturation degree of the current flywheel system, obtain the angular velocity Ω, and execute the following function to determine the saturation degree,

where Ω＝[Ω ₁ Ω ₂ Ω ₃ ] ^T is the angular velocity of each flywheel in the flywheel system, Q∈R ^3×3 pair weight matrix, is a weighted two-norm, ||Ω ₀ || _∞ ＝max{Ω _i ,i=1,2,3} is an infinite norm;

Step ₃ : The spaceborne control computer calculates the frame angle trajectory δ when only the torque is output by the control torque gyro according to the torque command sequence τc, and at the same time obtains the value sequence S of the singular metric function of the control torque gyro.

Step 3-1: Determine the current frame angle combination δ of the control moment gyro system by the attitude measurement and feedback system;

Step 3-2: Execute the singularity metric function to obtain the singularity degree of the control moment gyro system of the current pyramid configuration:

S=det(J ^T J), where J∈R ^3×4 is the Jacobian matrix of the control moment gyro system, and the current frame angle of the control moment gyro δ=(δ ₁ ,δ ₂ ,δ ₃ ,δ ₄ ) Sure.

Step 4: According to the singularity metric sequence S, judge the degree of singularity of the control moment gyroscope at any moment in the entire task period. If it exceeds the initial singularity threshold Then it is considered that the system is in a singular state, record the first singular moment as t _s , and go to step 5, otherwise go to step 9;

Step 5: The system enters the singularity correction stage of the control torque gyro, and adds a small singularity correction torque to the torque command at time Δt before time _ts Get a new sequence of torque commands Among them, the singular correction torque will be generated by the flywheel system. Execute step 4 and step 5 until the control moment gyro system is far away from singularity in the whole control cycle. Then, go to step 6.

Step 6: According to the correction torque sequence obtained in step 5 and the current state of the flywheel system, calculate the angular velocity sequence Ω when the flywheel system outputs the correction torque _TM . If the initial angular velocity of the flywheel meets the output requirement of _TM , go to step 8, otherwise go to step 7.

Step 7: According to the requirements of the initial state of the flywheel angular velocity sequence Ω, when the control torque gyro system does not fall into a singular state, add the flywheel initial state correction torque N _M to the spacecraft control torque sequence _τc , and adjust the flywheel initial angular velocity to meet the control torque Gyro singularity correction needs. Then, go to step 8.

Step 8: Determine the new spacecraft control torque sequence τ _c according to the final control torque gyro singular correction torque and the flywheel initial state correction torque N _M , and perform step 9.

Step 9: The frame angular velocity command sequence determined according to the control moment gyro steering law and the flywheel acceleration command sequence determined by the flywheel control law The control moment gyro system and the flywheel system are activated to generate output torques N _CMG and N _RW , which act on the spacecraft for attitude control and maneuvering.

Step 10: At the end of the spacecraft attitude maneuver mission cycle, turn off the control moment gyro system. And the moderate output torque of the flywheel system is used for attitude correction and fine alignment.

The invention proposes an optimal control method for a hybrid executive mechanism based on a reaction flywheel (Reaction Wheel, RW) and a single-frame control moment gyroscope. Similar to the single-frame control-moment gyro system, the reaction flywheel system is non-singular, but has dead-band and saturation characteristics. The present invention adopts a 4-SGCMG system with a pyramid configuration and an orthogonal configuration 3-RW to form a hybrid actuator for spacecraft attitude control, as shown in Figure 3, and aims to use the flywheel system to weaken and overcome the singularity of the control moment gyro system At the same time, the control moment gyro system is used to adjust the flywheel, so that the two can play their respective characteristics, ensuring that the hybrid actuator system can be stable for a long time, and complete the high-precision and agile attitude maneuver of the spacecraft.

There are many specific application approaches of the present invention, and the above description is only a preferred embodiment of the present invention. It should be pointed out that for those of ordinary skill in the art, some improvements can also be made without departing from the principles of the present invention. Improvements should also be regarded as the protection scope of the present invention.

Claims (3)

1. a spacecraft attitude control hybrid executive mechanism optimal control method, is characterized in that comprising the following steps: Step 1: In each space mission, according to the target attitude Requirements, the corresponding control torque τ _c sequence is calculated by the spacecraft on-board control computer, as the torque that needs to be generated by the actuator to control the torque gyro system and the reaction flywheel system; Step 2: Control the torque gyro and reaction flywheel to perform a self-test to determine the current degree of singularity and saturation of the system; Step ₃ : The spaceborne control computer calculates the frame angle trajectory δ when only the torque is output by the control torque gyro according to the torque command sequence τc, and at the same time obtains the value sequence S of the singular metric function of the control torque gyro; Step 4: According to the singularity measurement sequence S, judge the degree of singularity of the control torque gyroscope at any time in the entire task cycle, if it exceeds the initial singularity threshold Then it is considered that the system is in a singular state, record the first singular moment as t _s , and go to step 5, otherwise go to step 9; Step 5: The system enters the singularity correction stage of the control torque gyro, and adds a small singularity correction torque to the torque command at time Δt before time _ts Get a new sequence of torque commands Wherein the singular correction torque will be generated by the flywheel system, perform steps 3, 4 and 5 until the control torque gyro system is away from the singularity in the entire control cycle, and then perform step 6; Step 6: According to the corrected torque sequence obtained in step 5 and the current state of the flywheel system, solve the angular velocity sequence Ω when the flywheel system outputs the corrected torque _TM . If the initial angular velocity of the flywheel meets the output requirements of _TM , then perform step 8. Otherwise, go to step 7; Step 7: According to the requirements of the initial state of the flywheel angular velocity sequence Ω, when the control torque gyro system does not fall into a singular state, add the flywheel initial state correction torque N _M to the spacecraft control torque sequence _τc , and adjust the flywheel initial angular velocity to meet the control torque Gyro singularity correction needs, then, go to step 8; Step 8: Determine the new spacecraft control torque sequence τ _c according to the final control torque gyro singular correction torque and the flywheel initial state correction torque N _M , and perform step 9; Step 9: The frame angular velocity command sequence determined according to the control moment gyro steering law and the flywheel acceleration command determined by the flywheel control law sequence, start the control moment gyro system and the flywheel system, generate output torques N _CMG and N _RW , act on the spacecraft, and perform attitude control and maneuvering; Step 10: At the end of the attitude maneuvering mission cycle of the spacecraft, turn off the control moment gyro system, and use the moderate output torque of the flywheel system for attitude correction and fine alignment. 2. The spacecraft attitude control hybrid actuator optimization control method according to claim 1, characterized in that, the self-test method described in the step 2 is: Step, 2-1: Calculate the degree to which the current control moment gyro system is away from the singular state, S＝det(J ^T J) Where J is the Jacobian matrix of the SGCMG system, and J is the frame angle function that is J=J(δ); Step 2-2: Calculate the distance saturation degree of the current flywheel system, obtain the angular velocity Ω, and execute the following function to determine the saturation degree, where Ω＝[Ω ₁ Ω ₂ Ω ₃ ] ^T is the angular velocity of each flywheel in the flywheel system, Q∈R ^3×3 pair weight matrix, is a weighted two-norm, and ||Ω ₀ || _∞ =max{Ω _i , i=1,2,3} is an infinite norm. 3. the spacecraft attitude control hybrid executive mechanism optimal control method according to claim 1, is characterized in that, the determination method of singularity metric value in the described step 3 is: Step 3-1: Determine the current frame angle combination δ of the control moment gyro system by the attitude measurement and feedback system; Step 3-2: Execute the singularity metric function to obtain the singularity degree of the control moment gyro system of the current pyramid configuration: S＝det(J ^T J) In the formula, J∈R ^3×4 is the Jacobian matrix of the control moment gyro system, which is determined by the current frame angle of the control moment gyro δ=(δ ₁ ,δ ₂ ,δ ₃ ,δ ₄ ).