Double-satellite multi-loop cooperative control method based on non-contact load vibration isolation system

Abstract

The invention discloses a double-satellite multi-loop cooperative control method based on a non-contact load vibration isolation system. The method comprises the following steps: dividing a whole control system into a satellite platform attitude control loop, a follow-up control loop of a load following satellite platform and a load cabin medium-high frequency vibration suppression loop; determining a follow-up control loop control law of a load following satellite platform; determining a satellite platform attitude control loop control law; determining a high-frequency vibration suppression loop control law in the load cabin; and the requirements of vibration isolation and high-precision rapid pointing adjustment of the load are met through a loop synergistic effect. According to the control method, the multi-time-scale system cooperative control problem of a satellite platform attitude control loop, a follow-up control loop of a load following satellite platform and an active vibration control loop is considered, and the three loops act synergistically, so that the control target of high precision and high stability is achieved.

Description

Technical Field

[0001] The present invention relates to a dual-satellite multi-loop cooperative control method based on a non-contact load vibration isolation system, belonging to the field of control technology. Background Art

[0002] Space missions, such as those involving space telescopes, laser communications, and high-precision Earth observation, place extremely high demands on the accuracy, stability, and agility of control systems for space-based satellite observation platforms. Currently, two main factors hinder the high-performance control accuracy, stability, and agility of satellite platforms: 1) Low-frequency vibrations (0.1-10 Hz) of flexible attachments such as solar panels, combined with high-speed rotation of onboard rotating components, payload scanning mechanism rotation, stepping motion of large controllable component drive mechanisms, thruster ignition during orbital maneuvers, mechanical motion of thermal control components, stimulated vibrations of large flexible structures, and thermal deformation disturbances induced by alternating hot and cold temperatures during entry and exit from shadows, induce a low-amplitude, high-frequency flutter response (10-200 Hz) in the spacecraft, severely impacting the platform's pointing accuracy and stability. 2) The high speed and dynamic delays of the spacecraft during motion affect pointing accuracy and agility.

[0003] In order to solve the impact of the above two factors on the satellite platform, adding a payload pointing control system with vibration isolation, vibration suppression and rapid maneuverability between the satellite platform and the payload to form a multi-level composite control system for the entire satellite has become a new option. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to overcome the deficiencies of the prior art and provide a dual-satellite multi-loop cooperative control method based on a non-contact load vibration isolation system.

[0005] The purpose of the present invention is achieved through the following technical solutions:

[0006] A dual-satellite multi-loop cooperative control method based on a non-contact load vibration isolation system comprises:

[0007] The entire control system is divided into a satellite platform attitude control loop, a payload follow-up control loop for the satellite platform, and a high-frequency vibration suppression loop in the payload cabin.

[0008] Determine the control law of the follow-up control loop for the payload to follow the satellite platform;

[0009] Determine the control law of the satellite platform attitude control loop;

[0010] Determine the control law for the high-frequency vibration suppression loop in the payload compartment;

[0011] The synergistic effect of the circuits satisfies the vibration isolation and high-precision and rapid pointing adjustment requirements of the load.

[0012] Compared with the prior art, the present invention has the following beneficial effects:

[0013] (1) The present invention is based on a magnetic levitation non-contact load vibration isolation system, and adds a load pointing control system with vibration isolation, vibration suppression and rapid maneuvering capabilities between the satellite platform and the payload to form a multi-level composite control system for the entire satellite.

[0014] (2) The present invention is based on a non-contact load vibration isolation system supported by magnetic levitation. Its multi-loop system control scheme utilizes the multi-loop collaborative control of satellite platform attitude control, follow-up control loop and active pointing ultra-quiet platform to achieve load vibration isolation and pointing adjustment functions, meeting the requirements of fast, accurate and stable attitude control.

[0015] (3) For the non-contact load vibration isolation system based on magnetic levitation support, the present invention establishes a dynamic model of the non-contact load vibration isolation system based on magnetic levitation support, laying a design foundation for the follow-up control loop of the load following satellite platform.

[0016] (4) The control method of the present invention takes into account the multi-time scale system collaborative control problem of the satellite platform attitude control loop, the payload following satellite platform follow-up control loop and the active vibration control loop. The three loops work together to achieve the control goal of high precision and high stability. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 Schematic diagram of a dual-body satellite based on a non-contact load vibration isolation system according to the present invention.

[0018] Figure 2 Schematic diagram of the non-contact load vibration isolation system of the present invention.

[0019] Figure 3 This is a block diagram of the dual-satellite multi-loop collaborative control system of the present invention. DETAILED DESCRIPTION

[0020] To make the objectives, technical solutions and advantages of the present invention more clear, the embodiments of the present invention will be described in further detail below with reference to the accompanying drawings.

[0021] A multi-loop collaborative control method for twin-body satellites based on a non-contact load vibration isolation system can be used for a high-precision, rapid maneuvering pointing control system for twin-body satellites. It mainly solves the problem of high-precision, rapid maneuvering pointing control of twin-body satellites based on a non-contact load vibration isolation system under the condition that there is a vibration disturbance source in the platform cabin.

[0022] The present invention will be described in further detail below with reference to the accompanying drawings.

[0023] Figure 1The schematic diagram of the double-body satellite based on the non-contact load vibration isolation system is shown in Figure 1. A is the satellite platform cabin, B is the non-contact load vibration isolation system, and C is the load cabin. Figure 2 As shown in FIG, the non-contact load vibration isolation system consists of four symmetrically installed magnetic bearings, which connect the satellite platform cabin and the payload cabin.

[0024] like Figure 3 As shown in the figure, the entire control system is divided into a satellite platform attitude control loop, a follow-up control loop for the payload to follow the satellite platform, and a high-frequency vibration suppression loop in the payload cabin, which work together to achieve the vibration isolation of the payload and high-precision and rapid pointing adjustment requirements.

[0025] Satellite platform attitude control loop: Utilize attitude measurement sensors such as the platform cabin star sensor and gyroscope to obtain platform cabin attitude measurement information, design the platform cabin's high-precision attitude maneuvering control law, thereby obtaining the desired instructions for the attitude control actuator, and ultimately outputting the control torque through attitude control actuators such as flywheels.

[0026] The follow-up control loop of the payload-following satellite platform: According to the relative posture dynamic model and magnetic equation of the float and stator, the relative posture and position measurement information is solved based on the relative position sensor of the magnetic levitation bearing float and stator. Combined with the attitude change of the platform cabin, the relative posture controller is divided into a combination of feedback controller and feedforward controller. A second-order full-drive linear quadratic regulation control method is designed to obtain the optimal virtual control law, and finally the relative posture current control law is obtained, which is the follow-up control current of the magnetic bearing, meeting the high-precision relative posture control requirements under maneuvering conditions, ensuring that the relative posture of the two cabins is close to 0 and there is no collision between the two cabins.

[0027] The medium- and high-frequency vibration suppression loop in the payload cabin: The absolute velocity sensor of the payload platform is used to obtain the medium- and high-frequency vibration measurement information of the payload platform (including the three-axis inertial angular velocity and translational velocity), and the medium- and high-frequency active vibration suppression current control law is designed. The medium- and high-frequency active vibration suppression current control law works synergistically with the relative posture current control law of the payload platform cabin to jointly achieve the control goals of high precision and high stability.

[0028] Since the mid- and high-frequency vibration suppression controller has the shortest control period, typically reaching 1000Hz, while the servo control loop has a relatively moderate control period of 100Hz, and the satellite platform attitude control loop has the longest control period of 10Hz, the dual-body satellite multi-loop coordinated control system is a multi-timescale system. Therefore, based on singular perturbation theory, the three loops are coupled to maintain the state of the slow-varying system unchanged in the fast-varying system, while the fast-varying parameters in the slow-varying system are equivalent. The three loops work together to achieve high-precision and high-stability control.

[0029] (1) The payload follows the satellite platform's follow-up control loop: The mass of the payload cabin is defined as MC and the moment of inertia is J C The relative attitude angle between the payload cabin coordinate system and the platform cabin coordinate system is χ CD ∈R 3 , platform cabin center of mass O D To the center of mass of the payload compartment O C The vector is r CD ∈R 3 , the angular velocity of the payload cabin relative to the platform cabin is ω CD ∈R 3 , the relative posture state of the payload platform cabin x1=[r CD T χ CD T ] T , the relative position change variable of the payload platform cabin is e1=x1-x d , where x d ∈R 6 It is expected that the relative posture of the two cabins will be maintained.

[0030] definition M f =diag(M C ,J C ),ω D is the three-axis attitude angular velocity of the platform cabin in the inertial system, is the three-axis attitude angular acceleration of the platform cabin in the inertial system, r D is the center of mass of the Earth in the inertial system O N To the center of mass of the payload compartment O C The kinematic and dynamic equations of the relative posture error of the payload platform cabin can be written as

[0031]

[0032] in, is the coupling term between the platform cabin attitude and relative attitude, N(e1,r D ) is the coupling term between the platform cabin orbital motion and relative position, is the term related to the angular acceleration of the platform cabin attitude. The 8 magnetic forces generated by the 4 magnetic bearings, C A is the conversion matrix from 8 magnetic forces to 6-DOF forces / torques, F C is the conversion matrix from the 6-DOF force / torque at the center of the magnetic levitation mechanism to the center of mass of the payload compartment.

[0033] Since the coil current adopts differential form, the linear expression of the i-th single electromagnetic force is:

[0034] f Li (Δδ i ,i ci)=k1i ci +k2Δδ i

[0035] Where: k1 is the current stiffness coefficient, k2 is the displacement stiffness coefficient, Δδ i is the change of the ith magnetic levitation gap, and the output current of the ith magnetic levitation controller is i ci , i=1……8. Define the control current of the magnetic levitation system I=[i c1 i c2 i c3 i c4 i c5 i c6 i c7 i c8 ] T , the maglev system gap change Δδ = [Δδ1 Δδ2 Δδ3 Δδ4 Δδ5 Δδ6 Δδ7 Δδ8] T ,but

[0036] F L =k1I+k2Δδ

[0037] The relationship between the gap change and the relative position change of the load platform cabin can be obtained by deduction:

[0038] e1=R x2d Δδ, Δδ=R d2x e1

[0039] Among them, R x2d is the conversion matrix from the gap change of the maglev system to the relative position change of the payload platform cabin, R d2x The conversion matrix from the relative posture change of the payload platform cabin to the gap change of the magnetic levitation system can be used to calculate the relative posture and position measurement information using the relative position sensor between the magnetic levitation bearing float and the stator, and further design the relative posture controller:

[0040]

[0041] Where G1 is the control input matrix, v is the virtual control input, I cf is the current feedback control law, I cq As the current feedforward control law, the second-order all-wheel drive linear model of the relative posture of the payload platform cabin can be obtained as

[0042]

[0043] The second-order full-drive linear quadratic regulation control method is adopted, and the relative posture error and error speed are selected as the state variables e=[e1 T e2 T ] T , the system can be transformed into

[0044]

[0045] Among them, the system matrices A and B are

[0046]

[0047] The optimal performance index is selected as

[0048]

[0049] Where Q∈R 12×12 and R c =R 6×6 The weighted matrices are semi-positive definite and positive definite respectively. By selecting appropriate Q and R, the system steady-state error can be reduced to meet the control accuracy requirements and obtain better dynamic response within the range allowed by the actuator.

[0050] Get the optimal controller

[0051] v * =-R c -1 B T Pe

[0052] Where P is the only positive definite solution of the Riccati matrix algebraic equation

[0053] A T P+PA-PB T R c -1 BP+Q=0

[0054] Thus we get the virtual optimal controller v * , then the relative posture current control law It meets the requirements of high-precision relative attitude control under maneuvering conditions, ensuring that the relative attitude of the two cabins is close to 0 and there is no collision between the two cabins.

[0055] (2) For the satellite platform attitude control loop, the platform cabin inertial angular velocity ω is obtained using attitude measurement sensors such as the platform cabin gyroscope D , the star sensor obtains attitude quaternion information q D , the control law of the attitude maneuver tracking control loop of the platform cabin is designed as

[0056] u D =-k d ω D -k p q D

[0057] Among them, u DThe high-precision attitude maneuver control law is designed to obtain the desired instruction of the attitude control actuator, and finally output the control torque through the attitude control actuator such as the flywheel. d and k p is a positive definite control matrix to be designed.

[0058] (3) For the high-frequency vibration suppression circuit of the payload cabin, the three-axis inertial velocity ξ of the high-frequency payload cabin is measured by the absolute velocity sensor installed on the payload platform. C =[ω C T v C T ] T ,ω C =[ω C1 ω C2 ω C2 ] T is the angular velocity of the medium and high frequency payload cabin in the inertial system, ω C1 ω C2 ω C2 are the three components of angular velocity, v C =[v C1 v C2 v C2 ] T is the translational velocity of the medium and high frequency payload cabin in the inertial system, v C1 v C2 v C2 For the three components of speed, design the current control law for medium and high frequency active vibration suppression

[0059] I g =c m G1 -1 ξ C

[0060] where c m The damping matrix to be designed is the medium and high frequency active vibration suppression current control law and the relative posture current control law of the payload platform cabin work together to achieve the control goal of high precision and high stability.

[0061] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

[0062] Although the present invention has been disclosed above in terms of preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solutions of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the scope of protection of the technical solutions of the present invention.

Claims

1. A dual-satellite multi-loop cooperative control method based on a non-contact load vibration isolation system, characterized in that: include: The entire control system is divided into a satellite platform attitude control loop, a payload follow-up control loop for the satellite platform, and a high-frequency vibration suppression loop in the payload cabin. Determine the control law of the follow-up control loop for the payload to follow the satellite platform; Determine the control law of the satellite platform attitude control loop; Determine the control law for the high-frequency vibration suppression loop in the payload compartment.

2. The dual-satellite multi-loop coordinated control method according to claim 1, characterized in that: For the satellite platform attitude control loop, the platform cabin attitude measurement information is used to determine the high-precision attitude maneuver control law of the platform cabin, thereby obtaining the desired instructions for the attitude control actuator, and finally outputting the control torque through the attitude control actuator.

3. The dual-satellite multi-loop coordinated control method according to claim 2, characterized in that: The platform cabin uses attitude measurement sensors to obtain platform cabin attitude measurement information.

4. The dual-satellite multi-loop coordinated control method according to claim 1, characterized in that: For the follow-up control loop of the payload following satellite platform, according to the relative posture dynamic model and magnetic equation of the float and stator, the relative posture and position measurement information is solved based on the relative position sensor of the magnetic levitation bearing float and stator. Combined with the attitude change of the platform cabin, the relative posture controller is divided into a combination of feedback controller and feedforward controller. The second-order full-drive linear quadratic regulation control method is adopted to obtain the optimal virtual control law, and further obtain the relative posture current control law, which is the follow-up control current of the magnetic bearing, to meet the high-precision relative posture control requirements under maneuvering conditions.

5. The dual-satellite multi-loop coordinated control method according to claim 4, characterized in that: The relative posture controller is divided into a combination of feedback controller and feedforward controller as follows: Where, I c is the relative pose controller, G1 is the control input matrix, v is the virtual control input, I cf is the current feedback control law, I cq is the current feedforward control law, is the coupling term between the platform cabin attitude and relative attitude, N(e1,r D ) is the coupling term between the platform cabin orbital motion and relative position, is the term related to the platform cabin attitude angular acceleration, is the three-axis attitude angular acceleration of the platform cabin in the inertial system, r D is the center of mass of the Earth in the inertial system O N To the center of mass of the payload compartment O C vector, e1 is the relative position change variable of the payload platform cabin, and the center of mass of the platform cabin O D To the center of mass of the payload compartment O C The vector is r CD , the angular velocity of the payload cabin relative to the platform cabin is ω CD .

6. The dual-satellite multi-loop coordinated control method according to claim 5, characterized in that: The second-order full-drive linear quadratic regulation control method is adopted, and the relative posture error and error speed are selected as the state variables e=[e1 T e2 T ] T , the system is transformed into Among them, v is the virtual control input, and the system matrices A and B are M f =diag(M C ,J C ) Among them, M C is the mass of the payload cabin, J C is the moment of inertia, k2 is the displacement stiffness coefficient, F C is the conversion matrix from the 6-DOF force / torque at the center of the magnetic levitation mechanism to the center of mass of the payload cabin, C A is the conversion matrix from 8 magnetic forces to 6-DOF forces / torques, R d2x The conversion matrix from the relative position change of the payload platform cabin to the gap change of the magnetic levitation system; The optimal performance index is selected as Where Q∈R 12×12 and R c =R 6×6 are weight matrices that are semi-positive definite and positive definite respectively; Get the optimal controller: V * =-R c -1 B T On Where P is the only positive definite solution of the Riccati matrix algebraic equation A T P+PA-PB T R c -1 BP+Q=0 Thus we get the virtual optimal controller v * .

7. The dual-satellite multi-loop coordinated control method according to claim 6, characterized in that: Relative posture current control law 8. The dual-satellite multi-loop coordinated control method according to claim 4, characterized in that: For the medium and high frequency vibration suppression circuit of the payload cabin, the absolute velocity sensor of the payload platform is used to obtain the medium and high frequency vibration measurement information of the payload platform, and determine the medium and high frequency active vibration suppression current control law. The medium and high frequency active vibration suppression current control law works together with the relative posture current control law of the payload platform cabin to jointly achieve the control goals of high precision and high stability.

9. The dual-satellite multi-loop coordinated control method according to claim 1, characterized in that: The control law of the attitude maneuver tracking control loop of the platform cabin is: u D =-k d ω D -k p q D Among them, u D is the designed high-precision attitude maneuver control law, k d and k p is the positive definite control matrix to be designed, ω D is the inertial angular velocity of the platform cabin.

10. The dual-satellite multi-loop coordinated control method according to claim 1, characterized in that: High-frequency active vibration suppression current control law I g =c m G1 -1 ξ C where c m is the damping matrix to be designed, ξ C is the three-axis inertial velocity of the medium and high frequency payload cabin, and G1 is the control input matrix.

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