Satellite in-orbit identification and feedforward feedback combined control method based on magnetic suspension bearing

Abstract

The invention provides a satellite in-orbit identification and feedforward feedback combined control method based on a magnetic suspension bearing, which is applied to a satellite of which a satellite platform and a load platform are both large-inertia rigid bodies. The method comprises the following steps: according to a first motion parameter of a load platform, a second motion parameter of a satellite platform, a position vector and the mass of the load platform in a rotation process, determining a position disturbance force and an attitude disturbance torque to the load platform and a disturbance torque to the satellite platform in the rotation process; the disturbance torque comprises a first disturbance torque introduced by non-coincidence of the mass center of the load platform and the rotating shaft, a second disturbance torque introduced by non-uniform mass distribution of the load platform in the rotating shaft direction, and a third disturbance torque introduced by non-uniform mass distribution of the load platform in the non-rotating shaft direction; and performing combined control on the satellite according to the position disturbance force and the attitude disturbance torque of the load platform and the disturbance torque of the satellite platform. According to the scheme, in-orbit identification is carried out on interference terms in the rotation process, feedforward control is designed for the controller, and the satellite attitude control precision and stability index are improved.

Description

Technical Field

[0001] The present invention relates to the field of satellite control technology, in particular to the field of satellite attitude control technology, and in particular to a satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings. Background Art

[0002] The magnetic bearing joint completes the connection between the stable satellite platform and the high-speed rotating load platform system. The magnetic bearing outputs the control force and control torque to complete the stable control of the load platform and the satellite platform. As the satellite imaging mission puts forward the requirement of high-speed rotation of the load platform, the load platform used is of higher quality. The large mass and high dynamic characteristics of the load platform determine that the interference characteristics of the load platform and the satellite platform are obvious during the high-speed rotation process. In addition, the motion characteristics of the satellite platform and the load platform determine that the interference characteristics of the load platform and the satellite platform are not the same. In order to ensure the accuracy and stability of the satellite attitude control, it is necessary to perform interference identification and feedforward controller design on the controllers of the load platform and the satellite platform respectively. Therefore, it is urgent to provide a satellite on-orbit identification and feedforward feedback joint control method based on magnetic bearings. Summary of the invention

[0003] The present invention provides a satellite on-orbit identification and feedforward feedback joint control method based on magnetic levitation bearings. By decomposing the position and attitude interference of the load platform and the satellite platform during the high-speed rotation of the load platform, interference identification and feedforward control are designed for the controller to improve the satellite attitude control accuracy and stability index.

[0004] In a first aspect, the present invention provides a satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings, which is applied to a satellite in which both the satellite platform and the payload platform are large inertia rigid bodies. The method comprises:

[0005] Acquire a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform;

[0006] According to the first motion parameter, the second motion parameter, the position vector and the mass of the load platform, determine the position interference force, attitude interference torque and interference torque on the load platform caused by the rotation process, wherein the interference torque includes a first interference torque introduced by the misalignment of the center of mass of the load platform and the rotation axis, a second interference torque introduced by the uneven mass distribution of the load platform in the direction of the rotation axis, and a third interference torque introduced by the uneven mass distribution of the load platform in the direction of the non-rotation axis;

[0007] The satellite is jointly controlled according to the position disturbance force, attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

[0008] In a second aspect, the present invention further provides a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings, which is applied to a satellite in which both the satellite platform and the payload platform are large inertia rigid bodies, comprising:

[0009] An acquisition module, used for acquiring a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform;

[0010] An on-orbit identification module is used to determine the position interference force, attitude interference torque and interference torque of the payload platform caused by the rotation process, and the satellite platform according to the first motion parameter, the second motion parameter, the position vector and the mass of the payload platform; wherein the interference torque includes a first interference torque introduced by the misalignment between the center of mass of the payload platform and the rotation axis, a second interference torque introduced by the uneven mass distribution of the payload platform in the direction of the rotation axis, and a third interference torque introduced by the uneven mass distribution of the payload platform in the direction of the non-rotation axis;

[0011] A control module is used to jointly control the satellite according to the position disturbance force and attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

[0012] In a third aspect, the present invention further provides a computing device, comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements any of the above-mentioned methods for in-orbit identification and feedforward-feedback joint control of satellites based on magnetic levitation bearings.

[0013] In a fourth aspect, the present invention further provides a computer-readable storage medium having a computer program stored thereon. When the computer program is executed in a computer, the computer is caused to execute any of the above-mentioned methods for in-orbit identification and feedforward-feedback joint control of a satellite based on a magnetic levitation bearing.

[0014] In a fifth aspect, an embodiment of the present invention further provides a computer program product, comprising computer instructions, which, when executed by a processor, implement the steps of the method described in any first aspect of this specification.

[0015] The present invention provides a satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings, which is applied to satellite platforms and load platforms that are both large inertia rigid bodies connected by magnetic suspension joints. The method analyzes the motion parameters and position vectors of the satellite platform and the load platform, and sequentially determines the position interference force and attitude interference torque of the load platform during the rotation of the large inertia load platform, as well as the interference torque of the satellite platform introduced by the misalignment of the center of mass of the load platform and the rotation axis, the uneven mass distribution of the load platform in the direction of the rotation axis, and the uneven mass distribution of the load platform in the direction of the non-rotation axis, and then uses feedforward and feedback joint control to realize control of the satellite based on the position interference force of the load platform, the attitude interference torque and the interference torque of the satellite platform. In this way, the present invention determines these interferences introduced by the large inertia load through on-orbit identification, and uses a joint control method of feedforward combined with feedback to process these interferences, thereby improving the control accuracy and ensuring the attitude accuracy and stability index of the satellite during the on-orbit operation. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0017] Figure 1 It is a flow chart of a satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings provided by one embodiment of the present invention;

[0018] Figure 2 is a schematic diagram of a reference coordinate system applied to a satellite system provided by an embodiment of the present invention;

[0019] Figure 3 It is a schematic diagram of the working principle of radial eddy current measurement provided by one embodiment of the present invention;

[0020] Figure 4 It is a schematic diagram of relative posture position decoupling calculation within a radial measurement section provided by an embodiment of the present invention;

[0021] Figure 5 is a schematic diagram of attitude interference of a satellite platform provided by an embodiment of the present invention;

[0022] Figure 6 It is a schematic diagram of the pitch attitude control effect of 500s pre-feedback control and 500s post-joint control provided by an embodiment of the present invention;

[0023] Figure 7It is a schematic diagram of the pitch attitude angular velocity control effect of 500s pre-feedback control and 500s post-joint control provided by an embodiment of the present invention;

[0024] Figure 8 is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0025] Fig. 9 It is a structural diagram of a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings provided in one embodiment of the present invention. DETAILED DESCRIPTION

[0026] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0027] Please refer to Figure 1 The embodiment of the present invention provides a satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings, which is applied to a satellite in which both the satellite platform and the payload platform are large inertia rigid bodies, including:

[0028] Step 100, obtaining a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform;

[0029] Step 102, determining the position disturbance force, attitude disturbance torque and disturbance torque of the load platform and the satellite platform caused by the rotation process according to the first motion parameter, the second motion parameter, the position vector and the mass of the load platform; wherein the disturbance torque includes a first disturbance torque introduced by the misalignment between the center of mass of the load platform and the rotation axis, a second disturbance torque introduced by the uneven mass distribution of the load platform in the direction of the rotation axis and a third disturbance torque introduced by the uneven mass distribution of the load platform in the direction of the non-rotation axis;

[0030] Step 104 , jointly controlling the satellite according to the position disturbance force and attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

[0031] In the present invention, in the satellite platform and the load platform, both of which are large inertia rigid bodies connected by magnetic levitation joints, the motion parameters and position vectors of the satellite platform and the load platform are analyzed, and the position interference force and attitude interference torque of the load platform during the rotation of the large inertia load platform are determined in turn, as well as the interference torque of the satellite platform introduced by the non-coincidence of the center of mass of the load platform and the rotation axis, the uneven mass distribution of the load platform in the direction of the rotation axis, and the uneven mass distribution of the load platform in the direction of the non-rotation axis, and then the satellite is controlled based on the position interference force, attitude interference torque of the load platform and the interference torque of the satellite platform by feedforward and feedback joint control. In this way, the present invention determines these interferences introduced by the large inertia load through on-orbit identification, and uses a joint control method of feedforward combined with feedback to process these interferences, thereby improving the control accuracy and ensuring the attitude accuracy and stability index of the satellite during the on-orbit operation.

[0032] It should be noted that the payload platform and the satellite platform are connected via a connection point (eg, a hinge connection point), and the satellite includes the payload platform and the satellite platform. The large inertia payload is preferably such that the mass of the payload platform itself is equivalent to the mass of the satellite platform.

[0033] Described below Figure 1 How the various steps are performed.

[0034] The coordinate systems used in the present invention include the geocentric inertial coordinate system, the satellite platform coordinate system, the whole satellite platform coordinate system and the payload platform coordinate system. Figure 2 As shown, for the geocentric inertial coordinate system, with the origin O E is the center of mass of the Earth, O E X E The axis points to the 2000.0 mean vernal equinox, O E Z E The axis is perpendicular to the 2000.0 flat equatorial plane and is consistent with the direction of the Earth's rotation angular velocity. E Y E Axis and O E X E , O E Z E The axes form a right-hand system;

[0035] For the satellite platform coordinate system, the origin O is the center of the connection surface between the satellite and the carrier, O x The axis is along the longitudinal axis of the satellite body, pointing to the payload platform as the positive direction, the positive direction of the z-axis points to the I quadrant of the spacecraft (pointing to the center of the earth when there is no attitude deviation), perpendicular to the longitudinal axis, and the y-axis, x-axis and z-axis form a right-hand system;

[0036] For the whole satellite platform coordinate system, the origin Oc is the center of mass of the whole satellite, and the XcYcZc axis is parallel to the satellite platform coordinate system;

[0037] For the payload platform coordinate system, the origin Oi is the connection point between the payload platform and the satellite platform, and the XiYiZi axis is parallel to the satellite platform coordinate system.

[0038] First, for step 100, the first motion parameter includes the angular velocity ω of the payload platform relative to the satellite platform i and the moment of inertia of the load platform J ic ;

[0039] The second motion parameter includes the rotation angular velocity ω of the satellite platform b ;

[0040] The position vector includes the first position vector r from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system ci , the second position vector r from the center of mass of the satellite to the center of mass of the payload platform oci .

[0041] Specifically, the angular velocity ω of the payload platform relative to the satellite platform i =[ω x 00] T ,ω b is the angular velocity of the satellite platform, which is equal to the platform orbital angular velocity ω under nominal conditions orb =[0ω00] T Consistent. ci =[r cix r ciy r ciz ] T is the first position vector from the magnetic suspension connection center to the mass center of the load platform; r oci =r oi +r ci -r c =[r x r y r z ] T is the second position vector from the center of mass of the whole satellite to the center of mass of the payload platform, wherein, in the satellite platform coordinate system, the third position vector r from the origin of the satellite platform coordinate system to the center of mass of the satellite is obtained. c , the fourth position vector r to the connection point oi . Load platform moment of inertia Defining the matrix inertia

[0042] In step 102, the position disturbance force of the load platform is the centrifugal disturbance force on the magnetic bearing joint when the center of mass of the load platform is not on the rotation axis during the rotation process;

[0043] The centrifugal interference force is determined by the following formula:

[0044] F i =m i ω i ×(ω i × ci )

[0045] Among them, F i is the position disturbance force of the load platform; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform; r ci It is the first position vector from the magnetic levitation connection point to the center of mass of the load platform in the load platform coordinate system.

[0046] In the present invention, the position interference analysis of the load platform is estimated on-orbit during the high-speed rotation of the load platform. When the center of mass of the load platform is not on the line connecting the rotation axes, the centrifugal interference force exerted on the magnetic bearing joints during the rotation of the rotating load is considered to be the position interference force of the load platform. The interference force appears as a fixed direction in the load platform coordinate system and does not change direction as the load platform rotates, thus meeting the feedforward control conditions.

[0047] In a preferred embodiment, the attitude disturbance torque of the load platform is introduced by the uneven mass distribution of the load platform during the rotation process;

[0048] The attitude disturbance torque of the load platform is determined by the following formula:

[0049]

[0050] Among them, M i is the attitude disturbance torque of the load platform; m i is the mass of the load platform; r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform; J sic is the matrix inertia defined based on the moment of inertia of the load platform; ω b is the angular velocity of the satellite platform; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, and θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis.

[0051] It should be noted that A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system. When small angle deviations are ignored, the rotation angle of the payload platform relative to the satellite platform around the rotation axis can be considered to be θ.

[0052] In the present invention, small items are ignored. During the high-speed rotation of the load platform, the load platform is affected by the interference torque mainly including the above-mentioned items, among which the attitude interference torque of the load platform is described in the load platform coordinate system as a constant item and a one-time frequency item, which meets the feedforward control conditions.

[0053] In step 102, the first disturbance torque is determined by the following formula:

[0054]

[0055] Among them, M b1 is the first disturbance torque; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform, ω i =[ω x 00] T ; r oci is the second position vector from the satellite's center of mass to the payload platform's center of mass, r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis.

[0056] In the present invention, during the on-orbit estimation of the high-speed rotation of the payload platform, if the center of mass of the payload platform is not on the line connecting the rotation axes, the interference generated on the satellite platform during the rotation of the rotating payload is described as shown by the first interference torque. At this time, the interference received by the satellite platform is expressed as a single-frequency term in the satellite platform coordinate system, and this interference can be feedforward compensated based on the on-orbit identification.

[0057] In a preferred embodiment, the second interference torque is determined by the following formula:

[0058]

[0059] Among them, M b2 is the second disturbance torque; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform.

[0060] In the present invention, the influence of the uneven distribution of mass characteristics of the large inertia load platform in the direction of the rotation axis is taken into account, and the interference generated by the rotating load on the satellite platform during the rotation is described as shown in the second interference torque. At this time, the interference to the satellite platform is expressed as a single-frequency term in the satellite platform coordinate system, and this interference can be feedforward compensated based on on-orbit identification.

[0061] In a preferred embodiment, the third interference torque is determined by the following formula:

[0062]

[0063] Among them, M b3 is the third disturbance torque; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J sic is the matrix inertia defined based on the moment of inertia of the load platform; ω b is the angular velocity of the satellite platform.

[0064] In the present invention, it can be seen from the third disturbance torque introduced by the uneven mass distribution of the load platform in the non-rotation axis direction that the disturbance term in the yaw direction of the satellite platform introduced by the rotation of the load platform around the rotation axis includes a constant value disturbance, and the disturbance torque is described as -J xx ω0ω x ; In addition to the constant interference in the yaw direction, it also includes the double frequency interference caused by the inertia asymmetry of the y / z axis (i.e., the non-rotating axis), and the interference amplitude is Therefore, feedforward compensation can be performed on the basis of on-orbit identification.

[0065] With respect to step 104, the satellite is jointly controlled according to the position disturbance force, attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform, including:

[0066] For each eddy current sampling cycle of the magnetic bearing joint, the following are executed:

[0067] The position disturbance force of the load platform is used as the first feedforward parameter;

[0068] Obtaining the current relative displacement deviation of the satellite platform in the eddy current sampling cycle and the last relative displacement deviation of the satellite platform in the last eddy current sampling cycle;

[0069] Jointly controlling the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the last relative displacement deviation and the first feedforward parameter;

[0070] The disturbance torque of the satellite platform is used as the second feedforward parameter;

[0071] Obtaining current posture information in the eddy current sampling period, and using the current posture information as a feedback parameter;

[0072] Implementing joint control of the attitude of the satellite platform according to the second feedforward parameter and the feedback parameter;

[0073] The attitude disturbance torque of the load platform is used as the third feedforward parameter;

[0074] The joint control of the attitude of the attitude platform is achieved according to the third feedforward parameter, the feedback parameter and the last attitude information of the load platform in the last eddy current sampling cycle.

[0075] In a specific embodiment, the relative displacement deviation in the eddy current sampling period is determined by the following method:

[0076] S1: Obtain the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform in the current eddy current sampling period;

[0077] S2: determining an eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship; the eddy current measurement vector includes a radial eddy current measurement vector and an axial eddy current measurement vector;

[0078] S3: For each dimension of the payload platform, the following steps are performed: according to the installation parameters and eddy current measurement vectors of each eddy current displacement sensor in the dimension, the relative attitude deviation and relative displacement deviation of the satellite platform are calculated; wherein, four eddy current displacement sensors are arranged in each dimension.

[0079] For step S1, the magnetic suspension bearing joint includes an eddy current displacement sensor, and the measured end of the eddy current displacement sensor is installed on the stator shaft of the satellite platform; the measuring end of the eddy current displacement sensor is installed on the rotor shaft of the load platform; the magnetic suspension joint includes a stator shaft and a rotor shaft;

[0080] The installation parameters include the first position vector r from the center of mass of the satellite platform to the measured end in the satellite platform coordinate system L_P , the second position vector r from the center of mass of the load platform to the measuring end in the load platform coordinate system U_Q +r I , the third position vector r from the center of mass of the satellite platform to the center of mass of the payload platform in the orbital system L_U , eddy current stator direction r in satellite platform coordinate system L_X , eddy current measurement rotor direction r in the load platform coordinate system U_X ;

[0081] The relative motion relationship includes: the transformation matrix C of the satellite platform coordinate system relative to the orbital system LO , the transformation matrix C of the load platform coordinate system relative to the track system UO .

[0082] It should be noted that the rotor of the magnetic bearing is installed on the rotor of the magnetic joint, and the rotor of the magnetic joint (i.e. the rotor shaft) is installed on the load platform; the stator of the magnetic bearing is installed on the stator of the joint (i.e. the stator shaft), and the stator of the joint is installed on the satellite platform. It should be noted that the working principle of the eddy current displacement sensor is to drive the magnetic field around the eddy current coil to change through a pulse current of a certain waveform, and generate eddy current at the induction point of the metal parts of the central bearing. Figure 3 In the schematic diagram of radial eddy current measurement working principle shown in FIG. 1 , the installation position of the eddy current displacement sensor is r U_X The direction of the rotor shaft parallel to the load platform, and the measurement direction is the load platform rotor shaft axis r U_X The vertical direction of the eddy current stator is related to the vector r L_X The direction is consistent. The vector direction is fixed in the satellite platform coordinate system. The direction is consistent and the vector size is ar L_X ; where a is the coefficient. The direction of the eddy current measurement rotor is related to r U_X The direction is consistent and is the fixed direction of this system.

[0083] In step S2, the radial eddy current measurement vector is determined by the following formula:

[0084]

[0085] Among them, r b is the radial eddy current measurement vector in the load platform coordinate system;

[0086] C UO is the transformation matrix of the load platform coordinate system relative to the track system, is the transposed matrix of the transformation matrix;

[0087] C LO is the transformation matrix of the satellite platform coordinate system relative to the orbital system, is the transposed matrix of the transformation matrix;

[0088] In the orbital system, r L_U is the third position vector from the center of mass of the satellite platform to the center of mass of the payload platform;

[0089] In the satellite platform coordinate system, r L_X is the direction of the stator axis; r L_P is the first position vector from the center of mass of the satellite platform to the measured end;

[0090] In the load platform coordinate system, r U_X is the direction of the rotor axis; r U_Q +r I is the second position vector from the center of mass of the load platform to the measuring end.

[0091] Specifically, Figure 3 As shown, the radial eddy current measurement vector r is uniformly expressed in the orbital system as follows:

[0092]

[0093] Because in Figure 3 In the same coordinate system, the radial eddy current measurement vector r and the eddy current installation direction r at a fixed position in the load platform coordinate system are U_X are mutually perpendicular vectors, so Considering the attitude motion of the satellite platform and the payload platform and the relative motion relationship between the platforms, the following expression for the coefficient a can be obtained:

[0094]

[0095] Substitute a into the above radial eddy current measurement vector r to obtain the radial eddy current measurement vector r in the load platform coordinate system. b :

[0096]

[0097] Similarly, the axial eddy current measurement vector is determined by the following formula:

[0098]

[0099] Among them, r s is the axial eddy current measurement vector in the load platform coordinate system, and b is the coefficient.

[0100] In the present invention, since the aerospace ultra-wide coverage high-resolution optical satellite has a large inertia payload, the relative attitude and relative position of the large inertia payload platform to be introduced need to be considered to ensure high measurement accuracy of the satellite.

[0101] For step S3, on the measurement section of each dimension, the eddy current displacement sensors are distributed on the upper eddy current surface and the lower eddy current surface, and two eddy current displacement sensors are installed symmetrically on the upper eddy current surface and the lower eddy current surface;

[0102] According to the installation parameters of each eddy current displacement sensor in this dimension and the eddy current measurement vector, the relative attitude deviation and relative displacement deviation of the satellite platform are calculated, including:

[0103] The projection of the eddy current measurement vector within the measurement cross section of this dimension is taken as the relative distance;

[0104] According to the installation parameters, the displacement of each eddy current displacement sensor from the geometric center of the rotating shaft of the magnetic suspension joint in the positive direction is determined;

[0105] According to the relative distance and displacement of each eddy current displacement sensor, the relative attitude deviation and relative displacement deviation are calculated.

[0106] It should be noted that the magnetic levitation joint has three translational degrees of freedom, each of which corresponds to each dimension of the load platform, and four eddy current displacement sensors are installed in each dimension.

[0107] In a specific embodiment, the relative posture deviation is determined by the following formula:

[0108]

[0109] The relative displacement deviation is determined by the following formula:

[0110]

[0111] Among them, θ is the relative posture deviation; r is the relative displacement deviation; d U_1 ,d U_2 are the relative distances between the first eddy current displacement sensor and the second eddy current displacement sensor on the upper eddy current surface within the measurement section; d D_1 ,d D_2 are the relative distances between the third eddy current displacement sensor and the fourth eddy current displacement sensor of the lower eddy current surface within the measuring section; L1 is the displacement of the eddy current displacement sensor located on the upper eddy current surface from the geometric center of the rotating shaft of the magnetic levitation joint in the positive direction; L2 is the displacement of the eddy current displacement sensor located on the lower eddy current surface from the geometric center of the rotating shaft of the magnetic levitation joint in the positive direction; L is the displacement of the upper eddy current surface and the lower eddy current surface in the positive direction, L=L1-L2.

[0112] It should be noted that d U_1 ,d U_2 ,d D_1 ,d D_2 All are scalars.

[0113] Specifically, the relative position and relative attitude of the eddy current measurement vectors measured by the four eddy current displacement sensors in each dimension are decoupled and calculated, such as Figure 4The eddy current data measurement situation within the radial measurement section shown in the figure, four eddies are installed as two upper and lower eddy current surfaces in the load platform coordinate system, and two eddy current displacement sensors are symmetrically installed at the center of each eddy current surface. The central bearing metal component fixed to the satellite platform in the nominal locked state is located at the center position of the eddy current measurement. After the lock is released, the load platform and the satellite platform produce relative position and posture changes. Among them, the first eddy current displacement sensor D1 and the second eddy current displacement sensor D2 on the upper eddy current surface, the third eddy current displacement sensor D3 and the fourth eddy current displacement sensor D4 on the lower eddy current surface. Ignoring small-order terms, the relative distance information measured by the four eddy current displacement sensors can decouple the calculation of the relative posture deviation θ and relative position deviation r of the satellite platform and the load platform. As shown Figure 4 As shown, the displacement of the eddy current displacement sensor located on the lower eddy current surface from the geometric center of the rotating shaft of the magnetic suspension joint in the positive direction is -L2; L=L1-(-L2).

[0114] In a preferred embodiment, the joint control of the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the last relative displacement deviation and the feedforward parameter includes:

[0115] The relative position control force is calculated according to the current relative displacement deviation, the previous relative displacement deviation and the first feedforward parameter;

[0116] Using relative position control force to achieve joint control of relative position;

[0117] Among them, the relative position control force is determined by the following formula:

[0118] F=-k i ∫rdt-k p rk d (r-r_lst) / dt-F i

[0119] Where F is the relative position control force; k i , k p , k d are the integral, proportional and differential coefficients of the PID controller respectively; F i is the position disturbance force of the load platform; r is the current relative displacement deviation; r_lst is the last relative displacement deviation.

[0120] It should be noted that k i , k p , k d They are all determined by engineering experience and are not specifically limited here.

[0121] In the present invention, based on PID feedback control combined with the position disturbance force F of the load platform as a feedforward parameter i, realizing the joint control of the relative positions of the payload platform and the satellite platform, which significantly improves the control accuracy and control stability compared with single feedback control.

[0122] In a preferred embodiment, the joint control of the attitude of the satellite platform is implemented according to the second feedforward parameter and the feedback parameter, including:

[0123] The control torque is calculated according to the second feedforward parameter and the feedback parameter;

[0124] Use control torque to achieve joint control of the satellite platform's attitude;

[0125] The control torque is determined by the following formula:

[0126] M b =-k ib ∫αdt-k pb α-k db α′-M b1 -M b2 -M b3

[0127] Among them, M b is the control torque; k ib , k pb , k db are the integral, proportional and differential coefficients of the PID controller respectively; α is the attitude information of the satellite platform to be controlled in the feedback parameter; M b1 、M b2 、M b3 They are respectively the first disturbance torque, the second disturbance torque and the third disturbance torque in the second feedforward parameter.

[0128] It should be noted that k ib , k pb , k db All are determined by engineering experience and are not specifically limited here. The attitude information to be controlled includes the angular velocity of the rotation axis; the pitch angle and yaw angle of the satellite platform of the non-rotation axis. α′=α / dt.

[0129] In the present invention, based on PID feedback control combined with the first disturbance torque, second disturbance torque and third disturbance torque of the satellite platform as feedforward parameters, joint control of the attitude of the satellite platform is realized. Compared with single feedback control, the control accuracy and control stability are significantly improved.

[0130] In the present invention, the on-orbit motion process of the satellite is controlled in real time through feedforward control and feedback control, thereby improving the attitude control accuracy and stability index, and ensuring the measurement accuracy and stability of the star sensor in practical applications.

[0131] In a preferred embodiment, the joint control of the attitude of the attitude platform is realized according to the third feedforward parameter, the feedback parameter and the last attitude information of the load platform in the last eddy current sampling cycle, including:

[0132] The control torque on the rotating axis and the non-rotating axis is calculated according to the third feedforward parameter, the feedback parameter and the last posture information;

[0133] The control torque is determined by the following formula:

[0134]

[0135] Among them, M_X, M_Y, M_Z are the control torques on the rotating axis X axis, the non-rotating axis Y axis, and the non-rotating axis Z axis respectively; ω_X, ω_Y, ω_Z are the components of the angular velocity of the load platform in the rolling direction on the X axis, Y axis, and Z axis respectively; ω_X_lst is the component of the angular velocity of the load platform in the rolling direction in the last attitude information on the X axis; A_Y, A_Z are the attitude deviations on the non-rotating axis Y axis and the non-rotating axis Z axis respectively; M iy 、M iz They are the attitude disturbance torque M of the load platform respectively. i Components on the non-rotating axes Y and Z; k ix , k px , k dx are the integral, proportional and differential coefficients of the PID controller on the rotating axis X-axis respectively; k iy , k py , k dy are the integral, proportional and differential coefficients of the PID controller on the Y axis of rotation respectively; k iz , k pz , k dz They are the integral, proportional and differential coefficients of the PID controller on the Z-axis of rotation.

[0136] It should be noted that the integral, proportional and differential coefficients of the above PID controller are determined by engineering experience and are not specifically limited here.

[0137] In a specific embodiment, the X axis is used as the rotation axis, and the Y and Z axes are used as non-rotation axes. Figure 5 The schematic diagram of the attitude interference of the satellite platform is shown, where mb2i = M b1 +M b2 +M b3 ; mb2i=[mb2ix mb2iy mb2iz] T ;Depend on Figure 5 It can be seen that the X-axis is less affected by the interference, while the Y and Z axes are more obviously affected by the interference, so it is necessary to control the interference of the non-rotating axis. Figure 6 A schematic diagram showing the control effect of the pitch attitude with feedback control before 500 seconds and joint control of feedforward and feedback after 500 seconds is shown; Figure 7 The figure shows the control effect of the pitch attitude angular velocity by the combined control of feedforward and feedback before 500s and after 500s. Figure 6 and Figure 7 It can be seen that the pitch angle and pitch angular velocity controlled only by feedback have poor stability, while the pitch angle and pitch angular velocity controlled by feedforward and feedback are obviously more stable and have higher control accuracy.

[0138] like Figure 8 , Fig. 9 As shown, an embodiment of the present invention provides a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, Figure 8 As shown in FIG. 1 , a hardware architecture diagram of a computing device where a satellite on-orbit identification and feedforward feedback joint control device based on a magnetic suspension bearing provided in an embodiment of the present invention is located is shown in FIG. Figure 8 In addition to the processor, memory, network interface, and non-volatile memory shown in the figure, the computing device in which the device is located in the embodiment may also generally include other hardware, such as a forwarding chip responsible for processing messages, etc. Taking software implementation as an example, Fig. 9 As shown, as a device in a logical sense, the CPU of the computing device in which it is located reads the corresponding computer program in the non-volatile memory into the memory and runs it. This embodiment provides a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings, which is applied to satellites in which both the satellite platform and the payload platform are large inertia rigid bodies, including:

[0139] An acquisition module 900 is used to acquire a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform;

[0140] An on-orbit identification module 902 is used to determine the position interference force, attitude interference torque and interference torque of the payload platform caused by the rotation process, and the satellite platform according to the first motion parameter, the second motion parameter, the position vector and the mass of the payload platform; wherein the interference torque includes a first interference torque introduced by the misalignment between the center of mass of the payload platform and the rotation axis, a second interference torque introduced by the uneven mass distribution of the payload platform in the direction of the rotation axis, and a third interference torque introduced by the uneven mass distribution of the payload platform in the direction of the non-rotation axis;

[0141] The control module 904 is used to jointly control the satellite according to the position disturbance force and attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

[0142] In some specific implementations, the acquisition module 900 may be used to execute the above step 100 , the on-track identification module 902 may be used to execute the above step 102 , and the control module 904 may be used to execute the above step 104 .

[0143] In some specific embodiments, the first motion parameter includes the angular velocity of the payload platform relative to the satellite platform and the moment of inertia of the payload platform;

[0144] The second motion parameter includes the rotational angular velocity of the satellite platform;

[0145] The position vectors include a first position vector from the magnetic levitation connection point to the mass center of the load platform in the load platform coordinate system, and a second position vector from the mass center of the satellite to the mass center of the load platform.

[0146] In some specific implementations, in the on-track identification module 902, the position disturbance force of the load platform is the centrifugal disturbance force on the magnetic bearing joint when the center of mass of the load platform is not on the rotation axis during the rotation process;

[0147] The centrifugal interference force is determined by the following formula:

[0148] F i =m i ω i ×(ω i × ci )

[0149] Among them, F i is the position disturbance force of the load platform; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform; r ci It is the first position vector from the magnetic levitation connection point to the center of mass of the load platform in the load platform coordinate system.

[0150] In some specific implementations, in the on-orbit identification module 902, the attitude disturbance torque of the payload platform is introduced by the uneven mass distribution of the payload platform during the rotation process;

[0151] The attitude disturbance torque of the load platform is determined by the following formula:

[0152] M i =m i r ci ×(ω i ×(ω i × ci ))+ωi ×J ic ω i +2ω i J sic (A(θ)ω b )

[0153] Among them, M i is the attitude disturbance torque of the load platform; m i is the mass of the load platform; r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform; J sic is the matrix inertia defined based on the moment of inertia of the load platform; ω b is the angular velocity of the satellite platform; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, and θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis.

[0154] In some specific implementations, in the on-track identification module 902, the first interference torque is determined by the following formula:

[0155]

[0156] Among them, M b1 is the first disturbance torque; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform, ω i =[ω x 00] T ; r oci is the second position vector from the satellite's center of mass to the payload platform's center of mass, r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis.

[0157] In some specific implementations, in the on-track identification module 902, the second interference torque is determined by the following formula:

[0158] M b2 =-A(θ) T ω i ×J ic ω i

[0159] Among them, M b2is the second disturbance torque; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform.

[0160] In some specific implementations, in the on-track identification module 902, the third interference torque is determined by the following formula:

[0161] M b3 =-2A(θ) T ω i J sic (A(θ)ω b )

[0162] Among them, M b3 is the third disturbance torque; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J sic is the matrix inertia defined based on the moment of inertia of the load platform; ω b is the angular velocity of the satellite platform.

[0163] In some specific implementations, the control module 904 is further configured to perform the following operations:

[0164] For each eddy current sampling cycle of the magnetic bearing joint, the following are executed:

[0165] A1: The position disturbance force of the load platform is used as the first feedforward parameter;

[0166] A2: Obtain the current relative displacement deviation of the satellite platform in the eddy current sampling cycle and the last relative displacement deviation of the satellite platform in the last eddy current sampling cycle;

[0167] A3: Joint control of the relative position of the payload platform and the satellite platform based on the current relative displacement deviation, the previous relative displacement deviation and the first feedforward parameter;

[0168] A4: The disturbance torque of the satellite platform is used as the second feedforward parameter;

[0169] A5: Obtain the current posture information in the eddy current sampling period, and use the current posture information as a feedback parameter;

[0170] A6: realizing joint control of the attitude of the satellite platform according to the second feedforward parameter and the feedback parameter;

[0171] A7: The attitude disturbance torque of the payload platform is used as the third feedforward parameter;

[0172] A8: A joint control of the attitude of the attitude platform is achieved according to the third feedforward parameter, the feedback parameter and the last attitude information of the load platform in the last eddy current sampling cycle.

[0173] In some specific embodiments, in step A3,

[0174] The relative position control force is calculated according to the current relative displacement deviation, the previous relative displacement deviation and the first feedforward parameter;

[0175] Using relative position control force to achieve joint control of relative position;

[0176] Among them, the relative position control force is determined by the following formula:

[0177] F=-k i ∫rdt-k p rk d (r-r_lst) / dt-F i

[0178] Where F is the relative position control force; k i , k p , k d are the integral, proportional and differential coefficients of the PID controller respectively; F i is the position disturbance force of the load platform; r is the current relative displacement deviation; r_lst is the last relative displacement deviation.

[0179] In some specific embodiments, in step A6,

[0180] The control torque is calculated according to the second feedforward parameter and the feedback parameter;

[0181] Use control torque to achieve joint control of the satellite platform's attitude;

[0182] The control torque is determined by the following formula:

[0183] M b =-k ib ∫αdt-k pb α-k db α′-M b1 -M b2 -M b3

[0184] Among them, M b is the control torque; k ib , k pb , k dbare the integral, proportional and differential coefficients of the PID controller respectively; α is the posture information to be controlled in the feedback parameter; M b1 、M b2 、M b3 They are respectively the first disturbance torque, the second disturbance torque and the third disturbance torque in the second feedforward parameter.

[0185] In some specific implementations, in step A8, the control torque on the rotating axis and the non-rotating axis is calculated based on the third feedforward parameter, the feedback parameter and the last posture information;

[0186] The control torque is determined by the following formula:

[0187]

[0188] Among them, M_X, M_Y, M_Z are the control torques on the rotating axis X axis, the non-rotating axis Y axis, and the non-rotating axis Z axis respectively; ω_X, ω_Y, ω_Z are the components of the angular velocity of the load platform in the rolling direction on the X axis, Y axis, and Z axis respectively; ω_X_lst is the component of the angular velocity of the load platform in the rolling direction in the last attitude information on the X axis; A_Y, A_Z are the attitude deviations on the non-rotating axis Y axis and the non-rotating axis Z axis respectively; M iy 、M iz They are the attitude disturbance torque M of the load platform respectively. i Components on the non-rotating axes Y and Z; k ix , k px , k dx are the integral, proportional and differential coefficients of the PID controller on the rotating axis X-axis respectively; k iy , k py , k dy are the integral, proportional and differential coefficients of the PID controller on the Y axis of rotation respectively; k iz , k pz , k dz They are the integral, proportional and differential coefficients of the PID controller on the Z-axis of rotation.

[0189] It is understandable that the structure illustrated in the embodiment of the present invention does not constitute a specific limitation on a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings. In other embodiments of the present invention, a satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings may include more or fewer components than shown in the figure, or combine some components, or split some components, or arrange the components differently. The components shown in the figure may be implemented in hardware, software, or a combination of software and hardware.

[0190] The information interaction, execution process and other contents between the modules in the above-mentioned device are based on the same concept as the embodiment of the method of the present invention. For specific contents, please refer to the description in the embodiment of the method of the present invention, and no further description is given here.

[0191] An embodiment of the present invention further provides a computing device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, a satellite in-orbit identification and feedforward feedback joint control method based on magnetic levitation bearings in any embodiment of the present invention is implemented.

[0192] An embodiment of the present invention also provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the processor executes a satellite in-orbit identification and feedforward feedback joint control method based on magnetic levitation bearings in any embodiment of the present invention.

[0193] An embodiment of the present application also provides a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium, and the processor executes the computer program, so that the computer device executes a satellite in-orbit identification and feedforward feedback joint control method based on magnetic levitation bearings as described in any of the above embodiments.

[0194] Specifically, a system or device equipped with a storage medium can be provided, on which software program code that implements the functions of any of the above-mentioned embodiments is stored, and a computer (or CPU or MPU) of the system or device can be enabled to read and execute the program code stored in the storage medium.

[0195] In this case, the program code itself read from the storage medium can realize the function of any one of the above-mentioned embodiments, and thus the program code and the storage medium storing the program code constitute a part of the present invention.

[0196] The storage medium embodiments for providing the program code include a floppy disk, a hard disk, a magneto-optical disk, an optical disk (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), a magnetic tape, a non-volatile memory card, and a ROM. Alternatively, the program code can be downloaded from a server computer by a communication network.

[0197] In addition, it should be clear that the functions of any of the above embodiments can be implemented not only by executing the program code read by the computer, but also by enabling an operating system operating on the computer to complete part or all of the actual operations based on instructions from the program code.

[0198] In addition, it can be understood that the program code read from the storage medium is written to a memory provided in an expansion board inserted into the computer or to a memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, a CPU installed on the expansion board or expansion module is enabled to perform part or all of the actual operations, thereby realizing the functions of any of the above-mentioned embodiments.

[0199] It should be noted that, in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the statement "comprise a ..." do not exclude the presence of other identical factors in the process, method, article or device including the elements.

[0200] A person of ordinary skill in the art can understand that all or part of the steps of implementing the above method embodiments can be completed by hardware related to program instructions, and the aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it executes the steps of the above method embodiments; and the aforementioned storage medium includes: ROM, RAM, magnetic disk or optical disk, etc., various media that can store program codes.

[0201] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A satellite on-orbit identification and feedforward feedback joint control method based on magnetic suspension bearings, characterized in that: Applied to satellites where both the satellite platform and the payload platform are large inertia rigid bodies, including: Acquire a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform; According to the first motion parameter, the second motion parameter, the position vector and the mass of the load platform, determine the position interference force, attitude interference torque and interference torque on the load platform caused by the rotation process, wherein the interference torque includes a first interference torque introduced by the misalignment of the center of mass of the load platform and the rotation axis, a second interference torque introduced by the uneven mass distribution of the load platform in the direction of the rotation axis, and a third interference torque introduced by the uneven mass distribution of the load platform in the direction of the non-rotation axis; The satellite is jointly controlled according to the position disturbance force, attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

2. The method according to claim 1, characterized in that The first motion parameter includes the angular velocity of the payload platform relative to the satellite platform and the moment of inertia of the payload platform; The second motion parameter includes the rotational angular velocity of the satellite platform; The position vector includes a first position vector from the magnetic levitation connection point to the center of mass of the payload platform in the payload platform coordinate system, and a second position vector from the center of mass of the satellite to the center of mass of the payload platform.

3. The method according to claim 2, characterized in that The position disturbance force of the load platform is the centrifugal disturbance force on the magnetic bearing joint when the center of mass of the load platform is not on the rotation axis during the rotation process; The centrifugal interference force is determined by the following formula: F i =m i oh i ×(ω i ×r ci ) Among them, F i is the position disturbance force of the load platform; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform; r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system; and / or, The attitude disturbance torque of the load platform is introduced by the uneven mass distribution of the load platform during the rotation process; The attitude disturbance torque of the load platform is determined by the following formula: M i =m i r ci ×(ω i ×(ω i ×r ci ))+ω i ×J ic oh i +2h i J sic (A(θ)ω b ) Among them, M i is the attitude disturbance torque of the load platform; m i is the mass of the load platform; r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform; J sic is the matrix inertia defined based on the rotational inertia of the load platform; ω b is the angular velocity of rotation of the satellite platform; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, and θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis.

4. The method according to claim 2, characterized in that: The first interference torque is determined by the following formula: Among them, M b1 is the first disturbance torque; m i is the mass of the load platform; ω i is the angular velocity of the payload platform relative to the satellite platform, ω i =[ω x 00] T ; r oci is the second position vector from the center of mass of the satellite to the center of mass of the payload platform, r oci =[r x r y r z ] T ; r ci is the first position vector from the magnetic suspension connection point to the center of mass of the load platform in the load platform coordinate system, r ci =[r cix r ciy r ciz ] T ; θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; and / or, The second interference torque is determined by the following formula: M b2 =-A(θ) T oh i ×J ic oh i Among them, M b2 is the second interference torque; A(θ) is the transformation matrix from the satellite platform coordinate system to the payload platform coordinate system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J ic is the moment of inertia of the load platform.

5. The method according to claim 2, characterized in that: The third interference torque is determined by the following formula: M b3 =-2A(θ) T ω i J sic (A(θ)ω b ) Among them, M b3 is the third disturbance torque; A(θ) is the direction from the satellite platform coordinate system to the load platform coordinate system The conversion matrix of the reference system, θ is the rotation angle of the payload platform relative to the satellite platform around the rotation axis; ω i is the angular velocity of the payload platform relative to the satellite platform; J sic is the matrix inertia defined based on the rotational inertia of the load platform; ω b is the rotation angular velocity of the satellite platform.

6. The method according to any one of claims 2 to 5, characterized in that: The jointly controlling the satellite according to the position disturbance force, attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform comprises: For each eddy current sampling cycle of the magnetic bearing joint, the following steps are performed: Using the position disturbance force of the load platform as a first feedforward parameter; Obtaining the current relative displacement deviation of the satellite platform in the eddy current sampling cycle and the last relative displacement deviation of the satellite platform in the last eddy current sampling cycle; Jointly controlling the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the last relative displacement deviation and the first feedforward parameter; Using the disturbance torque of the satellite platform as a second feedforward parameter; Acquire current posture information in the eddy current sampling period, and use the current posture information as a feedback parameter; Implementing joint control of the attitude of the satellite platform according to the second feedforward parameter and the feedback parameter; Using the attitude disturbance torque of the load platform as the third feedforward parameter; The joint control of the attitude of the attitude platform is achieved according to the third feedforward parameter, the feedback parameter and the last attitude information of the load platform in the last eddy current sampling cycle.

7. The method according to claim 6, characterized in that The joint control of the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the last relative displacement deviation and the feedforward parameter includes: Calculating a relative position control force according to the current relative displacement deviation, the previous relative displacement deviation and the first feedforward parameter; Using the relative position control force to achieve joint control of the relative position; Wherein, the relative position control force is determined by the following formula: F=-k i ∫rdt-k p r-k d (r-r_lst) / dt-F i Wherein, F is the relative position control force; k i , k p , k d are the integral, proportional and differential coefficients of the PID controller respectively; F i is the position disturbance force of the load platform; r is the current relative displacement deviation; r_lst is the last relative displacement deviation; and / or, The step of realizing joint control of the attitude of the satellite platform according to the second feedforward parameter and the feedback parameter includes: Calculating a control torque according to the second feedforward parameter and the feedback parameter; Using the control torque to achieve joint control of the attitude of the satellite platform; Wherein, the control torque is determined by the following formula: M b =-k ib ∫αdt-k pb a-k db α′-M b1 -M b2 -M b3 Among them, M b is the control torque; k ib , k pb , k db are the integral, proportional and differential coefficients of the PID controller respectively; α is the posture information to be controlled in the feedback parameter; M b1 、M b2 、M b3 They are respectively the first interference torque, the second interference torque, and the third interference torque in the second feedforward parameters.

8. A satellite on-orbit identification and feedforward feedback joint control device based on magnetic suspension bearings, characterized in that: Applied to satellites where both the satellite platform and the payload platform are large inertia rigid bodies, including: An acquisition module, used for acquiring a first motion parameter of the payload platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the payload platform in the satellite during the rotation of the payload platform; An on-orbit identification module is used to determine the position interference force, attitude interference torque and interference torque of the payload platform caused by the rotation process, and the satellite platform according to the first motion parameter, the second motion parameter, the position vector and the mass of the payload platform; wherein the interference torque includes a first interference torque introduced by the misalignment between the center of mass of the payload platform and the rotation axis, a second interference torque introduced by the uneven mass distribution of the payload platform in the direction of the rotation axis, and a third interference torque introduced by the uneven mass distribution of the payload platform in the direction of the non-rotation axis; A control module is used to jointly control the satellite according to the position disturbance force and attitude disturbance torque of the payload platform and the disturbance torque of the satellite platform.

9. A computing device, comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the method according to any one of claims 1 to 7 is implemented.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to execute the method according to any one of claims 1 to 7.

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