Satellite On-Orbit Identification and Combined Feedforward and Feedback Control Method Based on Magnetic Levitation Bearings

Abstract

The present invention provides a method for on-orbit identification and combined feedforward and feedback control of a satellite based on a magnetic levitation bearing, which is applied to a satellite in which both the satellite platform and the payload platform are large-inertia rigid bodies. The method includes: determining the position disturbing force, attitude disturbing moment on the payload platform, and disturbing moment on the satellite platform during the rotation process according to the first motion parameter of the payload platform, the second motion parameter of the satellite platform, the position vector, and the mass of the payload platform during the rotation process; the disturbing moment includes a first disturbing moment introduced by the non-coincidence of the center of mass and the rotation axis of the payload platform, a second disturbing moment introduced by the uneven mass distribution of the payload platform in the rotation axis direction, and a third disturbing moment introduced by the uneven mass distribution of the payload platform in the non-rotation axis direction; performing combined control on the satellite according to the position disturbing force, attitude disturbing moment of the payload platform, and disturbing moment of the satellite platform. This solution performs on-orbit identification of the disturbing terms during the rotation process, designs feedforward control for the controller, and improves the satellite attitude control accuracy and stability index.

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 particularly to a satellite on-orbit identification and feedforward-feedback combined control method based on a magnetic levitation bearing. Background Art

[0002] The magnetic levitation bearing joint completes the connection between the stable satellite platform and the high-speed rotating load platform system, and outputs control force and control torque by the magnetic levitation bearing to complete the stable control of the load platform and the satellite platform. With the requirement of high-speed rotation for the load platform in satellite imaging missions, the load platform adopted has a higher mass, and the characteristics of large mass and high dynamics of the load platform determine that the interference characteristics during the high-speed rotation process for the load platform and the satellite platform are obvious, and the motion characteristics of the satellite platform and the load platform determine that the manifestation forms of the interference characteristics received by the load platform and the satellite platform are different. In order to ensure the satellite attitude control accuracy and stability index, it is necessary to separately perform interference identification and feedforward controller design for the controllers of the load platform and the satellite platform. Therefore, there is an urgent need to provide a satellite on-orbit identification and feedforward-feedback combined control method based on a magnetic levitation bearing. Summary of the Invention

[0003] The present invention provides a satellite on-orbit identification and feedforward-feedback combined control method based on a magnetic levitation bearing, which decomposes the position and attitude interference of the load platform and the satellite platform during the high-speed rotation process of the load platform, and designs interference identification and feedforward control 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 combined control method based on a magnetic levitation bearing, which is applied to a satellite in which both the satellite platform and the load platform are large inertia rigid bodies. The method includes:

[0005] Obtaining the first motion parameter of the load platform, the second motion parameter of the satellite platform, and the position vectors of the satellite platform and the load platform in the satellite during the rotation process of the load platform;

[0006] Determining the position interference force, attitude interference torque of the rotation process on the load platform, and the interference torque on the satellite platform according to the first motion parameter, the second motion parameter, the position vectors, and the mass of the load platform; wherein, the interference torque includes a first interference torque introduced by the non-coincidence of the centroid and the rotation axis of the load platform, a second interference torque introduced by the uneven mass distribution of the load platform in the rotation axis direction, and a third interference torque introduced by the uneven mass distribution of the load platform in the non-rotation axis direction;

[0007] Performing combined control on the satellite according to the position interference force, attitude interference torque of the load platform, and the interference torque of the satellite platform.

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

[0009] An acquisition module, configured 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 during the rotation of the payload platform;

[0010] An on-orbit identification module, configured to determine a position disturbing force, an attitude disturbing torque on the payload platform, and a disturbing torque on the satellite platform during the rotation according to the first motion parameter, the second motion parameter, the position vector, and the mass of the payload platform; wherein, the disturbing torque includes a first disturbing torque introduced by the non-coincidence of the center of mass and the rotation axis of the payload platform, a second disturbing torque introduced by the uneven mass distribution of the payload platform in the rotation axis direction, and a third disturbing torque introduced by the uneven mass distribution of the payload platform in the non-rotation axis direction;

[0011] A control module, configured to perform combined control on the satellite according to the position disturbing force, the attitude disturbing torque of the payload platform, and the disturbing torque of the satellite platform.

[0012] In a third aspect, the present invention further provides a computing device, including a memory and a processor, where a computer program is stored in the memory, and when the processor executes the computer program, the steps of the satellite on-orbit identification and feedforward-feedback combined control method according to any one of the above are implemented.

[0013] In a fourth aspect, the present invention further provides a computer-readable storage medium, on which a computer program is stored, and when the computer program is executed in a computer, the computer is made to execute the satellite on-orbit identification and feedforward-feedback combined control method according to any one of the above.

[0014] In a fifth aspect, an embodiment of the present invention further provides a computer program product, including computer instructions, and when the computer instructions are executed by a processor, the steps of the method according to any first aspect of the present specification are implemented.

[0015] The present invention provides a satellite on-orbit identification and feedforward-feedback combined control method based on magnetic levitation bearings, which is applied to a satellite platform and a payload platform that are both large-inertia rigid bodies connected by magnetic levitation joints. By analyzing the motion parameters and position vectors of the satellite platform and the payload platform, this method sequentially determines the position interference force and attitude interference torque received by the payload platform during the rotation of the large-inertia payload platform, as well as the interference torques received by the satellite platform, which are introduced by the non-coincidence of the center of mass and the rotation axis of the payload platform, by the uneven mass distribution of the payload platform in the rotation axis direction, and by the uneven mass distribution of the payload platform in the non-rotation axis direction. Then, the feedforward and feedback combined control is used to control the satellite based on the position interference force, attitude interference torque of the payload platform, and the interference torque of the satellite platform. Thus, the present invention determines these interferences introduced by the large-inertia payload through on-orbit identification and uses the combined control method of feedforward combined with feedback to process these interferences, improving the control accuracy and ensuring the attitude accuracy and stability index during the on-orbit operation of the satellite. Brief Description of the Drawings

[0016] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0017] Figure 1 is a flowchart of a satellite on-orbit identification and feedforward-feedback combined control method based on magnetic levitation bearings provided by an 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 is a schematic diagram of the working principle of radial eddy current measurement provided by an embodiment of the present invention;

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

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

[0022] Figure 6 is a schematic diagram of the pitch attitude control effect of the feedforward control before 500s and the combined control after 500s provided by an embodiment of the present invention;

[0023] Figure 7It is a schematic diagram of the control effect of pitch attitude angular velocity with feedback control before 500s and combined control after 500s provided by an embodiment of the present invention;

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

[0025] Figure 9 It is a structural diagram of a satellite on-orbit identification and feedforward-feedback combined control device based on a magnetic levitation bearing provided by an embodiment of the present invention. Detailed implementation manners

[0026] To make the objectives, 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 with reference to the accompanying drawings in the embodiments of the present invention. Apparently, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

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

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

[0029] Step 102, determining the position interference force, attitude interference torque on the payload platform, and interference torque on the satellite platform during the rotation process according to the first motion parameter, the second motion parameter, the position vectors, and the mass of the payload platform; wherein, the interference torque includes a first interference torque introduced by the non-coincidence of the centroid and the rotation axis of the payload platform, a second interference torque introduced by the uneven mass distribution of the payload platform in the rotation axis direction, and a third interference torque introduced by the uneven mass distribution of the payload platform in the non-rotation axis direction;

[0030] Step 104, performing combined control on the satellite according to the position interference force of the payload platform, the attitude interference torque, and the interference torque of the satellite platform.

[0031] In the present invention, in a satellite platform and a payload platform that are both large-inertia rigid bodies connected by a magnetic levitation joint, by analyzing the motion parameters and position vectors of the satellite platform and the payload platform, the position disturbance force and attitude disturbance torque received by the payload platform during the rotation process of the large-inertia payload platform are determined in sequence, as well as the disturbance torques received by the satellite platform, which are introduced respectively by the non-coincidence of the center of mass and the rotation axis of the payload platform, by the uneven mass distribution of the payload platform in the rotation axis direction, and by the uneven mass distribution of the payload platform in the non-rotation axis direction. Then, the satellite is controlled by using a combined feedforward and feedback control based on the position disturbance force, attitude disturbance torque of the payload platform, and the disturbance torque of the satellite platform. Thus, the present invention determines these disturbances introduced by the large-inertia payload through on-orbit identification, and uses a combined control method of feedforward and feedback to process these disturbances, improving the control accuracy while ensuring the attitude accuracy and stability index during the on-orbit operation of the satellite.

[0032] It should be noted that the payload platform and the satellite platform are connected through a connection point (for example, 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 is comparable to the mass of the satellite platform.

[0033] The following describes Figure 1 the execution manner of each step shown.

[0034] The coordinate systems adopted 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. Specifically, as Figure 2 shown, for the geocentric inertial coordinate system, with the origin O E being the center of mass of the earth, O E the X E axis points to the mean equinox of 2000.0, O E the Z E axis is perpendicular to the mean equatorial plane of 2000.0 and is in the same direction as the earth's angular velocity of rotation, O E the Y E axis and the O E X E axis, O E Z E axis form a right-handed coordinate system;

[0035] For the satellite platform coordinate system, the origin O is the center of the connection end face between the satellite and the launch vehicle, the O x axis is along the longitudinal axis direction of the satellite body, with the direction pointing to the payload platform being 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) and is perpendicular to the longitudinal axis, and the y-axis and the x, z axes form a right-handed coordinate system;

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

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

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

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

[0040] The position vector includes the first position vector r from the magnetic levitation connection point to the centroid of the load platform in the load platform coordinate system ci and the second position vector r from the centroid of the satellite to the centroid of the load platform oci .

[0041] Specifically, the angular velocity ω of the load platform relative to the satellite platform i = [ω x 0 0] T , where ω b is the angular velocity of the satellite platform, which is consistent with the orbital angular velocity ω orb = [0 ω 0 0] T under nominal conditions. r ci = [r cix r ciy r ciz T is the first position vector from the magnetic levitation connection center to the centroid 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 centroid of the entire satellite to the centroid of the load platform. Among them, in the satellite platform coordinate system, the third position vector r from the origin of the satellite platform coordinate system to the centroid of the satellite c and the fourth position vector r from the origin to the connection point oi are obtained. The moment of inertia of the load platform defines the matrix inertia

[0042] In step 102, the position disturbing force of the load platform is the centrifugal disturbing force on the magnetic levitation bearing joint when the centroid of the load platform is not on the rotation axis during rotation;

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

[0044] F i = m i ω i × (ω i × r ci )

[0045] Wherein, F i is the position interference force of the payload platform; m i is the mass of the payload platform; ω i is the angular velocity of motion of the payload platform relative to the satellite platform; r ci is the first position vector from the magnetic levitation connection point in the payload platform coordinate system to the centroid of the payload platform.

[0046] In the present invention, for the analysis of the position interference of the payload platform during high-speed rotation in orbit, when the centroid of the payload platform is not on the rotation axis connection line, it is considered that the centrifugal interference force received by the magnetic bearing joint during the rotation of the rotating payload is the position interference force of the payload platform. This interference force shows a fixed direction in the payload platform coordinate system and does not change direction with the rotation of the payload platform, meeting the conditions for feedforward control.

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

[0048] The attitude interference torque of the payload platform is determined by the following formula:

[0049]

[0050] Wherein, M i is the attitude interference torque of the payload platform; m i is the mass of the payload platform; r ci is the first position vector from the magnetic levitation connection point in the payload platform coordinate system to the centroid of the payload platform; ω i is the angular velocity of motion of the payload platform relative to the satellite platform; J ic is the moment of inertia of the payload platform; J sic is the matrix inertia defined based on the moment of inertia of the payload 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.

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

[0052] In the present invention, when ignoring minor terms, during the high-speed rotation of the payload platform, the influence of the disturbance torque on the payload platform mainly includes the above items. Among them, the attitude disturbance torque of the payload platform is described in the payload platform coordinate system and expressed as a constant term and a first-order frequency term, satisfying the condition for feedforward control.

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

[0054]

[0055] Wherein, M b1 is the first disturbance torque; m i is the mass of the payload platform; ω i is the angular velocity of the relative motion of the payload platform with respect to the satellite platform, ω i = [ω x 0 0] T ; r oci is the second position vector from the centroid of the satellite to the centroid of the payload platform, r ci is the first position vector from the magnetic levitation connection point in the payload platform coordinate system to the centroid of the payload platform, θ is the rotation angle of the payload platform relative to the satellite platform about the rotation axis.

[0056] In the present invention, during the on-orbit estimation of the high-speed rotation of the payload platform, if the centroid of the payload platform is not on the rotation axis connection line, the disturbance generated by the rotating payload during the rotation process on the satellite platform is described as the first disturbance torque. At this time, the disturbance received by the satellite platform is expressed as a first-order frequency term in the satellite platform coordinate system, and feedforward compensation can be performed on this disturbance based on on-orbit identification.

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

[0058]

[0059] Wherein, 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 about the rotation axis; ω i is the angular velocity of the relative motion of the payload platform with respect to the satellite platform; J ic is the moment of inertia of the payload platform.

[0060] In the present invention, considering the influence of the uneven mass characteristic distribution of the large-inertia payload platform in the rotation axis direction, the disturbance generated by the rotating payload during the rotation process on the satellite platform is described as the second disturbance torque. At this time, the disturbance received by the satellite platform is expressed as a first-order frequency term in the satellite platform coordinate system, and feedforward compensation can be performed on this disturbance based on on-orbit identification.

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

[0062]

[0063] where M b3 is the third interference torque; 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 about 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 payload platform; ω b is the angular velocity of the satellite platform rotation.

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

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

[0066] For each eddy current sampling period of the magnetic levitation bearing joint, the following operations are performed:

[0067] Take the position interference force of the payload platform as the first feedforward parameter;

[0068] Obtain the current relative displacement deviation of the satellite platform in this eddy current sampling period and the previous relative displacement deviation of the satellite platform in the previous eddy current sampling period;

[0069] Perform joint control of the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the previous relative displacement deviation, and the first feedforward parameter;

[0070] Take the interference torque of the satellite platform as the second feedforward parameter;

[0071] Obtain the current attitude information in this eddy current sampling period and take the current attitude information as the feedback parameter;

[0072] Realize 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 payload platform is used as the third feedforward parameter;

[0074] Based on the third feedforward parameter, the feedback parameter, and the previous attitude information of the payload platform in the previous eddy current sampling period, the attitude of the attitude platform is jointly controlled.

[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 in the current eddy current sampling period and the relative motion relationship between the satellite platform and the payload platform;

[0077] S2: Determine the 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, perform: calculate the relative attitude deviation and relative displacement deviation of the satellite platform according to the installation parameters and the eddy current measurement vector of each eddy current displacement sensor in this dimension; among them, 4 eddy current displacement sensors are provided in each dimension.

[0079] Regarding step S1, the magnetic levitation bearing joint includes an eddy current displacement sensor, 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 payload platform; the magnetic levitation joint includes a stator shaft and a rotor shaft;

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

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

[0082] It should be noted that the rotor of the magnetic levitation bearing is installed on the rotor of the magnetic levitation joint, and the rotor of the magnetic levitation joint (i.e., the rotor shaft) is installed on the load platform; the stator of the magnetic levitation 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 that a pulsed current of a certain waveform drives the change of the magnetic field around the eddy current coil, and an eddy current is induced at the metal part of the central bearing. In the schematic diagram of the working principle of radial eddy current measurement as shown in Figure 3 , the installation position r U_X of the eddy current displacement sensor is parallel to the axis direction of the rotor shaft of the load platform, and the measurement direction is perpendicular to the axis r U_X of the rotor shaft of the load platform. The direction of the eddy current stator is consistent with the vector r L_X direction. In the satellite platform coordinate system, the vector direction is fixed. Take the direction as consistent, and the vector magnitude is ar L_X ; where a is a coefficient. The direction of the eddy current measurement rotor is consistent with the r U_X direction, which is a fixed direction in this system.

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

[0084]

[0085] where 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 of the load platform relative to the orbital system, is the transpose matrix of this transformation matrix;

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

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

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

[0090] In the load platform coordinate system, r U_X is the direction of the rotor shaft; r U_Q +r I is the second position vector from the centroid of the load platform to the measurement end.

[0091] Specifically, as Figure 3 shown, the vector description relationship of the radial eddy current measurement vector r represented uniformly in the orbital system is:

[0092]

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

[0094]

[0095] Substituting a into the above radial eddy current measurement vector r, the radial eddy current measurement vector r in the load platform coordinate system is obtained b :

[0096]

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

[0098]

[0099] where 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 load, it is necessary to consider the relative attitude and relative position introduced by the large inertia load platform to ensure the high measurement accuracy of the satellite.

[0101] For step S3, on the measurement section in each dimension, eddy current displacement sensors are distributed on the upper eddy current surface and the lower eddy current surface, and 2 eddy current displacement sensors are symmetrically installed at the center on both the upper eddy current surface and the lower eddy current surface;

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

[0103] Taking the projection of the eddy current measurement vector in the measurement section of this dimension as the relative distance;

[0104] According to the installation parameters, determine the displacement amount of each eddy current displacement sensor from the geometric center of the axis of the magnetic suspension joint in the positive direction;

[0105] Based on the relative distances and displacement amounts 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 translational degree of freedom corresponds to each dimension of the load platform, and 4 eddy current displacement sensors are installed in each dimension.

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

[0108]

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

[0110]

[0111] Where, θ is the relative attitude deviation; r is the relative displacement deviation; d U_1 、d U_2 are the relative distances of the first eddy current displacement sensor and the second eddy current displacement sensor on the upper eddy current surface in the measurement section respectively; d D_1 、d D_2 are the relative distances of the third eddy current displacement sensor and the fourth eddy current displacement sensor on the lower eddy current surface in the measurement section respectively; L1 is the displacement amount of the eddy current displacement sensor located on the upper eddy current surface in the positive direction from the geometric center of the axis of the magnetic levitation joint; L2 is the displacement amount of the eddy current displacement sensor located on the lower eddy current surface in the positive direction from the geometric center of the axis of the magnetic levitation joint; L is the displacement amount of the upper eddy current surface and the lower eddy current surface in the positive direction, and L = L1 - L2.

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

[0113] Specifically, decoupling calculations of the relative position and relative attitude are performed on the eddy current measurement vectors measured by the 4 eddy current displacement sensors in each dimension, such as Figure 4The measurement of eddy current data within the indicated radial measurement section. Four eddy currents are installed as two upper and lower eddy current surfaces in the load platform coordinate system. Two eddy current displacement sensors are symmetrically installed at the center of each eddy current surface. The central bearing metal component fixedly connected to the satellite platform under the nominal locked state is located at the center of the eddy current measurement. After unlocking, relative position and attitude changes occur between the load platform and the satellite platform. Among them, the first eddy current displacement sensor D1 and the second eddy current displacement sensor D2 on the upper eddy current surface, and 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 be decoupled to calculate the relative attitude deviation θ and the relative position deviation r between the satellite platform and the load platform. As Figure 4 shown, the displacement of the eddy current displacement sensor located on the lower eddy current surface in the positive direction from the geometric center of the axis of the magnetic levitation joint is -L2; L = L1 - (-L2).

[0114] In a preferred embodiment, the joint control of the relative position between the load platform and the satellite platform based on the current relative displacement deviation, the previous relative displacement deviation, and the feedforward parameter includes:

[0115] Calculating the relative position control force based on the current relative displacement deviation, the previous relative displacement deviation, and the first feedforward parameter;

[0116] Realizing the joint control of the relative position by using the relative position control force;

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

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

[0119] Among them, F is the relative position control force; k i 、k p 、k d are the integral, proportional, and differential coefficient 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 previous relative displacement deviation.

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

[0121] In the present invention, based on the PID feedback control combined with the position disturbance force F of the load platform as the feedforward parameter i, the relative position of the payload platform and the satellite platform is jointly controlled, and compared with single feedback control, the control accuracy and control stability are significantly improved.

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

[0123] Calculate the control torque according to the second feedforward parameter and the feedback parameter;

[0124] Use the control torque to jointly control the attitude of the satellite platform;

[0125] Among them, 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 derivative coefficient of the PID controller respectively; α is the attitude information to be controlled of the satellite platform in the feedback parameter; M b1 , M b2 , M b3 are the first disturbance torque, the second disturbance torque, and the third disturbance torque in the second feedforward parameter respectively.

[0128] It should be noted that k ib , k pb , k db are all 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 on the non-rotation axis. α' = α / dt.

[0129] In the present invention, based on the PID feedback control combined with the first disturbance torque, the second disturbance torque, and the third disturbance torque of the satellite platform as the feedforward parameters, the attitude of the satellite platform is jointly controlled, and 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 attitude of the attitude platform is jointly controlled according to the third feedforward parameter, the feedback parameter, and the previous attitude information of the load platform in the previous eddy current sampling period, including:

[0132] Calculate the control torques on the rotating axis and the non-rotating axis according to the third feedforward parameter, the feedback parameter, and the previous attitude information;

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

[0134]

[0135] where M_X, M_Y, and 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, and ω_Z are the components of the rotational angular velocity in the rolling direction of the load platform on the X-axis, Y-axis, and Z-axis, respectively; ω_X_lst is the component of the rotational angular velocity in the rolling direction of the load platform on the X-axis in the previous attitude information; A_Y and 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 are the components of the attitude disturbance torque M i of the load platform on the non-rotating axes Y and Z, respectively; k ix 、k px 、k dx are the integral, proportional, and derivative coefficient of the PID controller on the rotating axis X-axis, respectively; k iy 、k py 、k dy are the integral, proportional, and derivative coefficient of the PID controller on the rotating axis Y-axis, respectively; k iz 、k pz 、k dz are the integral, proportional, and derivative coefficient of the PID controller on the rotating axis Z-axis, respectively.

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

[0137] In a specific embodiment, taking the X-axis as the rotating axis and the Y and Z axes as the non-rotating axes, Figure 5 shows the attitude disturbance schematic diagram of the satellite platform, where mb2i = M b1 +M b2 +M b3 ; mb2i = [mb2ix mb2iy mb2iz] T ; It can be seen from Figure 5 that the influence of the interference on the X-axis is relatively low, and the influence of the interference on the Y and Z axes is more obvious. Therefore, it is necessary to control the interference on the non-rotating axes.Figure 6 It shows a schematic diagram of the control effect of the pitch attitude with feedback control before 500 s and the combined feedforward and feedback control after 500 s; Figure 7 It shows a schematic diagram of the control effect of the pitch attitude angular velocity with feedback control before 500 s and the combined feedforward and feedback control after 500 s. From Figure 6 and Figure 7 it can be seen that the stability of the pitch angle and pitch angular velocity with only feedback control is poor, while the stability and control accuracy of the pitch angle and pitch angular velocity with the combined feedforward and feedback control are significantly higher.

[0138] As Figure 8 , Figure 9 shown, the embodiment of the present invention provides a satellite on-orbit identification and combined feedforward and feedback control device based on a magnetic levitation bearing. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. In terms of the hardware level, as Figure 8 shown, it is a hardware architecture diagram of a computing device where the satellite on-orbit identification and combined feedforward and feedback control device based on a magnetic levitation bearing provided by the embodiment of the present invention is located. In addition to Figure 8 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device is located in the embodiment usually also includes other hardware, such as a forwarding chip responsible for processing packets, etc. Taking the software implementation as an example, as Figure 9 shown, as a logically meaningful device, it is formed by the CPU of its computing device reading the corresponding computer program in the non-volatile memory into the memory and running. A satellite on-orbit identification and combined feedforward and feedback control device provided by the present embodiment is applied to a satellite in which both the satellite platform and the payload platform are large-inertia rigid bodies, and includes:

[0139] An acquisition module 900, configured to acquire the first motion parameter of the payload platform, the second motion parameter of the satellite platform, and the position vectors of the satellite platform and the payload platform in the satellite during the rotation process of the payload platform;

[0140] An on-orbit identification module 902, configured to determine the position interference force, attitude interference torque on the payload platform, and interference torque on the satellite platform during the rotation process according to the first motion parameter, the second motion parameter, the position vectors, and the mass of the payload platform; wherein, the interference torque includes a first interference torque introduced by the non-coincidence of the centroid and the rotation axis of the payload platform, a second interference torque introduced by the uneven mass distribution of the payload platform in the rotation axis direction, and a third interference torque introduced by the uneven mass distribution of the payload platform in the non-rotation axis direction;

[0141] The control module 904 is configured to perform joint control on the satellite according to the position disturbing force, attitude disturbing torque of the payload platform, and disturbing torque of the satellite platform.

[0142] In some specific embodiments, the acquisition module 900 can be used to execute the above step 100, the on-orbit identification module 902 can be used to execute the above step 102, and the control module 904 can 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 vector includes a first position vector from the magnetic levitation connection point in the payload platform coordinate system to the center of mass of the payload platform, and a second position vector from the center of mass of the satellite to the center of mass of the payload platform.

[0146] In some specific embodiments, in the on-orbit identification module 902, the position disturbing force of the payload platform is the centrifugal disturbing force received by the magnetic levitation bearing joint when the center of mass of the payload platform is not on the rotation axis during rotation;

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

[0148] F i =m i ω i ×(ω i ×r ci )

[0149] where F i is the position disturbing force of the payload platform; m i is the mass of the payload 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 levitation connection point in the payload platform coordinate system to the center of mass of the payload platform.

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

[0151] The attitude disturbing torque of the payload platform is determined by the following formula:

[0152] M i =m i r ci ×(ω i ×(ω i ×r 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 payload platform; m i is the mass of the payload platform; r ci is the first position vector from the magnetic levitation connection point to the centroid of the payload platform in the payload 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 payload platform; J sic is the matrix inertia defined based on the moment of inertia of the payload 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 embodiments, in the on-orbit identification module 902, the first disturbance 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 payload 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 centroid of the satellite to the centroid of the payload platform, r ci is the first position vector from the magnetic levitation connection point to the centroid of the payload platform in the payload 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 embodiments, in the on-orbit identification module 902, the second disturbance 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 interference torque; 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; ω i is the angular velocity of the relative motion of the payload platform with respect to the satellite platform; J ic is the moment of inertia of the payload platform.

[0160] In some specific embodiments, in the on-orbit 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] where M b3 is the third interference torque; 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; ω i is the angular velocity of the relative motion of the payload platform with respect to the satellite platform; J sic is the matrix inertia defined based on the moment of inertia of the payload platform; ω b is the angular velocity of the rotation of the satellite platform.

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

[0164] For each eddy current sampling period of the magnetic levitation bearing joint, the following operations are performed:

[0165] A1: Use the position interference force of the payload platform as the first feedforward parameter;

[0166] A2: Obtain the current relative displacement deviation of the satellite platform in this eddy current sampling period and the previous relative displacement deviation of the satellite platform in the previous eddy current sampling period;

[0167] A3: Perform joint control of the relative position between 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: Use the interference torque of the satellite platform as the second feedforward parameter;

[0169] A5: Obtain the current attitude information in this eddy current sampling period and use the current attitude information as the feedback parameter;

[0170] A6: Realize joint control of the attitude of the satellite platform based on the second feedforward parameter and the feedback parameter;

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

[0172] A8: Achieve the joint control of the attitude of the attitude platform according to the third feedforward parameter, the feedback parameter, and the previous attitude information of the payload platform in the previous eddy current sampling period.

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

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

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

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

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

[0178] Among them, F is the relative position control force; k i 、k p 、k d Are the integral, proportional, and differential coefficient terms of the PID controller respectively; F i Is the position disturbance force of the payload platform; r is the current relative displacement deviation; r_lst is the previous relative displacement deviation.

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

[0180] Calculate the control torque according to the second feedforward parameter and the feedback parameter;

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

[0182] Among them, 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 dbThey are the integral, proportional, and differential coefficients of the PID controller respectively; α is the attitude information to be controlled in the feedback parameters; M b1 、M b2 、M b3 They are the first disturbance torque, the second disturbance torque, and the third disturbance torque in the second feedforward parameters respectively.

[0185] In some specific embodiments, in step A8, according to the third feedforward parameter, the feedback parameter, and the previous attitude information, the control torques on the rotation axis and the non-rotation axis are calculated;

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

[0187]

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

[0189] It can be understood that the structure illustrated in the embodiments of the present invention does not constitute a specific limitation on a satellite on-orbit identification and feedforward-feedback joint control device based on a magnetic levitation bearing. In other embodiments of the present invention, a satellite on-orbit identification and feedforward-feedback joint control device based on a magnetic levitation bearing may include more or fewer components than those shown in the figure, or combine certain components, or split certain components, or have different component arrangements. The illustrated components can be implemented in hardware, software, or a combination of software and hardware.

[0190] For the information interaction, execution process, etc. among the modules in the above device, since they are based on the same concept as the method embodiments of the present invention, the specific content can be referred to the description in the method embodiments of the present invention, and will not be elaborated here.

[0191] An embodiment of the present invention further provides a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, a satellite on-orbit identification and feedforward-feedback combined control method according to any one of the embodiments of the present invention is implemented.

[0192] An embodiment of the present invention further provides a computer-readable storage medium. A computer program is stored on the computer-readable storage medium. When the computer program is executed by a processor, the processor is caused to execute a satellite on-orbit identification and feedforward-feedback combined control method according to any one of the embodiments of the present invention.

[0193] An embodiment of the present application further provides a computer program product. The computer program product 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 on-orbit identification and feedforward-feedback combined control method according to any one of the above embodiments.

[0194] Specifically, a system or device equipped with a storage medium can be provided. A software program code for implementing the functions of any one of the above embodiments is stored on the storage medium, and the computer (or CPU or MPU) of the system or device reads and executes the program code stored in the storage medium.

[0195] In this case, the program code read from the storage medium itself can implement the functions of any one of the above embodiments. Therefore, the program code and the storage medium storing the program code constitute a part of the present invention.

[0196] Embodiments of the storage medium for providing program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Optionally, the program code can be downloaded from a server computer via a communication network.

[0197] In addition, it should be clear that not only can the actual operations be completed in part or in whole by executing the program code read by the computer, but also by an operating system or the like operating on the computer based on the instructions of the program code, so as to implement the functions of any one of the above embodiments.

[0198] In addition, it can be understood that the program code read from the storage medium is written into the memory provided in the expansion board inserted into the computer or into the memory provided in the expansion module connected to the computer, and then based on the instructions of the program code, the CPU or the like installed on the expansion board or the expansion module is made to execute part or all of the actual operations, thereby implementing the functions of any one of the above 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 actual relationship or order between these entities or operations. Moreover, the term "comprising", "including" or any other variant thereof is intended to cover non-exclusive inclusion, so that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device. Without further limitation, an element defined by the statement "comprising an..." does not exclude the presence of additional identical elements in the process, method, article or device comprising the element.

[0200] Those of ordinary skill in the art can understand that all or part of the steps for implementing the above method embodiments can be completed by hardware related to program instructions. The foregoing program can be stored in a computer-readable storage medium. When the program is executed, it executes the steps including the above method embodiments; and the foregoing storage medium includes various media such as ROM, RAM, magnetic disk or optical disc that can store program code.

[0201] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A satellite on-orbit identification and combined feedforward and feedback control method based on magnetic levitation bearings, characterized in that, Applied to a satellite in which both the satellite platform and the payload platform are large-inertia rigid bodies, including: Obtaining the first motion parameters of the payload platform, the second motion parameters of the satellite platform, and the position vectors of the satellite platform and the payload platform in the satellite during the rotation of the payload platform; Determining the position interference force, attitude interference torque on the payload platform, and interference torque on the satellite platform during the rotation according to the first motion parameters, the second motion parameters, the position vectors, and the mass of the payload platform; wherein, the interference torque includes a first interference torque introduced by the non-coincidence of the centroid and the rotation axis of the payload platform, a second interference torque introduced by the uneven mass distribution of the payload platform in the rotation axis direction, and a third interference torque introduced by the uneven mass distribution of the payload platform in the non-rotation axis direction; Performing joint control on the satellite according to the position interference force, attitude interference torque of the payload platform, and the interference torque of the satellite platform.

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

3. The method according to claim 2, wherein The position interference force of the payload platform is the centrifugal interference force received by the magnetic levitation bearing joint when the centroid of the payload platform is not on the rotation axis during the rotation; The centrifugal interference force is determined by the following formula: F i = m i ω i × (ω i × r ci ) Among them, F i is the position interference force of the load platform; m i is the mass of the load platform; ω i is the angular velocity of the relative motion of the load platform with respect to the satellite platform; r ci is the first position vector from the magnetic levitation connection point to the centroid of the load platform in the load platform coordinate system.

4. The method according to claim 2, wherein The attitude interference torque of the payload platform is introduced by the uneven mass distribution of the payload platform during the rotation; The attitude interference torque of the payload platform is determined by the following formula: M i = m i r ci × (ω i × (ω i × r ci )) + ω i × J ic ω i + 2ω i J sic (A(θ)ω b ) Among them, M i is the attitude disturbance moment of the load platform; m i is the mass of the load platform; r ci is the first position vector from the magnetic levitation connection point to the centroid of the load platform in the load platform coordinate system; ω i is the angular velocity of the load 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 rotation of the satellite platform; A(θ) is the transformation matrix from the satellite platform coordinate system to the load platform coordinate system, and θ is the rotation angle of the load platform relative to the satellite platform about the rotation axis.

5. The method according to claim 2, wherein The first interference torque is determined by the following formula: Among them, M b1 is the first interference torque; m i is the mass of the load platform; ω i is the angular velocity of the load platform relative to the satellite platform, ω i =[ω x 00] T ; r oci is the second position vector from the centroid of the satellite to the centroid of the load platform, r oci =[r x r y r z T ; r ci is the first position vector from the magnetic levitation connection point in the load platform coordinate system to the centroid of the load platform, r ci =[r cix r ciy r ciz T ; θ is the rotation angle of the load platform relative to the satellite platform about the rotation axis.​​ 6. The method according to claim 2, characterized in that, The second interference torque is determined by the following formula: M b2 = -A(θ) T ω i × J ic ω i 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, and 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 payload platform.

7. The method according to claim 2, wherein 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 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; 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 payload platform; ω b is the angular velocity of the rotation of the satellite platform.

8. The method according to any one of claims 2 to 7, characterized in that, The performing joint control on the satellite according to the position interference force, attitude interference torque of the payload platform, and the interference torque of the satellite platform includes: For each eddy current sampling period of the magnetic levitation bearing joint, the following operations are performed: Taking the position interference force of the payload platform as the first feedforward parameter; Obtaining the current relative displacement deviation of the satellite platform in this eddy current sampling period and the previous relative displacement deviation of the satellite platform in the previous eddy current sampling period; Performing joint control on the relative position of the payload platform and the satellite platform according to the current relative displacement deviation, the previous relative displacement deviation, and the first feedforward parameter; Taking the interference torque of the satellite platform as the second feedforward parameter; Obtaining the current attitude information in this eddy current sampling period and taking the current attitude information as the feedback parameter; Realizing joint control of the attitude of the satellite platform according to the second feedforward parameter and the feedback parameter; Taking the attitude interference torque of the payload platform as the third feedforward parameter; Realizing joint control of the attitude of the payload platform according to the third feedforward parameter, the feedback parameter, and the previous attitude information of the payload platform in the previous eddy current sampling period.

9. The method according to claim 8, wherein The joint control of the relative position between the load platform and the satellite platform according to the current relative displacement deviation, the previous 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; Realizing the joint control of the relative position by using the relative position control force; 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 where F is the relative position control force; k i , k p , k d are the integral, proportional, and differential coefficient 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 previous relative displacement deviation.

10. The method according to claim 8, wherein The 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; Realizing the joint control of the attitude of the satellite platform by using the control torque; Wherein, the control torque is determined by the following formula: M b = -k ib ∫αdt - k pb α - 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 derivative coefficients of the PID controller respectively; α is the attitude information to be controlled in the feedback parameters; M b1 , M b2 , M b3 are the first disturbance torque, the second disturbance torque, and the third disturbance torque in the second feedforward parameter respectively.

11. A satellite on-orbit identification and feedforward-feedback joint control device based on a magnetic levitation bearing, characterized in that, Applied to a satellite in which both the satellite platform and the load platform are large inertia rigid bodies, including: An acquisition module, configured to acquire a first motion parameter of the load platform, a second motion parameter of the satellite platform, and a position vector of the satellite platform and the load platform in the satellite during the rotation of the load platform; An on-orbit identification module, configured to determine a position disturbance force on the load platform, an attitude disturbance torque, and a disturbance torque on the satellite platform during the rotation 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 non-coincidence of the center of mass and the rotation axis of the load platform, a second disturbance torque introduced by the uneven mass distribution of the load platform in the rotation axis direction, and a third disturbance torque introduced by the uneven mass distribution of the load platform in the non-rotation axis direction; A control module, configured to perform joint control on the satellite according to the position disturbance force of the load platform, the attitude disturbance torque, and the disturbance torque of the satellite platform.

12. A computing device, including a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, the method according to any one of claims 1-10 is implemented.

13. A computer-readable storage medium, on which a computer program is stored, and when the computer program is executed in a computer, the computer is made to execute the method according to any one of claims 1-10.

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