A Payload-Centered Cooperative Control Method for Satellite Platforms

Abstract

The present invention provides a collaborative control method for a payload-centered satellite platform. The method includes: performing pointing control on the magnetic levitation joint according to the attitude deviation and attitude angular velocity deviation of the payload platform relative to the payload platform rotation coordinate system in the payload platform coordinate system; resolving the measurement data of the eddy current sensor of the magnetic levitation joint to obtain the displacement of the magnetic levitation joint, so as to control the relative position between the payload platform and the satellite platform; determining the output control torque and output control force of the magnetic levitation joints at different installation positions according to the desired control torque and desired control force of the payload platform; calculating the attitude error of the satellite platform relative to the payload platform according to the attitude deviation; and controlling the satellite platform according to the attitude deviation, attitude angular velocity deviation, displacement, output control torque, output control force, and attitude error. This solution improves the control accuracy and control stability of the payload platform, and further improves the control stability of the satellite platform.

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 method for collaborative control of a satellite platform centered on a payload. Background Art

[0002] The ultra-wide coverage high-resolution optical satellite consists of two platforms, a large payload platform rotating at high speed and a satellite platform, forming a complex connected multi-body system composed of a satellite body - a rotating joint subsystem - a payload. The rotating joint subsystem is connected by a magnetic suspension bearing joint. The magnetic suspension bearing joint is a six-degree-of-freedom joint system, including three translational degrees of freedom and three rotational degrees of freedom. In this satellite system, six-degree-of-freedom control of the payload platform relative to the satellite platform is achieved by utilizing the magnetic suspension working principle, which has strong anti-interference ability and high reliability; a relatively high accuracy and measurement precision are obtained by using non-contact eddy current displacement sensors in the magnetic suspension bearing joint. Therefore, as the devices carried by the satellite become more and more precise, extremely high requirements are put forward for the pointing accuracy and stability of the satellite platform, and there is an urgent need to provide a method for collaborative control of a satellite platform centered on a payload. Summary of the Invention

[0003] The present invention provides a method for collaborative control of a satellite platform centered on a payload, which can improve the control accuracy of the payload platform, and further improve the control stability of the satellite platform by stably controlling the payload platform.

[0004] In a first aspect, the present invention provides a method for collaborative control of a satellite platform centered on a payload, the method comprising:

[0005] Performing pointing control on the magnetic suspension joint according to the attitude deviation and attitude angular velocity deviation of the payload platform in the payload platform coordinate system relative to the payload platform rotation coordinate system; wherein, the payload platform and the satellite platform are connected by the magnetic suspension joint;

[0006] Calculating the displacement of the magnetic suspension joint by resolving the measurement data of the eddy current sensor of the magnetic suspension joint to control the relative position of the payload platform and the satellite platform;

[0007] Determining the output control torque and output control force of the magnetic suspension joints at different installation positions according to the desired control torque and desired control force of the payload platform;

[0008] Calculating the attitude error of the satellite platform relative to the payload platform according to the attitude deviation;

[0009] Controlling the satellite platform according to the attitude deviation, the attitude angular velocity deviation, the displacement, the output control torque, the output control force and the attitude error.

[0010] In a second aspect, the present invention further provides a payload-centered satellite platform cooperative control device, including:

[0011] A first operation module, configured to perform pointing control on the magnetic levitation joint according to the attitude deviation and attitude angular velocity deviation of the payload platform relative to the payload platform rotation coordinate system in the payload platform coordinate system; wherein, the payload platform and the satellite platform are connected through the magnetic levitation joint;

[0012] A second operation module, configured to resolve the measurement data of the eddy current sensor of the magnetic levitation joint to obtain the displacement amount of the magnetic levitation joint, so as to control the relative position of the payload platform and the satellite platform;

[0013] A third operation module, configured to determine the output control torque and output control force of the magnetic levitation joints at different installation positions according to the desired control torque and desired control force of the payload platform;

[0014] A fourth operation module, configured to calculate the attitude error of the satellite platform relative to the payload platform according to the attitude deviation;

[0015] A control module, configured to control the satellite platform according to the attitude deviation, the attitude angular velocity deviation, the displacement amount, the output control torque, the output control force and the attitude error.

[0016] In a third aspect, 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, the payload-centered satellite platform cooperative control method described in any one of the above is implemented.

[0017] In a fourth aspect, the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed in a computer, the computer is made to execute the payload-centered satellite platform cooperative control method described in any one of the above.

[0018] In a fifth aspect, an embodiment of the present invention further provides a computer program product, including computer instructions. When the computer instructions are executed by a processor, the steps of the method described in any first aspect of this specification are implemented.

[0019] The present invention provides a collaborative control method for a payload-centered satellite platform. The magnetic levitation joint uses the attitude deviation and attitude angular velocity deviation of the payload platform in the payload platform coordinate system relative to the payload platform rotation coordinate system as inputs to perform pointing control on the magnetic levitation joint; uses the displacement of the magnetic levitation joint calculated from the measurement data of the eddy current sensor as an input to control the relative position of the payload platform and the satellite platform; then determines the distribution of the output control torque and output control force of the magnetic levitation joint bearing based on the desired control torque and desired control force of the payload platform, and decouples and determines the attitude error of the satellite platform based on the attitude deviation, so as to perform absolute control on the payload platform based on the attitude deviation, attitude angular velocity deviation, output control torque, and output control force. When the satellite platform moves with the payload platform, collaborative control of the satellite platform is performed based on the displacement and attitude error on the basis of the star sensor of the satellite platform. Thus, the present invention controls the payload platform by controlling the magnetic levitation joint. When the satellite platform follows the payload platform, the attitude of the satellite platform is corrected by using the displacement and attitude error, so as to realize the collaborative control of the payload-centered satellite platform, improve the payload pointing accuracy, and realize the high-precision and high-stability control of the satellite payload platform camera. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] 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 use in 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, other drawings can be obtained based on these drawings without creative efforts.

[0021] Figure 1 is a flowchart of a collaborative control method for a payload-centered satellite platform provided by an embodiment of the present invention;

[0022] Figure 2 is a schematic diagram of the relationship between the payload platform coordinate system and the payload platform rotation coordinate system provided by an embodiment of the present invention;

[0023] Figure 3 is a schematic diagram of the working principle of radial eddy current measurement provided by an embodiment of the present invention;

[0024] Figure 4 is a schematic diagram of the working principle of axial eddy current measurement provided by an embodiment of the present invention;

[0025] Figure 5 is a schematic diagram of decoupling and calculating the displacement in the radial measurement section provided by an embodiment of the present invention;

[0026] Figure 6 is a schematic diagram of the output principle of the control force of the magnetic levitation joint provided by an embodiment of the present invention;

[0027] Figure 7 It is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0028] Figure 8 It is a structural diagram of a payload-centered satellite platform cooperative control device provided by an embodiment of the present invention. Specific embodiments

[0029] 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. Obviously, 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.

[0030] Please refer to Figure 1 , an embodiment of the present invention provides a payload-centered satellite platform cooperative control method, including:

[0031] Step 100: Perform pointing control on the magnetic suspension joint according to the attitude deviation and attitude angular velocity deviation of the payload platform relative to the payload platform rotation coordinate system in the payload platform coordinate system; wherein, the payload platform and the satellite platform are connected by a magnetic suspension joint;

[0032] Step 102: Solve the measurement data of the eddy current sensor of the magnetic suspension joint to obtain the displacement of the magnetic suspension joint, so as to control the relative position of the payload platform and the satellite platform;

[0033] Step 104: Determine the output control torque and output control force of the magnetic suspension joints at different installation positions according to the desired control torque and desired control force of the payload platform;

[0034] Step 106: Calculate the attitude error of the satellite platform relative to the payload platform according to the attitude deviation;

[0035] Step 108: Control the satellite platform according to the attitude deviation, attitude angular velocity deviation, displacement, output control torque, output control force, and attitude error.

[0036] In the present invention, first, the magnetic levitation joint uses the attitude deviation and attitude angular velocity deviation of the load platform coordinate system relative to the load platform rotation coordinate system as inputs to perform pointing control on the magnetic levitation joint; uses the displacement of the magnetic levitation joint calculated from the measurement data of the eddy current sensor as an input to control the relative position of the load platform and the satellite platform; then determines the distribution of the output control torque and output control force of the magnetic levitation joint bearing based on the desired control torque and desired control force of the load platform, and decouples the attitude error of the satellite platform based on the attitude deviation, so as to perform absolute control on the load platform based on the attitude deviation, attitude angular velocity deviation, output control torque, and output control force. When the satellite platform moves with the load platform, cooperative control of the satellite platform is performed based on the displacement and attitude error on the basis of the star sensor of the satellite platform. In this way, the present invention controls the load platform by controlling the magnetic levitation joint. When the satellite platform follows the load platform, the attitude of the satellite platform is corrected by using the displacement and attitude error, so as to realize the cooperative control of the satellite platform centered on the load, improve the load pointing accuracy, and realize the high-precision and high-stability control of the satellite load platform camera.

[0037] The execution manner of each step described below Figure 1 is shown.

[0038] Specifically, the coordinate systems adopted in the present invention include: the satellite platform coordinate system, the load platform coordinate system, the load platform orbital coordinate system, and the load platform rotation coordinate system:

[0039] In the satellite platform coordinate system, the origin OL is the centroid of the satellite platform, the Ox axis is along the longitudinal axis of the satellite body, the positive direction points to the load platform, the positive direction of the z axis points to the first quadrant of the satellite platform (points to the center of the earth when there is no attitude deviation), perpendicular to the longitudinal axis, and the y axis forms a right-handed system with the x and z axes;

[0040] In the load platform coordinate system, the origin OU is the centroid of the load platform, the uox axis is along the longitudinal axis of the satellite body, the positive direction points to the front end of the load platform, the positive direction of the uoz axis points to the first quadrant of the load platform (points to the center of the earth when there is no attitude deviation), perpendicular to the longitudinal axis, and the uoy axis forms a right-handed system with the uox and uoz axes;

[0041] In the load platform orbital coordinate system, the origin Oi is the centroid of the load platform, the xi axis points to the satellite running direction, the zi axis points to the center of the earth, and the yi axis forms a right-handed system with the xi and zi axes;

[0042] In the load platform rotation coordinate system, the origin OU is the centroid of the load platform. Before rotation, the load platform coincides with the load platform orbital coordinate system, and the rotation coordinate system is the rotation coordinate system obtained by rotating the orbital coordinate system by a rotation angle around the rotation axis;

[0043] In the geocentric inertial coordinate system, with the origin O Eis the centroid of the Earth, O E X E The axis points to the mean equinox of 2000.0, O E Z E The axis is perpendicular to the mean equatorial plane of 2000.0 and is in the same direction as the angular velocity of the Earth's rotation, O E Y E The axis and O E X E 、O E Z E axis form a right - hand system.

[0044] First, for step 100, according to the attitude deviation and attitude angular velocity deviation of the payload platform relative to the payload platform rotation coordinate system in the payload platform coordinate system, the pointing control of the magnetic suspension joint is performed, including:

[0045] Obtain the attitude matrix C of the payload platform coordinate system relative to the payload platform rotation coordinate system pr 、the platform angular velocity ω of the payload platform XG 、the nominal angular velocity ω of the payload platform in the payload platform rotation coordinate system T and the transformation matrix C of the payload platform coordinate system relative to the orbital system bo ;

[0046] Calculate the attitude deviation of the non - rotating axis according to the attitude matrix;

[0047] According to the platform angular velocity, use the orbital extrapolation algorithm to obtain the orbital angular velocity ω oi ;

[0048] According to the platform angular velocity, transformation matrix, orbital angular velocity and nominal angular velocity, calculate the attitude angular velocity deviation ω r .

[0049] In a preferred embodiment, the attitude deviation is determined by the following formula:

[0050] A_Z = arctan(C pr21 / C pr11 )

[0051] A_Y = - arctan(C pr31 / C pr11 )

[0052]

[0053] where, A_Z is the attitude deviation of the Z - axis; A_Y is the attitude deviation of the Y - axis; where, both the Z - axis and the Y - axis are non - rotating axes; C pr is the attitude matrix.

[0054] Specifically, for attitude deviation, the payload platform is equipped with gyro measurement sensors and star sensors. The inertial attitude C of the payload platform can be obtained through the method of extrapolating gyro data and correcting star sensor data pi , and the precise orbit information of the satellite is determined by using the orbit extrapolation algorithm oi . Assume that the payload platform rotates around the X-axis. Then, the angular velocity of the direction of the rotation axis of the payload platform's rotating coordinate system relative to the rotation axis of the payload platform's orbital coordinate system is ω x , the rotation angle around the rotation axis is θ, and the transformation matrix of the payload platform's rotating coordinate system relative to the payload platform's orbital coordinate system RX is described as follows

[0055]

[0056] The transformation matrix of the payload platform's rotating coordinate system relative to the geocentric inertial coordinate system ri is described as: C ri = C RX · C oi ;

[0057] After the attitude matrix C of the payload platform's coordinate system relative to the geocentric inertial system is obtained by filtering the star sensors and gyro measurement sensors of the payload platform pi , the attitude matrix C of the payload platform's coordinate system relative to the payload platform's rotating coordinate system is further obtained through transformation pr and is described as follows

[0058]

[0059] As Figure 2 shown, lox is the direction of the x-axis of the payload platform's rotating coordinate system, and uox, uoy, uoz are the x, y, and z axes of the payload platform's coordinate system. The description of lox in the uo_xyz coordinate system is

[0060]

[0061] From this, the description of the attitude deviation of the non-rotating axes (Y and Z axes) of the payload platform's coordinate system relative to the payload platform's rotating coordinate system is calculated as follows

[0062]

[0063] In a preferred embodiment, the attitude angular velocity deviation is determined by the following formula

[0064] ω r = ω XG - C bo ω oi - ω T

[0065] where ωr is the attitude angular velocity deviation; ω XG is the platform angular velocity; C bo is the transformation matrix; ω oi is the orbital angular velocity; ω T is the nominal angular velocity.

[0066] Specifically, the platform angular velocity information measured by the gyro sensor configured on the payload platform is ω XG , and the orbital angular velocity obtained by using the orbit extrapolation algorithm is ω oi . The nominal angular velocity in the rotating coordinate system of the payload platform is ω T , ω T = [ω x , 0, 0] T .

[0067] In the present invention, the magnetic levitation joint adopts an absolute pointing control method, using the attitude deviations of the Y and Z axes of the payload platform coordinate system relative to the rotating coordinate system of the payload platform as the input as the pointing control command to perform the pointing control of the rotating shaft.

[0068] In step 102, the magnetic levitation joint has three translational degrees of freedom, and 4 eddy current displacement sensors are installed on each translational degree of freedom;

[0069] The measurement data of the eddy current sensors of the magnetic levitation joint is resolved to obtain the displacement amount of the magnetic levitation joint, including:

[0070] For each translational degree of freedom of the payload platform, the following operations are all performed:

[0071] Obtain the eddy current sensor measurement data collected by each eddy current displacement sensor of this translational degree of freedom;

[0072] According to the installation parameters of the eddy current displacement sensors and the relative motion relationship between the satellite platform and the payload platform, calculate the eddy current measurement vector;

[0073] Take the projection of the eddy current measurement vector in the measurement section of this translational degree of freedom as the relative distance;

[0074] According to the installation parameters, determine the first displacement amount of each eddy current displacement sensor from the bearing geometric center of the magnetic levitation joint in the positive direction;

[0075] According to the relative distance and the first displacement amount of each eddy current displacement sensor, calculate the displacement amount of the magnetic levitation joint in this translational degree of freedom.

[0076] In the present invention, 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;

[0077] The installation parameters include the first position vector r from the centroid of the satellite platform to the measured end in the satellite platform coordinate system L_P and the second position vector r from the centroid of the payload platform to the measurement end in the payload platform coordinate system U_Q +r I and the third position vector r from the centroid of the satellite platform to the centroid of the payload platform in the payload platform orbital coordinate system L_U and the eddy current stator direction r in the satellite platform coordinate system L_X and the eddy current measurement rotor direction r in the payload platform coordinate system U_X ;

[0078] The relative motion relationships include: the transformation matrix C of the satellite platform coordinate system of the satellite platform relative to the orbital system LO and the transformation matrix C of the payload platform coordinate system of the payload platform relative to the orbital system UO .

[0079] It should be noted that the working principle of the eddy current displacement sensor is that the magnetic field around the eddy current coil changes by driving a pulsed current with a certain waveform, and eddy currents are induced at the metal parts of the central bearing. The eddy current measurement vectors include the radial eddy current measurement vector and the axial eddy current measurement vector. In the schematic diagram of the working principle of the radial eddy current measurement shown in Figure 3 , the installation position r of the eddy current displacement sensor U_X is parallel to the axis direction of the rotor shaft of the payload platform, and the measurement direction is the vertical direction of the axis r U_X of the rotor shaft of the payload platform. The eddy current stator direction is consistent with the direction of the vector r L_X , and the vector direction is fixed in the satellite platform coordinate system. Taking the direction as consistent, the vector magnitude is ar L_X ; where a is a coefficient. The eddy current measurement rotor direction is consistent with the direction of r U_X , which is a fixed direction in this system

[0080] Specifically, as shown in Figure 3 , the vector description relationship of the unified radial eddy current measurement vector r represented in the payload platform orbital coordinate system is:

[0081]

[0082] 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 a fixed position in the payload platform coordinate system U_X are perpendicular vector relationships, so considering the attitude motions 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:

[0083]

[0084] 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 :

[0085]

[0086] In the present invention, it is necessary to consider the relative attitude and relative position introduced by the large-inertia load platform to ensure high measurement accuracy and high control accuracy of the satellite. It should be noted that the large-inertia load is preferably such that the self-mass of the load platform is comparable to the mass of the satellite platform.

[0087] In a preferred embodiment, in the schematic diagram of the working principle of axial eddy current measurement as shown in Figure 4 , the installation position of the eddy current displacement sensor is perpendicular to the axis of the rotor shaft of the load platform, and the measurement direction is parallel to the axis of the rotor shaft of the load platform. Considering that the axial eddy current installation surface is perpendicular to the rotor shaft and the measurement direction is consistent with the rotor shaft, since the eddy current is installed on the load platform coordinate system, the representation of the installation position vector in this system is r U_Q +r I . It can be seen from the eddy current measurement that the axial eddy current measurement vector r in the orbital system can be obtained through vector calculation: z :

[0088]

[0089] Since in Figure 4 , the axial eddy current measurement vector r z and the central bearing metal component r L_X fixedly connected to the satellite platform are in a mutually perpendicular vector relationship, therefore considering the attitude motion of the satellite platform and the load platform and the relative motion relationship between the platforms, the following expression for the coefficient b can be obtained:

[0090]

[0091] Substitute b into the above axial eddy current measurement vector r z to obtain the axial eddy current measurement vector r in the load platform coordinate system: s :

[0092]

[0093] In a preferred embodiment, on the measurement cross-section of each translational degree of freedom, the 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;

[0094] When the translational degree of freedom is in the non-rotating axis direction, the displacement of the magnetic suspension joint is determined by the following formula:

[0095]

[0096] where y r is the displacement of the magnetic levitation joint in the non-rotating axis direction; d1 and d2 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; d3 and d4 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 first displacement of the eddy current displacement sensor located on the upper eddy current surface in the positive direction from the bearing geometric center of the magnetic levitation joint; L2 is the first displacement of the eddy current displacement sensor located on the lower eddy current surface in the positive direction from the bearing geometric center of the magnetic levitation joint; L is the displacement of the upper eddy current surface and the lower eddy current surface in the positive direction, and L = L1 - L2.

[0097] It should be noted that d1, d2, d3, and d4 are all scalars. When determining the first displacement, in the nominal locked state, the central bearing metal component fixed to the satellite platform is located at the center of the eddy current measurement.

[0098] Specifically, since the three translational degrees of freedom respectively correspond to translational motions along the X, Y, and Z axes, as described in the previous example, taking the X axis as the rotation axis and the Y and Z axes as the non-rotating axes respectively. For the non-rotating axes, decoupling calculations of the displacement amounts need to be performed for the radial eddy current measurement vectors measured by the 4 eddy current displacement sensors for the Y and Z axes respectively; similarly, for the rotating axis, decoupling calculations of the displacement amounts need to be performed for the axial eddy current measurement vectors measured by the 4 eddy current displacement sensors for the X axis. As Figure 5 shown by the eddy current sensor measurement data in the radial measurement section, the 4 eddy currents are installed as upper and lower eddy current surfaces in the load platform coordinate system, and 2 eddy current displacement sensors are symmetrically installed at the center of each eddy current surface. In the nominal locked state, the central bearing metal component fixed to the satellite platform 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 displacement amount of the magnetic levitation joint in the non-rotating axis direction can be decoupled and calculated through the relative distance information measured by the 4 eddy current displacement sensors. As Figure 5 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 rotating shaft of the magnetic levitation joint is -L2; L = L1 - (-L2).

[0099] As described in the previous example, for the rotating axis X axis, the measurement data of the eddy current sensors in the axial measurement section where the rotating axis is located is used for calculation to obtain the displacement amount y of the bearing geometric center of the X-axis suspension joint x, the displacement in the Y-axis direction is obtained as y y , the displacement in the Z-axis direction is obtained as y z . In the present invention, the displacement rs = [y x , y y , y z of the geometric center of the bearing of the magnetic suspension joint is used as the input to complete the design of the relative position controller for the payload platform and the satellite platform.

[0100] For step 104, according to the desired control torque and the desired control force of the payload platform, determine the output control torque and the output control force of the magnetic suspension joints at different installation positions, including:

[0101] Obtain the desired control torque according to the preset attitude information and the preset attitude angular velocity of the payload platform;

[0102] Obtain the desired control force according to the displacement of the magnetic suspension joint;

[0103] Obtain the installation information of the magnetic suspension joints at different positions, and determine the lever arms of the magnetic suspension joints at different positions from the satellite platform according to the installation information;

[0104] Using the output control principle of the magnetic suspension joint, based on the desired control torque, the desired control force and the lever arm, calculate the output control torque and the output control force of the magnetic suspension joints at different installation positions.

[0105] Specifically, obtain the preset attitude angular velocity ω of the payload platform according to the camera target pointing x , determine the preset attitude information according to the star sensor and the gyro measurement sensor of the payload platform, and design a feedforward-feedback combined controller for the payload platform using the preset attitude angular velocity and the preset attitude information to obtain the desired control torque Design a feedforward-feedback combined controller according to the displacement of the magnetic suspension joint obtained in step 102, and calculate the desired control force In the schematic diagram of the output principle of the magnetic suspension joint control force as shown in Figure 6 , at this time, the eddy current and the magnetic bearing joint coincide. The upper box is used to represent the payload platform, and the lower box is used to represent the satellite platform. The position control and the torque control output are realized through the output force distribution of the magnetic suspension joints at different installation positions. Taking the control distribution in the Y-axis direction as an example:

[0106]

[0107] Among them, are respectively the output control forces of the magnetic suspension joints at installation positions 1 and 2 in the Y-axis direction; is the desired control torque in the Z-axis direction; is the desired control force in the Y-axis direction; l 1y, l 2y are the corresponding lever arms respectively.

[0108] In step 106, the attitude error is determined by the following formula:

[0109] ψ bm = arcsin(l bx_b2 )

[0110]

[0111] l bx2 / l bx1 = tan(A_Z)

[0112] l bx3 / l bx1 = -tan(A_Y)

[0113] l bx1 2 + l bx2 2 + l bx3 2 = 1

[0114] where ψ bm , θ bm are the yaw direction error and pitch direction error in the attitude error respectively; A_Z is the attitude deviation of the Z-axis; A_Y is the attitude deviation of the Y-axis; C RX is the transformation matrix of the load platform rotation coordinate system relative to the orbit system; where the Z-axis and Y-axis are both non-rotating axes; l is the intermediate term.

[0115] In the embodiment of the present invention, following the previous example, taking the X-axis as the rotation axis and the Y-axis and Z-axis as non-rotating axes, the yaw direction error ψ bm and pitch direction error θ bm of the satellite platform are calculated using the attitude deviations A_Z and A_Y obtained in step 100. Specifically:

[0116] Let and there is From this, the normalized l bx is calculated;

[0117] l bx is transformed into the satellite platform coordinate system to obtain l bx_b : Considering the transfer order output as: Therefore, the attitude error of the satellite platform relative to the load platform is obtained

[0118] In the present invention, the payload platform adopts a star sensor for absolute control. By taking the attitude of the satellite platform relative to the payload platform as the input of the controller for free, the attitude correction of the satellite platform centered on the payload platform is completed.

[0119] To ensure the imaging mission requirements of the payload platform and improve the payload pointing accuracy. In the present invention, for a complex connected multi-body system with a large-inertia and high-dynamic payload platform, a satellite platform cooperative control strategy centered on the payload platform is proposed to meet the mission requirements of the control accuracy and stability of an optical satellite. In this strategy, the payload platform adopts absolute pointing control. The rotation axis pointing control is carried out based on the attitude deviations of the Y and Z axes of the payload platform body coordinate system relative to the nominal rotation coordinate system, and the rotational speed control is carried out based on the attitude angular velocity deviation of the rotation axis rotation. The relative motion between the satellite platform and the payload platform is measured by an eddy current displacement sensor. The relative displacement of the bearing geometric center of the magnetic levitation joint is calculated using the measurement data of the eddy current sensor. The relative position control is carried out through the magnetic levitation joint, and the relative pointing is controlled by the satellite platform. The satellite platform comprehensively uses the attitude errors of the Y and Z axes of the satellite platform relative to the payload platform calculated based on the eddy current displacement sensor and the attitude deviation information of the X axis relative to the orbital system obtained by the combined filtering of the star sensor and gyroscope of the satellite platform, and completes the relative attitude cooperative control of the Y and Z axes and the absolute attitude stability control of the X axis through a large torque momentum wheel.

[0120] It should be noted that the superscript T in the above formula is used to represent the transpose of a matrix.

[0121] As Figure 7 、 Figure 8 shown, the embodiment of the present invention provides a satellite platform cooperative control device centered on the payload. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, as Figure 7 shown, it is a hardware architecture diagram of a computing device where the satellite platform cooperative control device centered on the payload provided by the embodiment of the present invention. In addition to Figure 7 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device is located in the embodiment usually may also include other hardware, such as a forwarding chip responsible for processing packets, etc. Taking software implementation as an example, as Figure 8 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. The satellite platform cooperative control device centered on the payload provided by this embodiment includes:

[0122] The first operation module 800 is configured to perform pointing control on the magnetic levitation joint according to the attitude deviation and attitude angular velocity deviation of the payload platform relative to the payload platform rotation coordinate system in the payload platform coordinate system; wherein, the payload platform and the satellite platform are connected by the magnetic levitation joint;

[0123] The second operation module 802 is configured to calculate the measurement data of the eddy current sensor of the magnetic levitation joint to obtain the displacement of the magnetic levitation joint, so as to control the relative position of the payload platform and the satellite platform;

[0124] The third operation module 804 is configured to determine the output control torque and output control force of the magnetic levitation joints at different installation positions according to the desired control torque and desired control force of the payload platform;

[0125] The fourth operation module 806 is configured to calculate the attitude error of the satellite platform relative to the payload platform according to the attitude deviation;

[0126] The control module 808 is configured to control the satellite platform according to the attitude deviation, attitude angular velocity deviation, displacement, output control torque, output control force and attitude error.

[0127] In some specific embodiments, the first operation module 800 can be used to execute step 100 above, the second operation module 802 can be used to execute step 102 above, the third operation module 804 can be used to execute step 104 above, the fourth operation module 806 can be used to execute step 106 above, and the control module 808 can be used to execute step 108 above.

[0128] In some specific embodiments, the first operation module 800 is further configured to perform the following operations:

[0129] Obtain the attitude matrix of the payload platform coordinate system relative to the payload platform rotation coordinate system, the platform angular velocity of the payload platform, the nominal angular velocity of the payload platform in the payload platform rotation coordinate system, and the transformation matrix of the payload platform coordinate system relative to the orbital system;

[0130] Calculate the attitude deviation of the non-rotating axis according to the attitude matrix; the attitude deviation is determined by the following formula:

[0131] A_Z = arctan(C pr21 / C pr11 )

[0132] A_Y = -arctan(C pr31 / C pr11 )

[0133]

[0134] Wherein, A_Z is the attitude deviation of the Z-axis; A_Y is the attitude deviation of the Y-axis; wherein, both the Z-axis and the Y-axis are non-rotating axes; C pr is the attitude matrix;

[0135] According to the platform angular velocity, the orbital angular velocity is obtained by using the orbital extrapolation algorithm;

[0136] According to the platform angular velocity, the transformation matrix, the orbital angular velocity and the nominal angular velocity, the attitude angular velocity deviation is calculated; the attitude angular velocity deviation is determined by the following formula:

[0137] ω r = ω XG - C bo ω oi - ω T

[0138] Wherein, ω r is the attitude angular velocity deviation; ω XG is the platform angular velocity; C bo is the transformation matrix; ω oi is the orbital angular velocity; ω T is the nominal angular velocity.

[0139] In some specific embodiments, the magnetic levitation joint has three translational degrees of freedom, and 4 eddy current displacement sensors are installed on each translational degree of freedom;

[0140] The second operation module 802 is further configured to perform the following operations:

[0141] For each translational degree of freedom of the payload platform, the following operations are performed:

[0142] Obtain the eddy current sensor measurement data collected by each eddy current displacement sensor of this translational degree of freedom; on the measurement section of each translational degree of freedom, the 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;

[0143] According to the installation parameters of the eddy current displacement sensors and the relative motion relationship between the satellite platform and the payload platform, the eddy current measurement vector is calculated;

[0144] Take the projection of the eddy current measurement vector in the measurement section of this translational degree of freedom as the relative distance;

[0145] According to the installation parameters, determine the first displacement amount of each eddy current displacement sensor from the bearing geometric center of the magnetic levitation joint in the positive direction;

[0146] According to the relative distance and the first displacement amount of each eddy current displacement sensor, calculate the displacement amount of the magnetic levitation joint in this translational degree of freedom;

[0147] When the translation degree of freedom is in the non-rotating axis direction, the displacement of the magnetic suspension joint is determined by the following formula:

[0148]

[0149] where y r is the displacement of the magnetic suspension joint in the non-rotating axis direction; d1 and d2 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; d3 and d4 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 first displacement of the eddy current displacement sensor located on the upper eddy current surface in the positive direction from the bearing geometric center of the magnetic suspension joint; L2 is the first displacement of the eddy current displacement sensor located on the lower eddy current surface in the positive direction from the bearing geometric center of the magnetic suspension joint; L is the displacement of the upper eddy current surface and the lower eddy current surface in the positive direction, and L = L1 - L2.

[0150] In some specific embodiments, the third operation module 804 is further configured to perform the following operations:

[0151] Obtain the desired control torque according to the preset attitude information and preset attitude angular velocity of the load platform;

[0152] Obtain the desired control force according to the displacement of the magnetic suspension joint;

[0153] Obtain the installation information of the magnetic suspension joints at different positions, and determine the lever arms of the magnetic suspension joints at different positions from the satellite platform according to the installation information;

[0154] Using the magnetic suspension joint control output principle, calculate the output control torque and output control force of the magnetic suspension joints at different installation positions based on the desired control torque, desired control force and lever arm.

[0155] In some specific embodiments, the attitude error is determined by the following formula:

[0156] ψ bm = arcsin(l bx_b2 )

[0157]

[0158] l bx2 / l bx1 = tan(A_Z)

[0159] l bx3 / l bx1 = -tan(A_Y)

[0160] l bx1 2 +lbx2 2 +l bx3 2 = 1

[0161] where ψ bm and θ bm are the yaw direction error and pitch direction error in the attitude error respectively; A_Z is the attitude deviation of the Z-axis; A_Y is the attitude deviation of the Y-axis; C RX is the transformation matrix of the load platform rotation coordinate system relative to the orbit system; where the Z-axis and Y-axis are both non-rotating axes; l is an intermediate term.

[0162] It can be understood that the structure schematically shown in the embodiments of the present invention does not constitute a specific limitation on a load-centered satellite platform cooperative control device. In other embodiments of the present invention, a load-centered satellite platform cooperative control device may include more or fewer components than 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.

[0163] Regarding the information interaction, execution process, etc. between the various modules within 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.

[0164] The embodiments of the present invention further provide a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, it implements a load-centered satellite platform cooperative control method in any one of the embodiments of the present invention.

[0165] The embodiments of the present invention further provide a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the processor is caused to execute a load-centered satellite platform cooperative control method in any one of the embodiments of the present invention.

[0166] The embodiments of the present application further provide a computer program product, which includes a computer program. The 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 load-centered satellite platform cooperative control method described in any one of the above embodiments.

[0167] Specifically, a system or device equipped with a storage medium can be provided. A software program code for implementing the functions in 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.

[0168] In this case, the program code read from the storage medium itself can implement the functions of any one of the above-described embodiments. Therefore, the program code and the storage medium storing the program code constitute a part of the present invention.

[0169] Examples of the storage medium for providing the 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. Alternatively, the program code can be downloaded from a server computer via a communication network.

[0170] In addition, it should be clear that not only can the functions of any one of the above-described embodiments be implemented by executing the program code read by the computer, but also by causing an operating system or the like operating on the computer to perform part or all of the actual operations based on the instructions of the program code.

[0171] 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 part or all of the actual operations are performed by causing a CPU or the like installed on the expansion board or the expansion module to execute based on the instructions of the program code, thereby implementing the functions of any one of the above-described embodiments.

[0172] It should be noted that, in this document, 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 variation thereof is intended to cover non-exclusive inclusion, such 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 one..." does not exclude the presence of additional identical elements in the process, method, article or device comprising the element.

[0173] 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 performs the steps including the above method embodiments; and the foregoing storage medium includes various media that can store program code, such as ROM, RAM, magnetic disks, or optical disks.

[0174] 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 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 on some of the technical features; and these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A payload-centric satellite platform collaborative control method, characterized in that: include: Obtaining the attitude matrix of the load platform coordinate system relative to the load platform rotation coordinate system, the platform angular velocity of the load platform, the nominal angular velocity of the load platform in the load platform rotation coordinate system, and the transformation matrix of the load platform coordinate system relative to the orbital system; Calculate the attitude deviation of the non-rotational axis according to the attitude matrix; According to the platform angular velocity, the orbital angular velocity is obtained by using an orbital extrapolation algorithm; The attitude angular velocity deviation is calculated according to the platform angular velocity, the conversion matrix, the orbital angular velocity and the nominal angular velocity; wherein the payload platform and the satellite platform are connected by a magnetic levitation joint; Calculating the measurement data of the eddy current sensor of the magnetic levitation joint to obtain the displacement of the magnetic levitation joint to control the relative position of the load platform and the satellite platform; Determining the output control torque and output control force of the magnetic suspension joint at different installation positions according to the expected control torque and expected control force of the load platform; Calculating an attitude error of the satellite platform relative to the payload platform according to the attitude deviation; The satellite platform is controlled according to the attitude deviation, the attitude angular velocity deviation, the displacement, the output control torque, the output control force and the attitude error.

2. The method according to claim 1, characterized in that The posture deviation is determined by the following formula: A_Z=arctan(C pr21 / C pr11 ) A_Y=-arctan(C pr31 / C pr11 ) Where A_Z is the posture deviation of the Z axis; A_Y is the posture deviation of the Y axis; both the Z axis and the Y axis are non-rotational axes; C pr is the posture matrix.

3. The method according to claim 1, characterized in that The attitude angular velocity deviation is determined by the following formula: oh r =ω XG -C bo oh oi -oh T Among them, ω r is the attitude angular velocity deviation; ω XG is the platform angular velocity; C bo is the transformation matrix; ω oi is the orbital angular velocity; ω T is the nominal angular velocity.

4. The method according to claim 1, characterized in that: The magnetic levitation joint has three translational degrees of freedom, and each translational degree of freedom is equipped with four eddy current displacement sensors; The step of calculating the eddy current sensor measurement data of the magnetic levitation joint to obtain the displacement of the magnetic levitation joint includes: For each translational degree of freedom of the load platform, perform: Acquire the eddy current sensor measurement data collected by each eddy current displacement sensor of the translational degree of freedom; Calculating an eddy current measurement vector according to installation parameters of the eddy current displacement sensor and a relative motion relationship between the satellite platform and the load platform; Taking the projection of the eddy current measurement vector within the measurement section of the translational degree of freedom as the relative distance; Determine, according to the installation parameters, a first displacement of each of the eddy current displacement sensors from a bearing geometric center of the magnetic levitation joint in a positive direction; The displacement of the magnetic suspension joint in the translational degree of freedom is calculated according to the relative distance of each of the eddy current displacement sensors and the first displacement.

5. The method according to claim 4, characterized in that On the measurement section of each translational degree of freedom, 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 centrally symmetrically installed on the upper eddy current surface and the lower eddy current surface; When the translational degree of freedom is in the direction of the non-rotational axis, the displacement of the magnetic levitation joint is determined by the following formula: Among them, y r is the displacement of the magnetic levitation joint in the direction of the non-rotating axis; d1 and d2 are the relative distances of the first eddy current displacement sensor and the second eddy current displacement sensor of the upper eddy current surface within the measuring section; d3 and d4 are the relative distances of the third eddy current displacement sensor and the fourth eddy current displacement sensor of the lower eddy current surface within the measuring section; L2 is the first displacement of the eddy current displacement sensor located on the lower eddy current surface from the geometric center of the bearing 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.

6. The method according to claim 1, characterized in that The step of determining the output control torque and output control force of the magnetic suspension joint at different installation positions according to the expected control torque and expected control force of the load platform includes: Obtaining the desired control torque according to the preset attitude information and preset attitude angular velocity of the load platform; Obtaining the desired control force according to the displacement of the magnetic levitation joint; Acquire installation information of the magnetic levitation joints at different positions, and determine the moment arms of the magnetic levitation joints at different positions from the satellite platform according to the installation information; The output control torque and output control force of the magnetic suspension joint at different installation positions are calculated based on the expected control torque, the expected control force and the lever arm by utilizing the output control principle of the magnetic suspension joint.

7. The method according to any one of claims 1 to 6, characterized in that: The attitude error is determined by the following formula: ψ bm =arcsin(l bx_b2 ) l bx2 / l bx1 =tan(A_Z) l bx3 / l bx1 =-tan(A_Y) l bx1 2 +l bx2 2 +l bx3 2 ＝1 Among them, ψ bm ,θ bm are the yaw direction error and pitch direction error in the attitude error respectively; A_Z is the attitude deviation of the Z axis; A_Y is the attitude deviation of the Y axis; C RX is the transformation matrix of the load platform rotation coordinate system relative to the orbital system; wherein the Z axis and the Y axis are both non-rotational axes; and l are all intermediate items.

8. A payload-centered satellite platform collaborative control device, characterized in that: Used to implement the method according to any one of claims 1 to 7, comprising: A first computing module is used to control the pointing of the magnetic levitation joint according to the attitude deviation and attitude angular velocity deviation of the load platform in the load platform coordinate system relative to the load platform rotation coordinate system; wherein the load platform and the satellite platform are connected through the magnetic levitation joint; A second computing module is used to solve the eddy current sensor measurement data of the magnetic levitation joint to obtain the displacement of the magnetic levitation joint so as to control the relative position of the load platform and the satellite platform; A third computing module is used to determine the output control torque and output control force of the magnetic suspension joint at different installation positions according to the expected control torque and expected control force of the load platform; A fourth computing module, used for calculating the attitude error of the satellite platform relative to the payload platform according to the attitude deviation; A control module is used to control the satellite platform according to the attitude deviation, the attitude angular velocity deviation, the displacement, the output control torque, the output control force and the attitude error.

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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