Simulation Modeling Based on Eddy Current Displacement Sensors and On-Orbit Star Sensor Calibration Method

Abstract

The present invention provides a simulation modeling and on-orbit star sensor calibration method based on an eddy current displacement sensor. The method includes: obtaining the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform; determining the eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship; for each dimension of the payload platform, perform: calculating the relative attitude deviation and the 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; wherein, 4 eddy current displacement sensors are arranged in each dimension; calibrating the star sensor of the payload platform according to the relative attitude deviation, the relative displacement deviation and the obtained installation error matrix of the star sensor of the payload platform. This solution uses the eddy current displacement sensor to realize the establishment of the simulation system and the calibration of the star sensor of the on-orbit payload platform, thereby improving the measurement accuracy.

Description

Technical Field

[0001] The present invention relates to the field of satellite control technology, in particular to the field of satellite simulation verification technology, and particularly to a simulation modeling based on eddy current displacement sensors and an in-orbit star sensor calibration method. Background Art

[0002] In recent years, with the demand for high resolution and ultra-wide coverage in space target measurement and imaging tasks, higher requirements have also been put forward for the dynamic control stability of complex connected multi-body satellites with large payloads and high dynamics. The high-speed rotating large payload platform and the satellite platform form a complex connected multi-body system consisting of a satellite body - a rotating joint subsystem - a payload. The rotating joint subsystem is connected by a magnetic levitation bearing joint, and the magnetic levitation bearing joint is a six-degree-of-freedom joint system. In this satellite system, the control of multiple degrees of freedom of the payload platform relative to the satellite platform is realized by using the magnetic levitation working principle, which has strong anti-interference ability and high reliability; a relatively high accuracy and measurement precision are obtained by using the non-contact eddy current displacement sensors in the magnetic levitation bearing joints.

[0003] Star sensors are important components in spacecraft. In order to ensure the accuracy of high-precision measurement and navigation, it is necessary to calibrate the star sensors of spacecraft. The existing calibration methods have relatively low measurement accuracy and do not meet the higher requirements for the dynamic control stability. Therefore, there is an urgent need to provide a simulation modeling based on eddy current displacement sensors and an in-orbit star sensor calibration method. Summary of the Invention

[0004] The present invention provides a simulation modeling based on eddy current displacement sensors and an in-orbit star sensor calibration method, which realizes the establishment of a simulation system and the calibration of the star sensors of the in-orbit payload platform by using eddy current displacement sensors, thereby improving the measurement accuracy.

[0005] In a first aspect, the present invention provides a simulation modeling based on eddy current displacement sensors and an in-orbit star sensor calibration method, and the method includes:

[0006] Obtain the installation parameters of the eddy current displacement sensors and the relative motion relationship between the satellite platform and the payload platform; the satellite platform and the payload platform are connected by a magnetic levitation joint; the eddy current displacement sensors are included in the magnetic levitation joint;

[0007] Determine the eddy current measurement vectors of the eddy current displacement sensors according to the installation parameters and the relative motion relationship;

[0008] 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 vectors of the eddy current displacement sensors in this dimension; 4 eddy current displacement sensors are provided in each dimension;

[0009] Calibrate the payload platform star sensor according to the relative attitude deviation, the relative displacement deviation, and the obtained installation error matrix of the payload platform star sensor.

[0010] In a second aspect, the present invention further provides a simulation modeling and on-orbit star sensor calibration device based on an eddy current displacement sensor, including:

[0011] An acquisition module, configured to acquire the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform; the satellite platform and the payload platform are connected by a magnetic levitation joint; the eddy current displacement sensor is included in the magnetic levitation joint;

[0012] An eddy current measurement module, configured to determine the eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship;

[0013] A decoupling module, configured to perform, for each dimension of the payload platform: calculate the relative attitude deviation and the relative displacement deviation of the satellite platform according to the installation parameters and the eddy current measurement vector of the eddy current displacement sensors in this dimension; where 4 eddy current displacement sensors are provided in each dimension;

[0014] A calibration module, configured to calibrate the payload platform star sensor according to the relative attitude deviation, the relative displacement deviation, and the obtained installation error matrix of the payload platform star sensor.

[0015] 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 simulation modeling and on-orbit star sensor calibration method according to any one of the above is implemented.

[0016] 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 simulation modeling and on-orbit star sensor calibration method according to any one of the above.

[0017] 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 one of the first aspects of this specification are implemented.

[0018] The present invention provides a simulation modeling and on-orbit star sensor calibration method based on an eddy current displacement sensor, which is applied to a satellite platform and a payload platform connected by a magnetic levitation joint, and the magnetic levitation joint includes an eddy current displacement sensor. The method first determines the eddy current measurement vector of the eddy current displacement sensor through the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform, so as to determine the relative attitude deviation and relative displacement deviation of the satellite platform by using the installation parameters of the 4 eddy current displacement sensors set in each dimension of the payload platform and the measured eddy current measurement vector. Finally, the calibration of the star sensor of the payload platform is realized by using the relative attitude deviation, relative displacement deviation and the installation error matrix of the star sensor of the payload platform. In this way, the present invention not only completes the dynamic modeling of the eddy current measurement sensor by using the working principle of the eddy current measurement sensor, but also realizes the calculation of the relative pose deviation, and then realizes the on-orbit calibration of the star sensor. The eddy current displacement sensor with high accuracy and measurement precision is used to further improve the measurement precision and navigation accuracy. At the same time, the ground calibration is realized based on the installation error matrix, and the final calibration of the star sensor of the payload platform is completed. Description of the Drawings

[0019] 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.

[0020] Figure 1 It is a flowchart of a simulation modeling and on-orbit star sensor calibration method based on an eddy current displacement sensor provided by an embodiment of the present invention;

[0021] Figure 2 It is a schematic diagram of the working principle of radial eddy current measurement provided by an embodiment of the present invention;

[0022] Figure 3 It is a schematic diagram of the working principle of axial eddy current measurement provided by an embodiment of the present invention;

[0023] Figure 4 It is a schematic diagram of decoupling calculation of relative attitude position in the radial measurement section provided by an embodiment of the present invention;

[0024] Figure 5 It is a schematic diagram of the relationship between the coordinate system of the payload platform and the rotating coordinate system of the payload platform provided by an embodiment of the present invention;

[0025] Figure 6 It is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0026] Figure 7This is a structural diagram of a simulation modeling and on-orbit star sensor calibration device based on an eddy current displacement sensor provided by an embodiment of the present invention. Detailed implementation manners

[0027] 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.

[0028] Please refer to Figure 1 , an embodiment of the present invention provides a simulation modeling and on-orbit star sensor calibration method based on an eddy current displacement sensor, including:

[0029] Step 100: Obtain the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform; the satellite platform and the payload platform are connected by a magnetic levitation joint; the magnetic levitation joint includes an eddy current displacement sensor;

[0030] Step 102: Determine the eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship;

[0031] Step 104: 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 the eddy current displacement sensors in this dimension; wherein, 4 eddy current displacement sensors are arranged in each dimension;

[0032] Step 106: Calibrate the star sensor of the payload platform according to the relative attitude deviation, relative displacement deviation, and the installation error matrix of the star sensor of the payload platform obtained.

[0033] In the present invention, in a satellite platform and a payload platform connected by a magnetic levitation joint, the magnetic levitation joint includes an eddy current displacement sensor. First, a three-dimensional model of the eddy current displacement sensor is constructed based on the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform. Then, based on this three-dimensional model, the eddy current measurement vector of the eddy current displacement sensor is determined. By using the installation parameters of the 4 eddy current displacement sensors arranged in each dimension of the payload platform and the measured eddy current measurement vector, the relative attitude deviation and relative displacement deviation of the satellite platform are determined. Finally, the calibration of the star sensor of the payload platform is realized by using the relative attitude deviation, relative displacement deviation, and the installation error matrix of the star sensor of the payload platform. Thus, the present invention not only completes the dynamic modeling of the eddy current measurement sensor by using the working principle of the eddy current measurement sensor, but also realizes the calculation of the deviation of the relative pose, and further realizes the on-orbit calibration of the star sensor. The eddy current displacement sensor with high accuracy and measurement precision is used to further improve the measurement precision and navigation accuracy. At the same time, the ground calibration is realized based on the installation error matrix to complete the final calibration of the star sensor of the payload platform.

[0034] The following describes Figure 1 the execution manner of each step shown.

[0035] Specifically, the coordinate systems in the present invention include: a satellite platform coordinate system, a payload platform coordinate system, a payload platform orbit coordinate system, and a payload platform rotation coordinate system:

[0036] 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, and the positive direction points to the payload platform. The positive direction of the z axis points to the I quadrant of the spacecraft (points to the center of the earth when there is no attitude deviation), perpendicular to the longitudinal axis. The y axis forms a right-handed system with the x and z axes;

[0037] In the payload platform coordinate system, the origin OU is the centroid of the payload platform. The uox axis is along the longitudinal axis of the satellite body, and the positive direction points away from the satellite platform. The positive direction of the uoz axis points to the I quadrant of the spacecraft (points to the center of the earth when there is no attitude deviation), perpendicular to the longitudinal axis. The uoy axis forms a right-handed system with the uox and uoz axes;

[0038] In the payload platform orbit coordinate system, the origin Oi is the centroid of the payload 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;

[0039] In the payload platform rotation coordinate system, the origin OU is the centroid of the payload platform. Before rotation, the payload platform coincides with the payload platform orbit coordinate system. The rotation coordinate system is the rotation coordinate system obtained by rotating the orbit coordinate system by a rotation angle around the rotation axis.

[0040] First, for step 100, 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 suspension joint includes a stator shaft and a rotor shaft;

[0041] 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 ;

[0042] 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 .

[0043] It should be noted that the magnetic suspension bearing rotor is installed on the rotor of the magnetic suspension joint, and the rotor of the magnetic suspension joint (i.e., the rotor shaft) is installed on the payload platform; the magnetic suspension bearing stator 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. The working principle of the eddy current displacement sensor is to drive the change of the magnetic field around the eddy current coil through a pulsed current of a certain waveform, and generate eddy current at the induction of the metal parts of the central bearing. In the Figure 2 schematic diagram of the working principle of radial eddy current measurement shown, 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 of the rotor shaft of the payload platform U_X . The eddy current stator direction is consistent with the vector r L_X direction, which is fixed in the satellite platform coordinate system. Taking the direction as the same, the vector magnitude is ar L_X ; where a is a coefficient. The eddy current measurement rotor direction is consistent with the r U_X direction, which is a fixed direction in this system.

[0044] In step 102, the radial eddy current measurement vector is determined by the following formula:

[0045]

[0046] where r b is the radial eddy current measurement vector in the payload platform coordinate system;

[0047] CUO $C_{lp}$ is the transformation matrix of the load platform coordinate system of the load platform relative to the orbit system, and $C_{lp}^T$ is the transpose matrix of this transformation matrix;

[0048] $C$ LO is the transformation matrix of the satellite platform coordinate system of the satellite platform relative to the orbit system, and $C^T$ is the transpose matrix of this transformation matrix;

[0049] In the orbit system, $r$ L_U is the third position vector from the center of mass of the satellite platform to the center of mass of the load platform;

[0050] In the satellite platform coordinate system, $r$ L_X is the direction of the stator axis; $r$ L_P is the first position vector from the center of mass of the satellite platform to the measured end;

[0051] In the load platform coordinate system, $r$ U_X is the direction of the rotor axis; $r$ U_Q $ + r$ I is the second position vector from the center of mass of the load platform to the measurement end.

[0052] Specifically, as Figure 2 shown, the vector description relationship of the radial eddy current measurement vector $r$ unified in the orbit system is:

[0053]

[0054] Since in Figure 2 , in the same coordinate system, the radial eddy current measurement vector $r$ and the installation direction $r$ U_X of the eddy current at the fixed position in the load platform coordinate system 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:

[0055]

[0056] Substituting $a$ into the above radial eddy current measurement vector $r$, the radial eddy current measurement vector $r$ b in the load platform coordinate system is obtained:

[0057]

[0058] In the present invention, it is necessary to consider the relative attitude and relative position introduced by the large-inertia load platform to ensure the 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.

[0059] In a preferred embodiment, in the schematic diagram of the working principle of axial eddy current measurement as shown in Figure 3 , 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 z in the orbital system can be obtained through vector calculation:

[0060]

[0061] Since in Figure 3 , the axial eddy current measurement vector r z and the central bearing metal part r L_X fixedly connected to the satellite platform are in a mutually perpendicular vector relationship, so 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:

[0062]

[0063] Substituting b into the above-mentioned axial eddy current measurement vector r z , the axial eddy current measurement vector r s in the load platform coordinate system is obtained:

[0064]

[0065] For step 104, on the measurement section in each dimension, 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;

[0066] According to the installation parameters of each eddy current displacement sensor in this dimension and the radial eddy current measurement vector, the relative attitude deviation and relative displacement deviation of the satellite platform are calculated, including:

[0067] The projection of the radial eddy current measurement vector in the measurement section of this dimension is used as the relative distance;

[0068] According to the installation parameters, determine the displacement of each eddy current displacement sensor from the geometric center of the rotating shaft of the magnetic levitation joint in the positive direction;

[0069] According to the relative distance and displacement of each eddy current displacement sensor, the relative attitude deviation and relative displacement deviation are calculated.

[0070] It should be noted that the maglev 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.

[0071] In a preferred embodiment, the relative attitude deviation is determined by the following formula:

[0072]

[0073] The relative displacement deviation is determined by the following formula:

[0074]

[0075] where θ is the relative attitude deviation; P is the relative displacement deviation; d U_1 and 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 within the measurement section, respectively; d D_1 and 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 within the measurement section, respectively; L1 is the displacement 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 maglev joint; L2 is 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 maglev 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.

[0076] It should be noted that d U_1 and d U_2 and d D_1 and d D_2 are all scalars.

[0077] Specifically, for the eddy current measurement vectors measured by the 4 eddy current displacement sensors in each dimension, decoupling calculations of the relative position and relative attitude are performed. As shown in the eddy current data measurement situation in the radial measurement section in Figure 4 , the 4 eddy currents are installed as two 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. The central bearing metal component fixedly connected to the satellite platform in 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 attitude deviation θ and relative position deviation P between the satellite platform and the load platform can be decoupled and calculated through the relative distance information measured by the 4 eddy current displacement sensors. As shown in Figure 4As shown in the figure, 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 rotation axis of the magnetic levitation joint is -L2; L = L1 - (-L2).

[0078] In step 106, the installation error matrix of the payload platform star sensor is obtained by the following method:

[0079] Obtain the attitude transformation matrix C of the payload platform relative to the satellite platform du ;

[0080] According to the first measurement input matrix C of the star sensor on the satellite platform obtained si_d and the first installation matrix C sb_d , the second measurement input matrix C of the star sensor on the payload platform si_u and the second installation matrix C sb_u as well as the attitude transformation matrix C du , calculate to obtain the installation error matrix;

[0081] The installation error matrix C Δ is determined by the following formula:

[0082]

[0083] where C Δ is the installation error matrix; C du is the attitude transformation matrix; C si_d is the first measurement input matrix; C sb_d is the first installation matrix; C si_u is the second measurement input matrix; C sb_u is the second installation matrix.

[0084] Specifically, as Figure 5 shown, taking the X-axis as the rotation axis, then A_Y and A_Z are the relative angles of the two axes measurable by the eddy current. Considering the eddy current measurement characteristics and using the third-party measurement data for the attitude angle in the rolling direction, the attitude transformation matrix of the payload platform relative to the satellite platform can be obtained as C du = dcm(A_X, A_Y, A_Z, 123); where dcm(a, b, c, 123) is the function description of the attitude transformation matrix calculated from the attitude angles according to the 123 rotation sequence, and dcm represents the direction cosine matrix. According to the first measurement output matrix C si_d and the first installation matrix C sb_d of the star sensor on the satellite platform, determine the attitude matrix of the satellite platform coordinate system relative to the inertial system as According to the second measurement output matrix C si_u and the second installation matrix C sb_u of the star sensor on the payload platform, determine the attitude matrix of the payload platform coordinate system relative to the inertial system as Considering the time-delay characteristics of the star sensor, with the eddy current measurement acquisition as the nominal time, the star sensor data is extrapolated to the nominal time data. It should be noted that the superscript T represents the matrix transpose. The installation error matrix of the star sensor on the payload platform is C Δ , and the following relational expressions are obtained according to the above installation relationship:

[0085]

[0086] Therefore, the installation error matrix of the star sensor on the payload platform obtained from the above reasoning is:

[0087]

[0088] For step 106, calibrate the star sensor on the payload platform according to the relative attitude deviation, relative displacement deviation, and the obtained installation error matrix of the star sensor on the payload platform, including:

[0089] Calibrate the star sensor on the payload platform using the installation error matrix;

[0090] Obtain the first absolute attitude, first absolute position of the satellite platform, and the second absolute attitude and second absolute position of the payload platform;

[0091] Sum the first absolute attitude and the relative attitude deviation to obtain the current second absolute attitude of the payload platform;

[0092] Sum the first absolute position and the relative displacement deviation to obtain the current second absolute position of the payload platform;

[0093] Adjust the current second absolute attitude to be the same as the second absolute attitude and adjust the current second absolute position to be the same as the second absolute position to complete the in-orbit installation calibration of the star sensor on the payload platform.

[0094] In the present invention, the ground calibration of the star sensor on the payload platform is achieved through the installation error matrix to ensure that its performance under ideal conditions can meet the design requirements. At the same time, by comparing the preset second absolute attitude and the current second absolute attitude of the payload platform in real time, the attitude of the payload platform is adjusted in real time; and by comparing the preset second absolute position and the current second absolute position of the payload platform in real time, the position of the payload platform is adjusted in real time, so as to achieve the in-orbit calibration of the star sensor on the payload platform based on the adjustment of the absolute attitude and absolute position, ensuring the measurement accuracy and stability of the star sensor in practical applications.

[0095] In the present invention, by performing eddy current vector measurement on the eddy current displacement sensors on the magnetic levitation joint connecting the satellite platform and the payload platform, and through decoupling calculation of the eddy current measurement vectors, the relative attitude deviation and relative displacement deviation of the satellite platform are obtained, so that the relative attitude deviation and relative displacement deviation obtained based on the highly reliable magnetic levitation joint and the eddy current displacement sensors with high accuracy and high measurement precision have higher reliability, higher accuracy and higher measurement precision. Therefore, it can significantly improve the measurement accuracy of the star sensor of the on-orbit payload platform calibrated based on the relative attitude deviation and relative displacement deviation, thereby improving the success rate and data accuracy of the entire space mission.

[0096] As Figure 6 , Figure 7 shown, the embodiment of the present invention provides a simulation modeling and on-orbit star sensor calibration device based on eddy current displacement sensors. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, as Figure 6 shown, it is a hardware architecture diagram of a computing device where a simulation modeling and on-orbit star sensor calibration device based on eddy current displacement sensors provided by an embodiment of the present invention is located. In addition to Figure 6 the shown processor, memory, network interface, and non-volatile memory, 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 7 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 for operation. A simulation modeling and on-orbit star sensor calibration device based on eddy current displacement sensors provided in this embodiment includes:

[0097] An acquisition module 700, configured to acquire the installation parameters of the eddy current displacement sensors and the relative motion relationship between the satellite platform and the payload platform; the satellite platform and the payload platform are connected by a magnetic levitation joint; the magnetic levitation joint includes an eddy current displacement sensor;

[0098] An eddy current measurement module 702, configured to determine the eddy current measurement vectors of the eddy current displacement sensors according to the installation parameters and the relative motion relationship;

[0099] A decoupling module 704, configured to perform, for each dimension of the payload platform: calculate the relative attitude deviation and relative displacement deviation of the satellite platform according to the installation parameters and the eddy current measurement vectors of the eddy current displacement sensors on this dimension; where 4 eddy current displacement sensors are provided on each dimension;

[0100] The calibration module 706 is used to calibrate the load platform star sensor according to the relative attitude deviation, relative displacement deviation, and the obtained installation error matrix of the load platform star sensor.

[0101] In some specific embodiments, the acquisition module 700 can be used to execute the above-mentioned step 100, the eddy current measurement module 702 can be used to execute the above-mentioned step 102, the decoupling module 704 can be used to execute the above-mentioned step 104, and the calibration module 706 can be used to execute the above-mentioned step 106.

[0102] In some specific embodiments, the measured end of the eddy current displacement sensor is installed on the stator shaft of the satellite platform; the measuring end of the eddy current displacement sensor is installed on the rotor shaft of the load platform; the magnetic suspension joint includes a stator shaft and a rotor shaft;

[0103] The installation parameters include the first position vector r of the center of mass of the satellite platform to the measured end in the satellite platform coordinate system L_P 、the second position vector r of the center of mass of the load platform to the measuring end in the load platform coordinate system U_Q +r I 、the third position vector r of the center of mass of the satellite platform to the center of mass of the load 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 load platform coordinate system U_X ;

[0104] 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 load platform coordinate system of the load platform relative to the orbital system UO .

[0105] In some specific embodiments, in the eddy current measurement module 702,

[0106] The radial eddy current measurement vector is determined by the following formula:

[0107]

[0108] The axial eddy current measurement vector is determined by the following formula:

[0109]

[0110] Where r b is the radial eddy current measurement vector in the load platform coordinate system; r s is the axial eddy current measurement vector in the load platform coordinate system;

[0111] C UO$C_{LP}$ is the transformation matrix of the load platform coordinate system of the load platform relative to the orbit system, and $C_{LP}^T$ is the transpose matrix of this transformation matrix;

[0112] $C$ LO is the transformation matrix of the satellite platform coordinate system of the satellite platform relative to the orbit system, and $C^T$ is the transpose matrix of this transformation matrix;

[0113] In the orbit system, $r$ L_U is the third position vector from the center of mass of the satellite platform to the center of mass of the load platform;

[0114] In the satellite platform coordinate system, $r$ L_X is the direction of the stator axis; $r$ L_P is the first position vector from the center of mass of the satellite platform to the measured end;

[0115] In the load platform coordinate system, $r$ U_X is the direction of the rotor axis; $r$ U_Q +$r$ I is the second position vector from the center of mass of the load platform to the measurement end.

[0116] In some specific embodiments, on the measurement cross-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.

[0117] In some specific embodiments, the decoupling module 704 is further configured to perform the following operations:

[0118] Take the projection of the eddy current measurement vector in the measurement cross-section of this dimension as the relative distance;

[0119] Determine the displacement of each eddy current displacement sensor from the geometric center of the axis of the magnetic levitation joint in the positive direction according to the installation parameters;

[0120] Calculate the relative attitude deviation and relative displacement deviation based on the relative distance and displacement of each eddy current displacement sensor; the relative attitude deviation is determined by the following formula:

[0121]

[0122] The relative displacement deviation is determined by the following formula:

[0123]

[0124] where, $\theta$ is the relative attitude deviation; $P$ 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 cross-section respectively; $d$D_1 、d D_2 are the relative distances of the third and fourth eddy current displacement sensors on the lower eddy current surface within the measurement section, respectively; L1 is the displacement of the eddy current displacement sensor on the upper eddy current surface in the positive direction from the geometric center of the axis of the magnetic suspension joint; L2 is the displacement of the eddy current displacement sensor on the lower eddy current surface in the positive direction from the geometric center of the axis of the magnetic suspension joint; L is the displacement of the upper and lower eddy current surfaces in the positive direction, and L = L1 - L2.

[0125] In some specific embodiments, the installation error matrix of the payload platform star sensor is obtained by the following method:

[0126] Obtain the attitude transformation matrix C of the payload platform relative to the satellite platform du ;

[0127] According to the first measurement input matrix C of the star sensor of the satellite platform obtained si_d and the first installation matrix C sb_d , the second measurement input matrix C of the star sensor of the payload platform si_u and the second installation matrix C sb_u as well as the attitude transformation matrix C du , calculate to obtain the installation error matrix;

[0128] The installation error matrix C Δ is determined by the following formula:

[0129]

[0130] where C Δ is the installation error matrix; C du is the attitude transformation matrix; C si_d is the first measurement input matrix; C sb_d is the first installation matrix; C si_u is the second measurement input matrix; C sb_u is the second installation matrix.

[0131] In some specific embodiments, the calibration module 706 is further configured to perform the following operations:

[0132] Calibrate the payload platform star sensor using the installation error matrix;

[0133] Obtain the first absolute attitude, the first absolute position of the satellite platform and the second absolute attitude and the second absolute position of the payload platform;

[0134] Sum the first absolute attitude and the relative attitude deviation to obtain the current second absolute attitude of the payload platform;

[0135] Sum the first absolute position and the relative displacement deviation to obtain the current second absolute position of the load platform;

[0136] Adjust the current second absolute attitude to be the same as the second absolute attitude and adjust the current second absolute position to be the same as the second absolute position, thereby completing the installation and calibration of the star sensor on the on-orbit load platform.

[0137] It can be understood that the structure illustrated in the embodiments of the present invention does not constitute a specific limitation on a simulation modeling and on-orbit star sensor calibration device based on an eddy current displacement sensor. In other embodiments of the present invention, a simulation modeling and on-orbit star sensor calibration device based on an eddy current displacement sensor 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 components shown in the figure can be implemented in hardware, software, or a combination of software and hardware.

[0138] Regarding the information interaction, execution process, etc. between the various modules within the above-mentioned 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.

[0139] 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, the simulation modeling and on-orbit star sensor calibration method according to any one of the embodiments of the present invention is implemented.

[0140] 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 enabled to execute the simulation modeling and on-orbit star sensor calibration method according to any one of the embodiments of the present invention.

[0141] 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 the simulation modeling and on-orbit star sensor calibration method according to any one of the above embodiments.

[0142] Specifically, a system or device equipped with a storage medium can be provided. On the storage medium, software program codes for implementing the functions of any one of the above embodiments are stored, and the computer (or CPU or MPU) of the system or device is enabled to read and execute the program codes stored in the storage medium.

[0143] 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.

[0144] Examples of the storage medium for providing the program code include a floppy disk, a hard disk, a magneto-optical disk, an optical disk (such as a CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), a magnetic tape, a non-volatile memory card, and a ROM. Optionally, the program code can be downloaded from a server computer via a communication network.

[0145] 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 a computer, but also by causing an operating system or the like operating on the computer based on the instructions of the program code to complete part or all of the actual operations.

[0146] In addition, it can be understood that the program code read from the storage medium is written into the memory provided in an expansion board inserted into the computer or into the memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, a CPU or the like installed on the expansion board or the expansion module is caused to execute part or all of the actual operations, thereby implementing the functions of any one of the above-described embodiments.

[0147] 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 variant thereof is intended to cover a 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.

[0148] 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 such as ROM, RAM, magnetic disks, or optical disks that can store program code.

[0149] 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 for 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 simulation modeling and on-orbit star sensor calibration method based on an eddy current displacement sensor, characterized in that, Including: Obtaining the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform; The satellite platform and the payload platform are connected by a magnetic levitation joint; the eddy current displacement sensor is included in the magnetic levitation joint; Determining the eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship; For each dimension of the payload platform, the following is performed: According to the installation parameters and the eddy current measurement vector of the eddy current displacement sensors in this dimension, the relative attitude deviation and the relative displacement deviation of the satellite platform are calculated; wherein, 4 eddy current displacement sensors are arranged in each dimension; Calibrating the star sensor of the payload platform according to the relative attitude deviation, the relative displacement deviation and the obtained installation error matrix of the star sensor of the payload platform; On the measurement section of each dimension, the eddy current displacement sensors are distributed on the upper eddy current surface and the lower eddy current surface, and 2 of the eddy current displacement sensors are symmetrically installed at the center on both the upper eddy current surface and the lower eddy current surface; The calculating the relative attitude deviation and the relative displacement deviation of the satellite platform according to the installation parameters and the eddy current measurement vector of the eddy current displacement sensors in this dimension includes: Taking the projection of the eddy current measurement vector in the measurement section of this dimension as the relative distance; Determining the displacement amount in the positive direction of the geometric center of the axis of the magnetic levitation joint by each of the eddy current displacement sensors according to the installation parameters; Calculating the relative attitude deviation and the relative displacement deviation according to the relative distance and the displacement amount of each of the eddy current displacement sensors; 2. The method according to claim 1, wherein 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; The installation parameters include the first position vector from the center of mass of the satellite platform to the measured end, the second position vector from the center of mass of the payload platform to the measuring end, the third position vector from the center of mass of the satellite platform to the center of mass of the payload platform, the eddy current stator direction, and the eddy current measurement rotor direction; The relative motion relationship includes: the transformation matrix of the satellite platform coordinate system of the satellite platform relative to the orbital system, and the transformation matrix of the payload platform coordinate system of the payload platform relative to the orbital system; 3. The method according to claim 2, wherein The eddy current measurement vector includes a radial eddy current measurement vector and an axial eddy current measurement vector; The radial eddy current measurement vector is determined by the following formula: The axial eddy current measurement vector is determined by the following formula: where r b is the radial eddy current measurement vector in the load platform coordinate system; r s is the axial eddy current measurement vector in the load platform coordinate system; a and b are both coefficients; 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; 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; Under the orbital system, r L_U is the third position vector from the centroid of the satellite platform to the centroid of the payload platform; In the satellite platform coordinate system, r L_X is the direction of the stator axis; r L_P is the first position vector from the centroid of the satellite platform to the measured end; 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 center of mass of the load platform to the measurement end.

4. The method according to claim 1, wherein The relative attitude deviation is determined by the following formula: The relative displacement deviation is determined by the following formula: where, θ is the relative attitude deviation; P 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 within 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 within the measurement section, respectively; L1 is the displacement 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 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 of the upper eddy current surface and the lower eddy current surface in the positive direction, and L = L1 - L2.

5. The method according to claim 1, wherein The installation error matrix of the star sensor of the payload platform is obtained by the following method: Obtaining the attitude transformation matrix of the payload platform relative to the satellite platform; Calculate the installation error matrix based on the first measurement output matrix and the first installation matrix of the star sensor of the satellite platform, the second measurement output matrix and the second installation matrix of the star sensor of the payload platform, and the attitude conversion matrix obtained; The installation error matrix is determined by the following formula: where C Δ is the installation error matrix; C du is the attitude conversion matrix; C si_d is the first measurement input matrix; C sb_d is the first installation matrix; C si_u is the second measurement input matrix; C sb_u is the second installation matrix.

6. The method according to any one of claims 1 to 5, characterized in that Calibrating the star sensor of the payload platform according to the relative attitude deviation, the relative displacement deviation, and the obtained installation error matrix of the star sensor of the payload platform includes: Calibrating the star sensor of the payload platform using the installation error matrix; Obtain the first absolute attitude, the first absolute position of the satellite platform, and the second absolute attitude and the second absolute position of the payload platform; Sum the first absolute attitude and the relative attitude deviation to obtain the current second absolute attitude of the payload platform; Sum the first absolute position and the relative displacement deviation to obtain the current second absolute position of the payload platform; Adjust the current second absolute attitude to be the same as the second absolute attitude, and adjust the current second absolute position to be the same as the second absolute position to complete the installation calibration of the star sensor of the on-orbit payload platform.

7. A simulation modeling and on-orbit star sensor calibration device based on an eddy current displacement sensor, characterized in that, Used to implement the method according to any one of claims 1 to 6, including: An acquisition module for acquiring the installation parameters of the eddy current displacement sensor and the relative motion relationship between the satellite platform and the payload platform; the satellite platform and the payload platform are connected by a magnetic levitation joint; the eddy current displacement sensor is included in the magnetic levitation joint; An eddy current measurement module for determining the eddy current measurement vector of the eddy current displacement sensor according to the installation parameters and the relative motion relationship; A decoupling module for performing, for each dimension of the payload platform: calculating the relative attitude deviation and the relative displacement deviation of the satellite platform according to the installation parameters and the eddy current measurement vector of the eddy current displacement sensors in this dimension; where 4 eddy current displacement sensors are provided in each dimension; A calibration module for calibrating the star sensor of the payload platform according to the relative attitude deviation, the relative displacement deviation, and the obtained installation error matrix of the star sensor of the payload platform.

8. 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 1-6 is implemented.

9. 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 1-6.

Images