Spacecraft pure magnetic control method and system based on magnetic momenter

Abstract

The application provides a spacecraft pure magnetic control sun safety control method and system based on a magnetic moment device, which comprises the following steps: calculating a normal vector of a rotation shaft plane through a sun vector and a sun-pointing vector of the system, normalizing a magnetic field vector and constructing a unit orthogonal vector in the normal plane, solving an optimal rotation shaft vector in combination with a rotation inertia array, dynamically adjusting a control parameter according to a geometric included angle between the optimal rotation shaft and an inertial angular velocity of the system, superimposing a magnetic control sun pointing directional control amount and a damping control amount to obtain a total control amount. The system comprises vector calculation, optimal rotation shaft solving, control parameter adjustment and magnetic control amount output modules connected in sequence, and can be additionally provided with an execution driving module. The application fully considers the geometric relationship of multiple vectors, avoids attitude oscillation divergence, and can realize a sun pointing directional accuracy of ≤15° in less than one orbit period, thereby guaranteeing the charging of a solar cell array and providing a reliable pure magnetic control scheme for a spacecraft safety mode, and the engineering application value is remarkable.

Description

Technical Field

[0001] This invention relates to the field of aerospace control technology, and more specifically, to a method and system for pure magnetic control of spacecraft for safe solar orbit based on a magnetic torque generator. Background Technology

[0002] With the development of space technology and the continuous growth of aerospace demands, the missions undertaken by spacecraft in orbit are becoming more diversified and complex, placing increasingly higher demands on spacecraft attitude and orbit control systems, requiring high reliability and safety. To address the issue of ensuring spacecraft energy security under all abnormal conditions of the main actuators (reaction flywheel, control moment gyroscope, jet propulsion, etc.) in orbit, this paper considers using a magnetic torque generator as an actuator for spacecraft solar-orbiting safety control, ensuring the charging conditions of the spacecraft's solar arrays, and guaranteeing spacecraft energy security, which has significant engineering application value. Current spacecraft magnetic control algorithms generally do not consider the geometric relationships between the magnetic field, the solar vector, the spacecraft's inertial angular velocity, and the solar vector. The design process typically begins by calculating the ideal control torque, followed by calculating the magnetic moment through a cross product of this ideal control torque and the spacecraft's intrinsic magnetic field. However, the magnetic moment generated by the interaction of this magnetic moment and the intrinsic magnetic field is generally not in the same direction as the ideal control torque, making it difficult to achieve the desired control effect under pure magnetic control conditions. Regarding the method described in the paper "A Single-Axis Pointing Magnetic Control Algorithm Based on Geometric Analysis" (Journal of Astronautics), when the parallelism of the intersection line between the spacecraft's inertial angular velocity and angular acceleration planes and the rotation plane is poor, the spacecraft's attitude will oscillate and diverge, limiting the application of this method in practical engineering. Summary of the Invention

[0003] To address the shortcomings of existing technologies, the purpose of this invention is to provide a method and system for pure magnetic control of spacecraft for solar orbit safety based on a magnetic torque generator.

[0004] A method for pure magnetic control of a spacecraft for solar orbit based on a magnetic torque generator, according to the present invention, includes: Step S1: Based on the solar vector and the solar-opposite vector in the spacecraft's own system, calculate the normal vector of the rotation axis plane and the rotation vector from the solar vector in the spacecraft's own system to the solar-opposite vector in the spacecraft's own system through cross product operation; Step S2: Normalize the magnetic field vector of the spacecraft system measured by the magnetometer, construct auxiliary vectors, and then calculate two unit orthogonal vectors in the magnetic field normal plane of the system. Step S3: Combine the spacecraft's rotational inertia matrix, the normal vector of the rotation axis plane, and two unit orthogonal vectors to calculate the optimal rotation axis vector for attitude control; Step S4: Based on the geometric angle between the inertial angular velocity of the spacecraft body and the optimal rotation axis vector, dynamically adjust the control parameters, set the magnetic control quantity for sun orientation and the damping control quantity, and superimpose the two to obtain the magnetic control quantity for sun orientation safety, and drive the magnetic torque device to realize the pure magnetic control for sun orientation of the spacecraft.

[0005] Preferably, the calculation formulas for the rotation axis plane normal vector and the rotation unit vector from the solar vector in the local system to the solar vector in the spacecraft's local system in step S1 are as follows:

[0006]

[0007] in, The normal vector of the plane of rotation. This represents the solar vector in this system. This represents the solar vector within the spacecraft's intrinsic system. Represents vector to vector The unit vector of rotation, This represents the vector cross product operation.

[0008] Preferably, step S2 includes: Step S2.1: Normalize the magnetic field vector: ; in, This represents the magnetic field of the spacecraft's intrinsic system as measured by the magnetometer.

[0009] Step S2.2: Construct auxiliary vectors ,like ,So ,otherwise , Represents vector The first component; Step S2.3: Through , Calculate orthogonal vectors , .

[0010] Preferably, step S3 includes: Step S3.1: The spacecraft rotates the inertial array. Calculate the projection vector , , ; Step S3.2: Calculate the optimal rotation axis vector : like , ,otherwise .

[0011] Preferably, step S4 includes: Step S4.1: Calculate the vector , The angle between and the change per beat :

[0012]

[0013] in, For vectors , The angle between them is the value of the previous beat, with an initial value of 0; Step S4.2: Calculate the vector Spacecraft inertial angular velocity The angle between :

[0014] Step S4.3: Through 、 The numerical range of the control parameters can be dynamically set. , , ,when When the damping control weight is small, reduce the damping control weight. Increase the damping control weight when the value is large; Step S4.4: Calculate the magnetic control parameters for solar safety. :

[0015]

[0016]

[0017] in, This indicates the magnetic control quantity for sun orientation. This indicates the damping control quantity.

[0018] According to the present invention, a spacecraft pure magnetic control system for solar orbit based on a magnetic torque generator is provided to implement the control method described above. The system includes a vector calculation module, an optimal rotation axis solution module, a control parameter adjustment module, and a magnetic control quantity output module connected in sequence. The vector calculation module is used to collect the solar vector, solar vector, and magnetic field vector of the spacecraft's own system, and to calculate the normal vector of the rotation axis plane and two unit orthogonal vectors in the normal plane of the magnetic field of the own system; The optimal rotation axis solving module is used to calculate the optimal rotation axis vector for attitude control based on the rotation axis plane normal vector and the unit orthogonal vector, in conjunction with the spacecraft rotation inertia matrix. The control parameter adjustment module is used to collect the inertial angular velocity of the spacecraft body and dynamically adjust the control parameters according to the geometric angle between the optimal rotation axis vector and the inertial angular velocity. The magnetic control output module is used to calculate the magnetic control sun orientation control quantity and the damping control quantity according to the adjusted control parameters, and after superposition, obtain the magnetic control sun orientation safety control quantity and convert it into a magnetic moment output to the magnetic torque device.

[0019] Preferably, the vector calculation module includes a magnetic field normalization submodule and a vector operation submodule; The magnetic field normalization submodule is connected to the magnetometer signal and is used to normalize the magnetic field vector of the system measured by the magnetometer. The vector operation submodule has built-in cross product and dot product operation logic, which is used to calculate the normal vector of the plane of rotation and the unit orthogonal vector.

[0020] Preferably, the optimal rotation axis solution module has a built-in spacecraft rotation inertia array storage unit and a projection calculation submodule; The projection calculation submodule is used to calculate the projection vectors of the normal vector and the unit orthogonal vector of the rotation axis plane through the rotation inertia matrix, and then select the optimal rotation axis vector based on the dot product result of the projection vectors.

[0021] Preferably, the control parameter adjustment module includes an angle calculation submodule and a parameter matching submodule; The angle calculation submodule is used to calculate the angle between the optimal rotation axis vector and the relevant vectors, as well as the change in the angle. The parameter matching submodule has built-in preset parameter matching rules, which automatically match control parameters according to the range of included angle values. , , The value of is used to achieve dynamic adjustment of the weights of directional control and damping control.

[0022] Preferably, the system further includes an execution drive module, which is signal-connected to the magnetic control output module and the magnetic torquer, respectively, for converting the magnetic control solar orientation safety control quantity into a drive signal for the magnetic torquer. The system can be implemented using computer-readable program code, logic gates, programmable logic controllers, or embedded microcontrollers, and is suitable for a safety mode when all main actuators of the spacecraft malfunction, ensuring the charging conditions of the solar array and achieving a solar orientation control accuracy of ≤15°.

[0023] Compared with the prior art, the present invention has the following beneficial effects: 1. This invention employs a method that combines the solar vector, the sun vector, and the magnetic field direction of the system to calculate the optimal rotation axis. It fully considers the geometric relationship between the magnetic field, the solar / sun vector, and the inertial angular velocity. Compared with the prior art, which ignores this geometric relationship and causes the magnetic torque to deviate from the direction of the ideal control torque, this invention improves the accuracy of pure magnetic attitude control.

[0024] 2. This invention dynamically adjusts the control parameters based on the geometric angle between the optimal rotation axis vector and the inertial angular velocity, and designs a composite magnetic control quantity of orientation and damping. Compared with the attitude oscillation divergence problem that is prone to occur in the existing single-axis pointing magnetic control algorithm, this invention achieves stable attitude control of spacecraft and avoids inertial angular velocity divergence.

[0025] 3. This invention uses a magnetic torque generator as the sole actuator to achieve pure magnetic control for solar orientation. Simulation verification shows that a solar orientation accuracy within 15 degrees can be achieved in less than one orbital cycle. Compared with the traditional magnetic control method, which has a solar orientation error of more than 36 degrees and cannot meet charging requirements, this invention ensures the charging conditions of the solar cell array when all main actuators are abnormal, thus ensuring the energy safety of the spacecraft. Attached Figure Description

[0026] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a schematic diagram of the working method of the present invention; Figure 2 This is the solar vector diagram of the spacecraft in its intrinsic system corresponding to the magnetic control of the sun according to the present invention; Figure 3 This is the solar vector diagram of the spacecraft's intrinsic system, corresponding to the traditional method of magnetically controlled solar control. Detailed Implementation

[0027] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0028] This invention is applicable to the magnetically controlled solar orbit safety control of spacecraft. The normal direction of the rotation axis plane can be calculated from the solar orbit vector (generally the normal direction of the solar array) and the solar vector within the system. Then, based on the magnetic field direction and the normal direction of the rotation axis plane, the optimal rotation axis for attitude control can be calculated. Finally, based on the optimal rotation axis for attitude control and the inertial angular velocity of the spacecraft itself, the magnetically controlled solar orbit safety control parameters can be calculated. This invention solves the problem of ensuring spacecraft energy safety under all abnormal conditions of the main actuators (reaction flywheel, control moment gyroscope, jet propulsion, etc.) in orbit, ensuring the charging conditions of the spacecraft's solar array, and thus has significant engineering application value in ensuring spacecraft energy safety.

[0029] This invention calculates the spacecraft's magnetically controlled solar alignment safety control quantity based on the solar vector, the solar vector, and the inertial angular velocity of the spacecraft itself. This control quantity enables solar alignment in the spacecraft's safe mode, ensuring energy security in the overall satellite magnetic control mode. It should be noted that the angle thresholds, control parameter values, and vector component thresholds involved in the following embodiments are optimized values ​​after simulation verification and engineering practice. These values ​​can be directly used as the engineering implementation basis for the spacecraft's pure magnetically controlled solar alignment safety control. All optimized values ​​are selected while considering magnetically controlled orientation accuracy, attitude stability, and damping convergence efficiency, adapting to the safe mode solar alignment control requirements of near-Earth orbit spacecraft in the event of main actuator anomalies.

[0030] According to a method for spacecraft magnetically controlled solar alignment safety control based on a magnetic torque generator, the control quantities for spacecraft magnetically controlled solar alignment safety can be calculated from the solar alignment vector, the solar vector, and the inertial angular velocity of the spacecraft itself. Figure 1 As shown, it includes the following steps: Step S1: Calculate the normal vector of the rotation axis plane; in step S1, the normal vector of the rotation axis plane is calculated. The formula for calculating the unit vector of rotation from the solar vector to the solar vector in the spacecraft's own system is as follows:

[0031]

[0032] in, The normal vector of the plane of rotation. This represents the solar vector in this system. This represents the solar vector within the spacecraft's intrinsic system. Represents vector to vector The unit vector of rotation, This represents the vector cross product operation; Step S2: Calculate the two unit orthogonal vectors in the magnetic field plane of this system. , In step S2, two unit orthogonal vectors within the magnetic field plane of the system are calculated. , : Step S2.1: Normalize the magnetic field vector: ; in, This represents the magnetic field of the spacecraft's intrinsic system as measured by the magnetometer.

[0033] Step S2.2: Construct auxiliary vectors Auxiliary vector The choice should avoid vector parallel if ,So ,otherwise , Represents vector The first component; Step S2.3: Calculate orthogonal vectors , :

[0034]

[0035] Step S3: Calculate the spacecraft's rotational inertia matrix, the normal vector of the rotation axis plane, and the orthogonal vectors. , Calculate the optimal rotation axis vector; in step S3, the spacecraft's rotation inertia matrix, the rotation axis plane normal vector, and the orthogonal vector are used. , The process of calculating the optimal axis of rotation vector is as follows: Step S3.1: The spacecraft rotates the inertial array. Calculate the projection vector , , ; Step S3.2: Calculate the optimal rotation axis vector : like , ,otherwise

[0036] Step S4: Calculate the magnetically controlled solar orbital safety control quantities based on the spacecraft's inertial angular velocity and optimal rotation axis vector. The process of calculating the magnetically controlled solar orbital safety control quantities based on the spacecraft's inertial angular velocity and optimal rotation axis vector in step S4 is as follows: Step S4.1: Calculate the vector , The angle between and the change per beat :

[0037]

[0038] in, For vectors , The angle between them is the value of the previous beat, with an initial value of 0.

[0039] Step S4.2: Calculate the vector Spacecraft inertial angular velocity The angle between :

[0040] Step S4.3: Set control parameters , , The control parameters should be set with full consideration of the optimal shaft vector. Spacecraft inertial angular velocity Vector , Geometric relationship between them: when the optimal rotation axis vector With spacecraft inertial angular velocity When the included angle between them is small, the rotational effect of the magnetic torque converter on the daytime control should be fully utilized. At this time, the damping control effect of the magnetic torque converter should be appropriately reduced. When the optimal rotation axis vector... With spacecraft inertial angular velocity When the included angle between the two is large, the magnetic torquer's rotational effect on the sun's control is small or even harmful, causing the celestial body's inertial angular velocity to diverge. In this case, the damping control effect of the magnetic torquer should be appropriately increased, and the specific control parameters should be set accordingly. , , The specific steps are as follows: if : like and ,So , , ; like and ,So , , ; otherwise , , .

[0041] otherwise: like and ,So , , ; like and ,So , , ; otherwise , , .

[0042] Step S4.4: Calculate the magnetic control parameters for solar safety. :

[0043]

[0044]

[0045] in, This indicates the magnetic control quantity for sun orientation. This indicates the damping control quantity.

[0046] Furthermore, the specific description of the magnetically controlled solar orientation method of the present invention is as follows, through simulation comparison: The spacecraft's Z-axis is oriented to the sun via magnetic control; therefore, the closer the Sz component of the solar vector in the spacecraft's own system is to -1, the better the magnetically controlled solar orientation effect. Figure 2 It can be seen that good magnetically controlled solar orientation can be achieved in less than one orbital cycle, with an orientation accuracy within 15 degrees, which can fully meet the charging requirements of the solar array; through Figure 3 It can be seen that the traditional method of magnetic control for solar orientation cannot achieve stable solar orientation control, with an orientation error of more than 36 degrees, which affects the charging effect of the solar cell array.

[0047] The present invention also provides a spacecraft magnetically controlled solar eclipse safety control system based on a magnetic torque generator. The spacecraft magnetically controlled solar eclipse safety control system based on a magnetic torque generator can be implemented by executing the process steps of the spacecraft magnetically controlled solar eclipse safety control method based on a magnetic torque generator. That is, those skilled in the art can understand the spacecraft magnetically controlled solar eclipse safety control method based on a magnetic torque generator as a preferred embodiment of the spacecraft magnetically controlled solar eclipse safety control system based on a magnetic torque generator.

[0048] According to the present invention, a spacecraft pure magnetic control system for solar orbit safety based on a magnetic torque generator is provided. The control method for implementing this system includes a vector calculation module, an optimal rotation axis solution module, a control parameter adjustment module, and a magnetic control quantity output module connected in sequence.

[0049] The vector calculation module is used to acquire the solar vector, solar-opposite vector, and magnetic field vector within the spacecraft's intrinsic system, and to calculate the normal vector of the rotation axis plane and two unit orthogonal vectors within the magnetic field normal plane of the intrinsic system. The vector calculation module includes a magnetic field normalization submodule and a vector operation submodule. The magnetic field normalization submodule is connected to the magnetometer signal and is used to normalize the intrinsic system magnetic field vector measured by the magnetometer. The vector operation submodule has built-in cross product and dot product operation logic to complete the calculation of the normal vector of the rotation axis plane and the unit orthogonal vectors.

[0050] The optimal rotation axis solution module is used to calculate the optimal rotation axis vector for attitude control based on the spacecraft's rotational inertia matrix, using the rotation axis plane normal vector and the unit orthogonal vector. This module includes a built-in spacecraft rotational inertia matrix storage unit and a projection calculation submodule. The projection calculation submodule calculates the projection vectors of the rotation axis plane normal vector and the unit orthogonal vector using the rotational inertia matrix, and then selects the optimal rotation axis vector based on the dot product of the projection vectors.

[0051] The control parameter adjustment module is used to acquire the inertial angular velocity of the spacecraft and dynamically adjust the control parameters based on the geometric angle between the optimal rotation axis vector and the inertial angular velocity. The control parameter adjustment module includes an angle calculation submodule and a parameter matching submodule. The angle calculation submodule calculates the angle between the optimal rotation axis vector and relevant vectors, as well as the change in angle. The parameter matching submodule has built-in preset parameter matching rules and automatically matches control parameters based on the range of angle values. , , The value of is used to achieve dynamic adjustment of the weights of directional control and damping control.

[0052] The magnetic control output module calculates the magnetic control and damping control quantities for solar orientation based on the adjusted control parameters. These are then superimposed to obtain the magnetic control safety control quantity for solar orientation and converted into a magnetic moment output to the magnetic torque generator. The system also includes an execution drive module, which is signal-connected to the magnetic control output module and the magnetic torque generator. This module converts the magnetic control safety control quantity for solar orientation into a drive signal for the magnetic torque generator. The system can be implemented using computer-readable program code, logic gates, programmable logic controllers, or embedded microcontrollers. It is suitable for a safety mode when all main actuators of the spacecraft malfunction, ensuring the charging conditions of the solar array and achieving a solar orientation control accuracy of ≤15°.

[0053] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, making the system and its various devices, modules, and units of this invention function in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0054] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for pure magnetic control of a spacecraft for solar orbit based on a magnetic torque generator, characterized in that, include: Step S1: Calculate the normal vector of the rotation axis plane by cross product operation based on the solar vector and the solar vector in the spacecraft's own system. Step S2: Normalize the magnetic field vector of the spacecraft system measured by the magnetometer, construct auxiliary vectors, and then calculate two unit orthogonal vectors in the magnetic field normal plane of the system. Step S3: Combine the spacecraft's rotational inertia matrix, the normal vector of the rotation axis plane, and two unit orthogonal vectors to calculate the optimal rotation axis vector for attitude control; Step S4: Based on the geometric angle between the inertial angular velocity of the spacecraft body and the optimal rotation axis vector, dynamically adjust the control parameters, set the magnetic control quantity for sun orientation and the damping control quantity, and superimpose the two to obtain the magnetic control quantity for sun orientation safety, and drive the magnetic torque device to realize the pure magnetic control for sun orientation of the spacecraft.

2. The spacecraft pure magnetic control method for solar orbit safety control based on a magnetic torque generator according to claim 1, characterized in that, The formulas for calculating the rotational axis plane normal vector and the rotational unit vector from the solar vector in this system to the solar vector in the spacecraft's system, as mentioned in step S1, are as follows: in, The normal vector of the plane of rotation. This represents the solar vector in this system. This represents the solar vector within the spacecraft's intrinsic system. Represents vector to vector The unit vector of rotation, This represents the vector cross product operation.

3. The spacecraft pure magnetic control method for solar orbit safety control based on a magnetic torque generator according to claim 1, characterized in that, Step S2 includes: Step S2.1: Normalize the magnetic field vector: ; in, This represents the magnetic field of the spacecraft's intrinsic system as measured by the magnetometer. Step S2.2: Construct auxiliary vectors ,like ,So ,otherwise , Represents vector The first component; Step S2.3: Through , Calculate orthogonal vectors , .

4. The spacecraft pure magnetic control method for solar orbit safety control based on a magnetic torque generator according to claim 3, characterized in that, Step S3 includes: Step S3.1: The spacecraft rotates the inertial array. Calculate the projection vector , , ; Step S3.2: Calculate the optimal rotation axis vector : like , ,otherwise .

5. The spacecraft pure magnetic control method for solar orbit safety control based on a magnetic torque generator according to claim 2, characterized in that, Step S4 includes: Step S4.1: Calculate the vector , The angle between and the change per beat : in, For vectors , The angle between them is the value of the previous beat, with an initial value of 0; Step S4.2: Calculate the vector Spacecraft inertial angular velocity The angle between : Step S4.3: Through 、 The numerical range of the control parameters can be dynamically set. , , ,when When the damping control weight is small, reduce the damping control weight. Increase the damping control weight when the value is large; Step S4.4: Calculate the magnetic control parameters for solar safety. : in, This indicates the magnetic control quantity for sun orientation. This indicates the damping control quantity.

6. A spacecraft pure magnetic control system for solar orbit safety based on a magnetic torque generator, characterized in that, The control method described in any one of claims 1-5 includes a vector calculation module, an optimal rotating shaft solution module, a control parameter adjustment module, and a magnetic control quantity output module connected in sequence by signals. The vector calculation module is used to collect the solar vector, solar vector, and magnetic field vector of the spacecraft's own system, and to calculate the normal vector of the rotation axis plane and two unit orthogonal vectors in the normal plane of the magnetic field of the own system; The optimal rotation axis solving module is used to calculate the optimal rotation axis vector for attitude control based on the rotation axis plane normal vector and the unit orthogonal vector, in conjunction with the spacecraft rotation inertia matrix. The control parameter adjustment module is used to collect the inertial angular velocity of the spacecraft body and dynamically adjust the control parameters according to the geometric angle between the optimal rotation axis vector and the inertial angular velocity. The magnetic control output module is used to calculate the magnetic control sun orientation control quantity and the damping control quantity according to the adjusted control parameters, and after superposition, obtain the magnetic control sun orientation safety control quantity and convert it into a magnetic moment output to the magnetic torque device.

7. The spacecraft pure magnetic control system for solar orbit based on a magnetic torque generator according to claim 6, characterized in that, The vector calculation module includes a magnetic field normalization submodule and a vector operation submodule; The magnetic field normalization submodule is connected to the magnetometer signal and is used to normalize the magnetic field vector of the system measured by the magnetometer. The vector operation submodule has built-in cross product and dot product operation logic, which is used to calculate the normal vector of the plane of rotation and the unit orthogonal vector.

8. The spacecraft pure magnetic control system for solar orbit based on a magnetic torque generator according to claim 6, characterized in that, The optimal rotation axis solution module has a built-in spacecraft rotation inertia array storage unit and a projection calculation submodule; The projection calculation submodule is used to calculate the projection vectors of the normal vector and the unit orthogonal vector of the rotation axis plane through the rotation inertia matrix, and then select the optimal rotation axis vector based on the dot product result of the projection vectors.

9. The spacecraft pure magnetic control system for solar orbit based on a magnetic torque generator according to claim 6, characterized in that, The control parameter adjustment module includes an angle calculation submodule and a parameter matching submodule; The angle calculation submodule is used to calculate the angle between the optimal rotation axis vector and the relevant vectors, as well as the change in the angle. The parameter matching submodule has built-in preset parameter matching rules, which automatically match control parameters according to the range of included angle values. , , The value of is used to achieve dynamic adjustment of the weights of directional control and damping control.

10. The spacecraft pure magnetic control system for solar orbit based on a magnetic torque generator according to claim 6, characterized in that, The system also includes an execution drive module, which is signal-connected to the magnetic control output module and the magnetic torquer, respectively, and is used to convert the magnetic control of the daytime safety control quantity into the drive signal of the magnetic torquer.

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