A method and system for magnetically controlling the solar capture of a satellite's intermediate inertia axis

Abstract

The present invention discloses a method and system for magnetically controlling solar capture of a satellite's intermediate inertia axis. When a satellite enters a safety mode, the solar angle of the satellite's safety mode in a continuous time period is obtained, the solar state of the satellite's intermediate inertia axis is determined based on the solar angle, and the satellite is determined to enter a solar capture mode or a spinning mode based on the solar state. When the satellite enters the solar capture mode, the solar state of the satellite's intermediate inertia axis is determined based on the solar angle of the main plane. When the intermediate inertia axis is within the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters a spinning mode or remains in the solar capture mode is determined based on the gyroscopic angular velocity and solar angle. The method enables the satellite to adopt a minimum control configuration and energy, controls the satellite's attitude based on the satellite's solar angle and gyroscopic angular velocity, ensures that a sailboard in a sunlit area can efficiently capture the solar energy of the intermediate inertia axis, and achieves rapid replenishment of satellite energy.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite design, and in particular to a method and system for magnetically controlling the sun capture of a satellite's intermediate inertia axis. Background Art

[0002] There are numerous methods for purely magnetically controlling a satellite's solar alignment, including spinning and non-spinning methods. Methods for obtaining the solar vector include gyro integration and omnidirectional simulated solar sensitivity. Typically, the normal direction of a satellite's sailboard is parallel to the satellite's axis of maximum or minimum inertia. Therefore, these magnetically controlled solar alignment methods can only achieve alignment with the axis of maximum or minimum inertia. If the sailboard is mounted at an angle, such that the normal direction of the sailboard is close to the intermediate axis of inertia, these existing solar alignment methods are less applicable.

[0003] For example, a magnetically controlled non-spin solar orientation method achieves single-axis solar orientation adjustment of the satellite's main inertia by controlling the magnetic torquer's magnetic moment on a plane with the magnetic field strength as the normal vector. The patented "A Magnetic Control Solar Capture Method Considering the Solar Angle Change Rate and Control Factor" describes a magnetically controlled solar capture method that considers the solar angle change rate and control factor with variable weights. This type of method is a traditional magnetic control solar orientation method for large satellites. It relies on flywheel rotation to generate bias momentum to maximize the built-in angular momentum of a particular main inertia axis. Once gyroscopic axis stability is achieved, the upward amplitude of the magnetic output magnetic moment is magnetically controlled based on the solar angle change rate to achieve stable solar orientation control. The method described in the patented "A Purely Magnetic Control Spin Solar Orientation Method" is similar to the present invention. It uses the geomagnetic vector and the solar vector to determine the stellar angular velocity vector, and determines whether a damping torque is required based on the stellar angular velocity vector. The solar angle ε is determined based on the solar vector and the normal vector of the solar sail, and the solar angle control torque is calculated. The output spin control torque is calculated based on the difference between the spin angular velocity and the target angular velocity.

[0004] However, the above methods are difficult to achieve rapid solar capture of the intermediate inertia axis (the Yb axis of the satellite body). Therefore, it is necessary to design a magnetically controlled solar capture method for the intermediate inertia axis to ensure that the sailboard in the sunlit area can efficiently achieve solar capture of the intermediate inertia axis under the satellite energy safety mode. Summary of the Invention

[0005] In response to the problems existing in the prior art, the present invention provides a method and system for magnetically controlling the solar capture of the intermediate inertia axis of a satellite. For obliquely mounted sailboard satellites, under the harsh conditions of strong constraints such as energy failure, the method ensures that the sailboards in the sunlit area can efficiently achieve solar capture of the intermediate inertia axis.

[0006] The present invention is achieved through the following technical solutions:

[0007] A method for magnetically controlling the sun's capture of a satellite's intermediate inertia axis, wherein the satellite's solar panels are tiltedly mounted on the satellite, comprises the following steps:

[0008] Step 1: When the output power of the satellite panel facing the sun is less than the predetermined power, the satellite enters the safety mode;

[0009] Step 2: Obtain the solar angle of the satellite in safety mode for a continuous period of time, determine the solar state of the satellite's intermediate inertia axis based on the solar angle, and, based on the solar state and in combination with the satellite's control instructions, enable the satellite to enter a solar capture mode or a spin mode. In the spin mode, the solar panels obtain maximum solar energy.

[0010] Step 3: When the satellite enters the sun capture mode;

[0011] Obtain the main plane solar angle of the satellite, and determine the solar state of the satellite's intermediate inertia axis based on the main plane solar angle;

[0012] When the middle inertia axis is in the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters the spinning mode or remains in the solar capture mode is determined according to the gyroscopic angular velocity and solar angle.

[0013] Preferably, in step 2, the satellite obtains the gyro angular velocity in the spinning mode. When the absolute value of the angular velocity of any axis of the gyro angular velocity is greater than or equal to the set angular velocity change rate W1, the satellite enters the autonomous control damping mode according to the control instruction, otherwise the satellite maintains the autonomous control spinning mode.

[0014] Preferably, in step 2, when the solar angle is less than or equal to the set value β1, the satellite enters the spinning mode; when the solar angle is greater than or equal to the set value α1, the satellite enters the autonomously controlled solar capture mode.

[0015] Preferably, in step 3, when the main plane solar angle is greater than or equal to the set value α2, the intermediate inertia axis of the satellite is located within the sun-facing range; otherwise, the intermediate inertia axis of the satellite is out of the sun-facing range.

[0016] Preferably, in step 3, when the gyroscope's Abs (Y-axis angular velocity) ≤ the set value W2, Abs (Z-axis angular velocity) ≤ W2, and the sun angle ≤ α3, the satellite switches to the autonomously controlled spin mode;

[0017] When Abs (Y-axis angular velocity) ≤ W3, Abs (Z-axis angular velocity) ≤ W3, and the safety mode sun angle > α3, the satellite maintains the autonomously controlled sun capture mode;

[0018] When Abs (Y-axis angular velocity) ≥ W4 or Abs (Z-axis angular velocity) ≥ W4, the satellite switches to the autonomously controlled damping mode.

[0019] Preferably, in step 3, when the solar angle is ≤α4, the satellite switches to the autonomous control mode; otherwise, the satellite switches to the magnetic uncontrolled mode.

[0020] Preferably, the satellite enters a damping mode;

[0021] When the sun angle is less than or equal to the set value α6, the gyro angular velocity Abs (Y-axis angular velocity) is less than or equal to W3, and Abs (Z-axis angular velocity) is less than or equal to W3, the satellite switches to the autonomously controlled spin mode. Otherwise, the satellite remains in the damping mode.

[0022] When the sun angle is greater than the set value α6, the gyro angular velocity Abs (Y-axis angular velocity) ≤ W3, and Abs (Z-axis angular velocity) ≤ W3, the satellite switches to the autonomously controlled sun capture mode, otherwise, the satellite remains in the damping mode.

[0023] A satellite intermediate inertia axis magnetically controlled solar capture system, wherein the solar panels of the satellite are tiltedly installed on the satellite, comprising:

[0024] A safety module, used to put the satellite into safety mode when the output power of the satellite's sailboard toward the sun is less than a predetermined power;

[0025] The mode switching module is used to obtain the solar angle of the satellite safety mode for a continuous period of time, determine the solar state of the satellite's intermediate inertia axis based on the solar angle, and determine whether the satellite enters the solar capture mode or the spin mode based on the solar state. In the spin mode, the solar panels obtain the maximum solar energy;

[0026] Capture module, used when the satellite enters the sun capture mode;

[0027] Obtain the main plane solar angle of the satellite, and determine the solar state of the satellite's intermediate inertia axis based on the main plane solar angle;

[0028] When the middle inertia axis is in the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters the spinning mode or remains in the solar capture mode is determined according to the gyroscopic angular velocity and solar angle.

[0029] Compared with the prior art, the present invention has the following beneficial technical effects:

[0030] The present invention provides a method for magnetically controlling the solar capture of a satellite's intermediate inertia axis. This method addresses the harsh conditions of energy failure on satellites with obliquely mounted sailboards, adopts a minimum control configuration and energy, controls the satellite's attitude according to the satellite's solar angle and gyro angular velocity, and controls the intermediate inertia axis of the obliquely mounted sailboard satellite, i.e., the Xb axis of the satellite body, to rapidly capture the sun in all directions, thereby acquiring energy from solar arrays and ensuring that the sailboards in the sunlit area can efficiently capture the sun on the intermediate inertia axis, thereby rapidly replenishing the satellite's energy. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 Schematic diagram of the satellite body coordinate system of the present invention.

[0032] Figure 2 The figure is a flow chart of the intermediate inertia axis magnetically controlled solar capture method of the present invention. DETAILED DESCRIPTION

[0033] The present invention will be further described in detail below with reference to the accompanying drawings, which are intended to explain rather than limit the present invention.

[0034] The solar angle and principal plane solar angle appearing in the following embodiments are defined as follows:

[0035] Sun angle: refers to the angle between the sun vector and the normal of the sailboard.

[0036] Main plane solar angle: Here it refers to the angle between satellite Xb and the sun vector.

[0037] See Figure 1 and 2 A method for magnetically controlling the sun's capture of a satellite's intermediate inertial axis, wherein the satellite's solar panels are tiltedly mounted on the satellite, comprises the following steps:

[0038] Step 1: When the output power of the satellite's sailboard toward the sun is less than a predetermined power, the satellite's energy consumption unit is turned off and the satellite enters a safe mode;

[0039] Specifically, if a satellite experiences a power failure, the solar array will only output approximately 20W of power, making it impossible to use wheel-guided solar alignment. The satellite will then enter safe mode, using solar array magnetic alignment to ensure basic energy security. In safe mode, all energy-consuming components, including the heater, flywheel, and star sensor, are shut down, leaving only the magnetic assembly, gyroscope, and simulated solar sensor operational.

[0040] Step 2: Obtain the solar angle of the satellite safety mode in a continuous time period, and compare the solar angle with the set value α1 and the set value β1 respectively. According to the comparison results, the satellite enters the solar capture mode or the spin mode according to the control instruction. In the spin mode, the solar panels obtain the maximum solar energy.

[0041] In the solar capture mode, the satellite's controlled inertia axis (the solar inertia axis) is controlled to approach the direction of the solar vector, reducing the solar angle and bringing the solar panels as close to the sun as possible, thereby providing the basic conditions for spin.

[0042] Entering the solar capture mode, the satellite's magnetic capture control command P c1 as follows:

[0043] P c1 =(B b ×T c2 ) / |B b | 2

[0044]

[0045] V s×l =(S b ×V c ) / |S b ×V c |

[0046]

[0047] In the above formula, T c2 To capture the torque, S b is the solar vector of the satellite system, The solar vector change rate of this system, V c is the direction of the controlled inertia axis, k ps 、k ds is the control negative gain coefficient; α s is the angle between the current sun angle and the desired sun angle vector.

[0048] The spin mode is the solar spin sub-mode. In this mode, it is necessary to control the solar angle and the satellite's gyro angular velocity at the same time, suppress the solar angle so that the satellite's intermediate inertia axis remains selected to face the sun, and at the same time, the satellite is rotated by controlling the gyro angular velocity to achieve a passive gyro stabilization effect.

[0049] Satellite magnetic control spin algorithm control law design:

[0050] k w (ω bic -ω bi )

[0051]

[0052] Entering the spin mode, the magnetic control spin control instruction P c2 Output algorithm:

[0053] P c2 =(B b ×T c3 ) / |B b | 2

[0054] Among them, k w is the spin-up control parameter, ω bic is the target spin axis angular velocity.

[0055] The following is a detailed description of the method for switching the satellite control mode according to the solar angle, as follows:

[0056] S2.1. Obtain the continuous solar angle for T1 seconds, where T1 is 3-5 seconds. When the solar angle is ≥ α1 and α1 is 20-30°, the satellite enters the autonomous solar capture mode. When the solar angle is greater than 45°, it indicates that the satellite's axis of inertia with respect to the sun deviates too much from the main axis. For obliquely mounted sailboard satellites, since the non-main inertia is facing the sun, it is more reasonable to choose α1 of 20-30°.

[0057] S2.2. When the solar angle is ≤ β1 and β1 is 5-8°, the satellite enters the spin mode. At this time, the satellite's middle inertia axis faces the sun, and the solar panels obtain the maximum solar energy. β1 of zero indicates that the controlled inertia is facing the sun. Considering that the attitude error is about 3°, a β1 of 5-8° is selected before entering the spin mode to ensure more stable spin.

[0058] In the spinning mode, the satellite obtains the gyro angular velocity of the satellite safety mode for a continuous period of time. When the absolute value of the angular velocity of any axis of the gyro angular velocity is greater than or equal to the set angular velocity change rate W1, the satellite enters the autonomous control damping mode according to the control command. Otherwise, the satellite maintains the autonomous control spinning mode.

[0059] Damping mode: The total energy of the satellite can be divided into kinetic energy and potential energy. The kinetic energy is mainly caused by the satellite angular velocity, and the potential energy is mainly caused by the gravity gradient and the satellite gyroscopic effect. Three-axis rate damping is performed through the magnetic torquer to damp the entire satellite, which has a large kinetic energy after release, to a smaller kinetic energy state.

[0060] Magnetic damping control law design:

[0061] T c1 =-k(ω bi ×B b )=-kB bx / y / zdot

[0062] Entering the damping mode, the satellite's magnetic damping control command P b Output algorithm:

[0063]

[0064] Where: T c1 is the damping torque, The angular velocity of the satellite body relative to the inertial space, B b is the magnetic field strength of the satellite itself, B bx / y / zdot is the rate of change of the intrinsic magnetic field intensity, k and P0 are coefficients.

[0065] For example, if the judgment is continuously made for T1 seconds and the constant angular velocity is valid, when Abs (X-axis angular velocity)||Abs (Y-axis angular velocity)|| or Abs (Z-axis angular velocity) ≥ angular velocity change rate W1, where W1 is selected based on the satellite inertia and is 1 to 1.5° / s, the satellite enters the autonomously controlled damping mode to reduce the satellite's angular velocity and prevent the satellite's intermediate inertia axis from leaving the solar range. Otherwise, the satellite maintains the spinning mode to obtain maximum solar energy.

[0066] In step 3 and step 2, when the satellite enters the sun capture mode, the intermediate inertia axis is determined to be within the solar range according to the main plane solar angle. When the intermediate inertia axis is within the solar range, the satellite control mode is determined according to the satellite's solar angle and gyro angular velocity. The specific process includes the following:

[0067] S3.1. After entering the acquisition mode, obtain the main plane solar angle. If it is ≥α2 (range 45-50°), set the main plane stay flag to 1; otherwise, set it to 0. If the flag is 1, it means that the satellite's intermediate inertia axis is within the solar range; if the flag is 0, it means that the satellite's intermediate inertia axis is out of the solar range.

[0068] S3.2. Continuously obtain the satellite's gyroscopic angular velocity and solar angle during a set time period;

[0069] When Abs (Y-axis angular velocity) ≤ W2, Abs (Z-axis angular velocity) ≤ W2, W2 is 0.16-0.18° / s, and the sun angle ≤ α3, α3 is 18-25°, the satellite switches to the autonomously controlled spin mode.

[0070] When Abs (Y-axis angular velocity) ≤ W3, Abs (Z-axis angular velocity) ≤ W3, W3 is 0.20-0.22° / s, and the safety mode solar angle is greater than α3, the satellite maintains the autonomously controlled solar capture mode.

[0071] When Abs (Y-axis angular velocity) ≥ W4 or Abs (Z-axis angular velocity) ≥ W4, and W4 is 0.28-0.32° / s, the satellite switches to autonomously controlled magnetically controlled damping.

[0072] S3.3. When the satellite's gyroscopic angular velocity and solar angle do not fall under all conditions of step S3.2, the satellite remains in the sun capture mode;

[0073] When the safety mode solar angle is ≤α4, and α4 is 10-12°, the satellite switches to the autonomous control mode; otherwise, the satellite switches to the magnetic uncontrolled mode.

[0074] When the main plane stay flag is 1 and the main plane solar angle is ≤α5 for T1 seconds, and α5 is 8~10°, the satellite switches to the autonomous control mode. Otherwise, the satellite switches to the magnetic uncontrolled mode.

[0075] Step 4: When the satellite enters the magnetron damping mode, the solar angle and gyro angular velocity of the satellite are obtained, and the satellite is determined to enter the spinning mode or the solar capture mode according to the solar angle and the gyro angular velocity.

[0076] S4.1. When the sun angle is ≤ α6, α6 is 18-20°, the gyro angular velocity Abs (Y-axis angular velocity) is ≤ W3, and Abs (Z-axis angular velocity) is ≤ W3, the satellite switches to the autonomously controlled spin mode. Otherwise, the satellite remains in the damping mode.

[0077] S4.2. When the sun angle is greater than α6, the gyro angular velocity Abs (Y-axis angular velocity) ≤ W3, and Abs (Z-axis angular velocity) ≤ W3, the satellite switches to the autonomously controlled sun capture mode; otherwise, the satellite remains in the damping mode.

[0078] It should be noted that the values ​​of W2, W3, W4, α4, α5, and α6 need to match the satellite inertia and the magnetic moment of the actuator magnetic torquer, and should be selected according to the simulation results.

[0079] Example 1

[0080] See Figure 1 , a certain on-orbit scientific experimental satellite, the satellite body coordinate system Xb, Yb, Zb, the satellite inertia array after the sailboard is deployed is as follows, the maximum inertia axis is on the X axis

[0081]

[0082] The normal vector of the sailboard in the body coordinate system is as follows, and the sailboard is installed at an angle of 40 degrees.

[0083] L b =[cos(40°) 0 -sin(40°)] T

[0084] Taking the above-mentioned on-orbit scientific experimental satellite as an example, the intermediate inertia axis magnetic control solar capture method of the present invention under the condition of satellite energy shortage is described, which includes the following steps:

[0085] Satellite in-orbit status: The satellite experienced an energy failure. The solar array could only output about 20W of power, making it impossible to use wheel control for solar alignment. Therefore, the satellite switched to safe mode for solar alignment using solar array magnetic control to ensure basic satellite energy security. The following control procedures and strategies were adopted:

[0086] Step 1: Initialize the satellite's attitude control mode to spin toward the sun, turn off the heater, flywheel, and star sensor; and turn on the magnetic assembly and gyroscope.

[0087] Step 2: The magnetically controlled spin mode processing method is as follows;

[0088] The solar angle is continuously judged for 10 seconds. If the solar angle in safety mode is ≥30°, the satellite switches to solar capture mode and the control mode is autonomous.

[0089] The sun angle is continuously judged for 10 seconds. If the safe mode sun angle is ≤5°, the satellite will remain in the spin mode.

[0090] Continuously judge the gyro angular velocity for 3 seconds. If the angular velocity is valid, Abs (X-axis angular velocity) || Abs (Y-axis angular velocity) || or Abs (Z-axis angular velocity) ≥ 1.5° / s, the satellite switches to damping mode and the control mode is autonomous; otherwise, the working mode is spin mode and the control mode is autonomous.

[0091] Step 3: The magnetically controlled solar capture processing method is as follows:

[0092] After entering the capture mode, the main plane solar angle of the first shot is recorded. If it is ≥50°, the main plane stay flag Flag is set to 1, otherwise it is set to 0.

[0093] If the gyro angular velocity is continuously judged for 5s, when Abs (Y-axis angular velocity) ≤ 0.18° / s and Abs (Z-axis angular velocity) ≤ 0.18° / s, and the safety mode sun angle ≤ 20 degrees, the satellite switches the working mode to spin mode and the control mode is autonomous.

[0094] If the gyro angular velocity is continuously measured for 5 seconds, when Abs (Y-axis angular velocity) ≤ 0.22° / s and Abs (Z-axis angular velocity) ≤ 0.22° / s, and the safety mode sun angle is greater than 20 degrees, the satellite maintains the sun capture mode and the control mode is autonomous; that is, the angular velocity during the capture period is no greater than 0.22° / s, but the safety mode sun angle has not fallen into the 20-degree cone angle, and the capture continues.

[0095] If the gyro angular velocity is continuously measured for 5 seconds, when Abs (Y-axis angular velocity) ≥ 0.3° / s or Abs (Z-axis angular velocity) ≥ 0.3° / s, the satellite switches to magnetically controlled damping and the control mode is autonomous; that is, the angular velocity during capture is greater than 0.3° / s. Considering that if the angular velocity is too high, even if the main solar angle can be captured, it will not be able to successfully rotate in a short time, so magnetically controlled damping is entered to reduce the gyro angular velocity.

[0096] When the satellite's attitude does not belong to the above three situations, the satellite maintains the sun capture mode with an angular velocity between 0.22 and 0.3° / s, and the satellite drifts without control.

[0097] If the safety mode solar angle is ≤12°, the satellite’s control mode is autonomous; otherwise, the satellite’s launch control mode is magnetically uncontrolled.

[0098] If the main plane stay flag is set to 1 and the main plane sun angle is ≤ 10° for 5 consecutive seconds, the satellite's control mode is autonomous, meaning the sun is perpendicular to the main plane and the Z angular velocity is the largest of the three axes. Otherwise, the satellite's launch control mode is magnetically uncontrolled.

[0099] The present invention also provides a satellite intermediate inertia axis magnetic control solar capture system, wherein the solar panels of the satellite are tiltedly installed on the satellite, comprising:

[0100] A safety module, used to put the satellite into safety mode when the output power of the satellite's sailboard toward the sun is less than a predetermined power;

[0101] The mode switching module is used to obtain the solar angle of the satellite safety mode for a continuous period of time, determine the solar state of the satellite's intermediate inertia axis based on the solar angle, and determine whether the satellite enters the solar capture mode or the spin mode based on the solar state. In the spin mode, the solar panels obtain the maximum solar energy;

[0102] Capture module, used when the satellite enters the sun capture mode;

[0103] Obtain the main plane solar angle of the satellite, and determine the solar state of the satellite's intermediate inertia axis based on the main plane solar angle;

[0104] When the middle inertia axis is in the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters the spinning mode or remains in the solar capture mode is determined according to the gyroscopic angular velocity and solar angle.

[0105] The present invention provides a method for magnetically controlling the solar capture of a satellite's intermediate inertia axis. This method addresses the problem that, after an energy failure on a satellite with obliquely mounted sailboards, batteries cannot be charged, causing the entire satellite to enter a safety mode and making it difficult to achieve rapid solar capture of the intermediate inertia axis. The method adopts a minimum control configuration and energy, controls the satellite's attitude according to the satellite's solar angle and gyro angular velocity, and controls the intermediate inertia axis of the obliquely mounted sailboard satellite, i.e., the Xb axis of the satellite body, to achieve omnidirectional rapid solar capture, thereby achieving energy acquisition from the solar array. This method solves the problem of strong constraints and harsh conditions in the event of a satellite energy failure, ensuring that the sailboard in the sunlit area can efficiently achieve solar capture of the intermediate inertia axis, thereby achieving rapid replenishment of satellite energy.

[0106] The above content is only for explaining the technical idea of ​​the present invention and cannot be used to limit the protection scope of the present invention. Any changes made on the basis of the technical solution in accordance with the technical idea proposed by the present invention shall fall within the protection scope of the claims of the present invention.

Claims

1. A method for magnetically controlling the sun's capture of a satellite's intermediate inertia axis, wherein the satellite's solar panels are tiltedly mounted on the satellite, characterized in that: The following steps are involved: Step 1: When the output power of the satellite panel facing the sun is less than the predetermined power, the satellite enters the safety mode; Step 2: Obtain the solar angle of the satellite in safety mode for a continuous period of time, determine the solar state of the satellite's intermediate inertia axis based on the solar angle, and, based on the solar state and in combination with the satellite's control instructions, enable the satellite to enter a solar capture mode or a spin mode. In the spin mode, the solar panels obtain maximum solar energy. Step 3: When the satellite enters the sun capture mode; Obtain the main plane solar angle of the satellite, and determine the solar state of the satellite's intermediate inertia axis based on the main plane solar angle; When the middle inertia axis is in the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters the spinning mode or remains in the solar capture mode is determined according to the gyroscopic angular velocity and solar angle.

2. The method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 1, characterized in that: In step 2, the satellite obtains the gyro angular velocity in the spinning mode. When the absolute value of the angular velocity of any axis of the gyro angular velocity is greater than or equal to the set angular velocity change rate W1, the satellite enters the autonomous control damping mode according to the control command, otherwise the satellite maintains the autonomous control spinning mode.

3. The method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 1, characterized in that: In step 2, when the solar angle is less than or equal to the set value β1, the satellite enters the spin mode; when the solar angle is greater than or equal to the set value α1, the satellite enters the autonomously controlled solar capture mode.

4. The method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 1, characterized in that: In step 3, when the main plane solar angle is greater than or equal to the set value α2, the satellite's intermediate inertia axis is within the sun-facing range; otherwise, the satellite's intermediate inertia axis is out of the sun-facing range.

5. The method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 1, characterized in that: In step 3, when the gyroscope's Abs (Y-axis angular velocity) ≤ the set value W2, Abs (Z-axis angular velocity) ≤ W2, and the sun angle ≤ α3, the satellite switches to the autonomous control spin mode; When Abs (Y-axis angular velocity) ≤ W3, Abs (Z-axis angular velocity) ≤ W3, and the safety mode sun angle > α3, the satellite maintains the autonomously controlled sun capture mode; When Abs (Y-axis angular velocity) ≥ W4 or Abs (Z-axis angular velocity) ≥ W4, the satellite switches to the autonomously controlled damping mode.

6. The method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 1, characterized in that: In step 3, when the solar angle ≤ α4, the satellite switches to the autonomous control mode, otherwise, the satellite switches to the magnetic uncontrolled mode.

7. A method for magnetically controlling the solar capture of a satellite intermediate inertia axis according to claim 2 or 5, characterized in that: The satellite enters a damping mode; When the sun angle is less than or equal to the set value α6, the gyro angular velocity Abs (Y-axis angular velocity) is less than or equal to W3, and Abs (Z-axis angular velocity) is less than or equal to W3, the satellite switches to the autonomously controlled spin mode. Otherwise, the satellite remains in the damping mode. When the sun angle is greater than the set value α6, the gyro angular velocity Abs (Y-axis angular velocity) ≤ W3, and Abs (Z-axis angular velocity) ≤ W3, the satellite switches to the autonomously controlled sun capture mode, otherwise, the satellite remains in the damping mode.

8. A satellite intermediate inertia axis magnetically controlled solar capture system, wherein the solar panels of the satellite are installed on the satellite at an angle, characterized in that: include, A safety module, used to put the satellite into safety mode when the output power of the satellite's sailboard toward the sun is less than a predetermined power; The mode switching module is used to obtain the solar angle of the satellite safety mode for a continuous period of time, determine the solar state of the satellite's intermediate inertia axis based on the solar angle, and according to the solar state and in combination with the satellite's control instructions, put the satellite into the solar capture mode or the spin mode. In the spin mode, the solar panels obtain the maximum solar energy; Capture module, used when the satellite enters the sun capture mode; Obtain the main plane solar angle of the satellite, and determine the solar state of the satellite's intermediate inertia axis based on the main plane solar angle; When the middle inertia axis is in the solar range, the satellite's gyroscopic angular velocity and solar angle are obtained, and whether the satellite enters the spinning mode or remains in the solar capture mode is determined according to the gyroscopic angular velocity and solar angle.

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