Satellite attitude control method and device based on mode conversion and satellite

Abstract

The application discloses a satellite attitude control method and device based on mode conversion and a satellite, relates to the field of satellite control, and is characterized in that: a fine control mode, a coarse control mode and a sun capture mode used for controlling the attitude of the satellite and having different attitude determination accuracies are set in advance, the fine control mode is determined as a current target attitude control mode when a first remote control signal is received, the stop control mode is determined to be entered when it is determined that the satellite failure this time is a system out-of-control failure, the attitude of the satellite is regulated according to a second remote control signal when the second remote control signal sent by a ground remote control module is received, and the target attitude control mode is determined again according to the type of the failure carried component when it is determined that the satellite failure this time is a failure of the carried component used for attitude control, so that the satellite is switched to a new target attitude control mode. According to the scheme, the satellite can be switched between different modes according to different states of the satellite, reliable control of the attitude of the satellite is realized, the scheme has high flexibility, strong self-adaptability and is beneficial to practical application.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of satellite control, in particular to a satellite attitude control method and device based on mode conversion and a satellite. BACKGROUND

[0002] Reliable determination and control of satellite attitude is the basis for performing various increasingly complex space missions. Currently, the main way to determine satellite attitude is to use a single attitude sensor, but the attitude control scheme is not combined with the actual running state of the satellite, resulting in a single attitude control mode for the satellite that cannot be flexibly changed to follow the actual running state of the satellite. Therefore, how to provide a solution to the above technical problems is a problem that needs to be solved by those skilled in the art. SUMMARY

[0003] Therefore, the present application provides a satellite attitude control method and device based on mode conversion, which can switch between different modes according to different states of the satellite to reliably control the satellite attitude, with high flexibility and strong adaptability.

[0004] To solve the above technical problems, the present application provides a satellite attitude control method based on mode conversion, comprising:

[0005] Upon receiving a first remote control signal sent by a ground remote control module, indicating that the satellite is normally starting to work, determining that the fine control mode is the current target attitude control mode;

[0006] Determining whether the satellite has failed;

[0007] If so, upon determining that this failure is a system out-of-control failure, determining that the satellite enters the stop control mode, and until a second remote control signal sent by the ground remote control module is received, adjusting the attitude of the satellite according to the second remote control signal;

[0008] Upon determining that this failure is a failure of a satellite-borne component used for attitude control, performing mode conversion and retreat processing according to the type of the failed satellite-borne component to re-determine the target attitude control mode, so that the satellite is converted from the fine control mode to a new target attitude control mode; wherein the new target attitude control mode is one of the coarse control mode and the sun capture mode, and the attitude determination accuracy of the fine control mode > the attitude determination accuracy of the coarse control mode > the attitude determination accuracy of the sun capture mode.

[0009] Further, determining that this failure is a system out-of-control failure comprises:

[0010] When it is determined that the absolute value of the angular velocity of the satellite in any one of the roll axis, the pitch axis and the yaw axis is greater than the first preset angular velocity threshold value, and the duration that the absolute value is greater than the preset safety angular velocity threshold value exceeds the first preset duration, it is determined that the failure is a system out-of-control failure.

[0011] Further, before receiving the first remote control signal sent by the ground remote control module and indicating that the satellite normally starts to work, the method further comprises:

[0012] When it is determined that the satellite has completed the satellite-rocket separation, it is determined that the sun capture mode is the current target attitude control mode, so as to control the attitude of the satellite according to the sun capture mode.

[0013] Further, determining that the failure is a failure of the carried component for attitude control comprises:

[0014] For each carried component for attitude control on the satellite, it is determined whether the accumulated number of times that the carried component appears invalid data is greater than a preset number threshold value corresponding to the carried component;

[0015] If yes, it is determined that the failure is a failure of the carried component for attitude control.

[0016] Further, the attitude of the satellite in the fine control mode is determined according to the star sensor and the gyroscope carried on the satellite;

[0017] The coarse control mode comprises a first coarse control mode and a second coarse control mode, the attitude of the satellite in the first coarse control mode is determined according to the sun sensor and the magnetometer carried on the satellite, and the attitude of the satellite in the second coarse control mode is determined according to the magnetometer and the gyroscope carried on the satellite;

[0018] The attitude of the satellite in the sun capture mode is determined according to the sun sensor.

[0019] Further, the attitude of the satellite is determined according to the sun sensor, comprising:

[0020] The satellite is subjected to rate damping, and it is determined whether the angular velocity of the satellite in each axis is less than a second preset angular velocity threshold value and the duration exceeds a second preset duration;

[0021] If yes, the pitch search is performed by the first sun sensor corresponding to the pitch axis to capture the sun, and the roll search is performed by the second sun sensor corresponding to the roll axis to capture the sun, and then the sun tracking is performed when the first sun sensor and the second sun sensor both capture the sun, so as to determine the attitude of the satellite according to the position of the captured sun.

[0022] Further, when the fault is determined to be a fault of the mounted component for attitude control, mode switching processing is performed according to the type of the fault mounted component to re-determine the target attitude control mode, including:

[0023] When the star sensor is determined to be faulty in the fine control mode, it is determined whether the sun sensor and the magnetometer are both not faulty;

[0024] If yes, the new target attitude control mode is determined to be the first coarse control mode;

[0025] If no, when the magnetometer is determined to be faulty and the sun sensor is not faulty, the new target attitude control mode is determined to be a sun capture mode;

[0026] When the sun sensor is determined to be faulty and the magnetometer is not faulty, the new target attitude control mode is determined to be the second coarse control mode.

[0027] Further, after the new target attitude control mode is determined to be the first coarse control mode, further including:

[0028] When the attitude angle determined in the first coarse control mode is greater than a preset safety attitude angle threshold, and the duration for which the attitude angle is greater than the preset safety attitude angle threshold exceeds a third preset duration, the new target attitude control mode is determined to be the sun capture mode.

[0029] To solve the above technical problem, the application further provides a satellite attitude control device based on mode switching, including:

[0030] A memory for storing a computer program;

[0031] A processor for executing the computer program to realize the steps of the satellite attitude control method based on mode switching as described above.

[0032] To solve the above technical problem, the application further provides a satellite including a star sensor, a gyroscope, a sun sensor, and a magnetometer, and further including the satellite attitude control device based on mode switching as described above;

[0033] The satellite attitude control device based on mode switching is connected with the star sensor, the gyroscope, the sun sensor, and the magnetometer, respectively.

[0034] The application provides a satellite attitude control method and device based on mode conversion and a satellite.

[0035] The above description is only a summary of the technical scheme of the application, in order to more clearly understand the technical means of the application, the application can be implemented according to the content of the specification, and in order to make the above and other purposes, characteristics and advantages of the application more obvious and easy to understand, the following specific embodiments of the application are described. BRIEF DESCRIPTION OF DRAWINGS

[0036] The accompanying drawings used to provide further understanding of the application constitute a part of the application, the schematic embodiments of the application and the description thereof are used to explain the application, and do not constitute an improper limitation on the application. In the drawings:

[0037] Figure 1 A flow chart of a satellite attitude control method based on mode conversion provided by the application;

[0038] Figure 2 A structural schematic diagram of a satellite attitude control device based on mode conversion provided by the application. DETAILED DESCRIPTION

[0039] The core of the application is to provide a satellite attitude control method, device and satellite based on mode conversion, which can convert between different modes according to different states of the satellite to realize reliable control of the satellite attitude, has high flexibility and strong self-adaptability.

[0040] With reference to the drawings of the embodiments of the present application, the technical solutions in the embodiments of the present application will be clearly described. Obviously, the described embodiments are only some of the embodiments of the present application, but not all of the embodiments of the present application. Based on the embodiments of the present application, all other embodiments obtained by those of ordinary skill in the art are within the scope of the present application.

[0041] The terms "first", "second", and the like in the specification and claims of the present application are used to distinguish similar objects, and are not used to describe a specific order or sequence. It should be understood that the data used in this way can be interchanged under appropriate circumstances, so that the embodiments of the present application can be implemented in an order other than that illustrated or described herein, and the objects distinguished by "first", "second", etc. are generally a category, and are not limited to the number of objects, for example, the first object can be one or more. In addition, "and / or" in the specification and claims indicates at least one of the connected objects, and the character " / ", generally indicates that the objects before and after are in a "or" relationship.

[0042] Please refer to Figure 1 , Figure 1 A flow chart of a satellite attitude control method based on mode conversion provided by the present application.

[0043] The satellite attitude control method based on mode conversion comprises:

[0044] S11: When a first remote control signal indicating that the satellite normally starts working is received from a ground remote control module, determining that the fine control mode is the current target attitude control mode;

[0045] S12: Judging whether the satellite has a fault; if yes, entering S13;

[0046] S13: When it is determined that the fault is a system out-of-control fault, determining that the satellite enters the stop control mode, and until a second remote control signal is received from the ground remote control module, adjusting the attitude of the satellite according to the second remote control signal;

[0047] S14: When it is determined that the fault is a fault of a carried component for attitude control, performing mode conversion processing according to the type of the faulted carried component to re-determine the target attitude control mode, so that the satellite is converted from the fine control mode to a new target attitude control mode; wherein the new target attitude control mode is one of the coarse control mode and the sun capture mode, and the attitude determination accuracy of the fine control mode > the attitude determination accuracy of the coarse control mode > the attitude determination accuracy of the sun capture mode.

[0048] In this embodiment, the satellite attitude control method can be applied to a control module on the satellite, the control module is in communication connection with a ground remote control module, and control modes with different attitude determination accuracies are set in advance, including a fine control mode, a coarse control mode and a sun capture mode, so as to adapt different attitude control modes according to the state of the satellite subsequently.

[0049] Specifically, after the satellite completes the separation from the rocket, the ground remote control module sends a first remote control signal for instructing the satellite to start working normally to perform various tasks. In order to ensure the reliability of the attitude control during the task execution, the fine control mode is determined as the current target attitude control mode, so as to control the satellite attitude according to the fine control mode. Then, the monitoring of whether the satellite fails is kept, that is, whether the satellite fails is continuously judged. If not, the control of the satellite attitude according to the fine control mode is kept. If yes, the type of the failure is further determined, including a system out-of-control failure and a failure of a carried component.

[0050] In detail, when the failure is determined as the system out-of-control failure, it is indicated that the satellite has lost control, such as the star body being in a rolling condition. At this time, the attitude cannot be adjusted by various preset modes. At this time, it is determined that the satellite enters a stop control mode to wait for the control instruction of the ground remote control module as the upper monitoring, that is, until the second remote control signal sent by the ground remote control module is received, the attitude of the satellite is adjusted according to the second remote control signal.

[0051] When the failure is determined as the failure of the carried component for attitude control, the mode rollback processing can be performed according to the type of the failed carried component to re-determine the target attitude control mode, so as to make the satellite rollback from the fine control mode with the highest attitude determination accuracy to the coarse control mode or the sun capture mode with relatively low attitude determination accuracy. It should be noted that the carried component can be a star sensor, a sun sensor or a magnetometer, which is not particularly limited here.

[0052] It should be further noted that the three different modes are different in the attitude determination accuracy. After the attitude determination is completed, the control signal for controlling the attitude actuator to act to adjust the actual attitude of the satellite can be output according to the current determined attitude and the target attitude combined with the PID control algorithm. Of course, the control signal can also be filtered by a filter to remove noise and smooth the signal before being transmitted to the attitude actuator, which is not particularly limited here. The attitude actuator can include a magnetic torque device and a flywheel, more specifically, a three-axis magnetic torque device and a zero momentum flywheel, so as to adjust the torque according to the control signal to achieve the purpose of attitude control, which is not particularly limited here.

[0053] Further, it can also be continuously monitored whether the magnetic torque device and the flywheel are faulty. Specifically, when the angular momentum increment of the flywheel per cycle exceeds a preset threshold, such as 0.00125 Nms, the corresponding cumulative value is +1. If it is determined that the angular momentum increment exceeds the threshold for 120 consecutive seconds, it is determined that the flywheel is faulty. At this time, if the remaining flywheels are sufficient to support the control according to the target attitude, the remaining flywheels continue to work; if not, the three-axis magnetic torque device performs control according to the target attitude. As for whether the magnetic torque device is faulty, the monitoring method is as follows: when the magnetic torque device loses power or communication is unsuccessful, the corresponding cumulative value is +1. When the cumulative value is greater than a first preset value, such as 100, the magnetic torque device is actively controlled to lose power and restart. When the cumulative value is greater than a second preset value, such as 240, it is determined that the magnetic torque device is faulty. Since the fault is a single point fault, a prompt signal indicating that the magnetic torque device is faulty and needs to be handled is output, so that the technician can timely know the situation.

[0054] It should be further noted that when it is determined that this time the fault is a system out-of-control fault, an out-of-control prompt signal can be actively fed back to the ground remote control module, so that the ground remote control module can know the situation as soon as possible, thereby facilitating the ground remote control module to feed back a second remote control signal as soon as possible to restore the attitude control of the satellite; when it is determined that this time the fault is a fault of the carried component, the system log can be recorded, or the redundant component corresponding to the faulty carried component on the satellite can be actively switched to replace the function of the faulty carried component. When the corresponding redundant component also has a fault, a system alarm is given to prompt the technician to know as soon as possible.

[0055] In addition, when a third remote control signal representing mode switching sent by the ground remote control module is received, it can be determined that the target attitude control mode is switched to the mode corresponding to the third remote control signal, such as from the fine control mode to the coarse control mode.

[0056] In summary, the satellite attitude control method based on mode conversion provided in the present application can switch between different modes according to different states of the satellite, thereby realizing reliable control of the satellite attitude, having high flexibility, strong self-adaptability, and being beneficial to practical application.

[0057] On the basis of the above-mentioned embodiments:

[0058] In some embodiments, determining that this time the fault is a system out-of-control fault comprises:

[0059] When it is determined that the absolute value of the angular velocity of the satellite in any one of the roll axis, the pitch axis and the yaw axis is greater than a first preset angular velocity threshold, and the duration for which the absolute value is greater than a preset safety angular velocity threshold exceeds a first preset duration, it is determined that this time the fault is a system out-of-control fault.

[0060] Specifically, the first preset angular velocity threshold value can be 10 degrees per second, and the first preset time length can be 250 seconds, so that whether the system out-of-control fault occurs can be reliably determined, and corresponding measures can be taken in time.

[0061] In some embodiments, before receiving the first remote control signal sent by the ground remote control module and indicating that the satellite starts to work normally, the method further comprises:

[0062] When it is determined that the satellite has completed the satellite-rocket separation, the sun capture mode is determined as the current target attitude control mode, so that the attitude control of the satellite is performed according to the sun capture mode.

[0063] In the embodiment, when it is determined that the satellite has completed the satellite-rocket separation, the sun capture mode is determined as the current target attitude control mode, because the first remote control signal has not been received at this time, so as to ensure the safe operation of the satellite.

[0064] In some embodiments, the determination that the fault is a fault of the carried component for attitude control comprises:

[0065] For each carried component for attitude control on the satellite, it is determined whether the accumulated number of times of invalid data of the carried component is greater than a preset number threshold corresponding to the carried component.

[0066] If yes, it is determined that the fault is a fault of the carried component for attitude control.

[0067] Specifically, when the carried component is a star sensor, the corresponding accumulated number of times is increased by 1 each time the star sensor has the determined invalid data, and when it is determined that the accumulated number of times is greater than a preset number threshold corresponding to the star sensor, such as 40, it is determined that the fault is a fault of the carried component for attitude control, and the fault is caused by the star sensor.

[0068] When the carried component is a magnetometer, the corresponding accumulated number of times is increased by 1 each time the magnetometer has the determined invalid data, and when it is determined that the accumulated number of times is greater than a preset number threshold corresponding to the magnetometer, such as 400, it is determined that the fault is a fault of the carried component for attitude control, and the fault is caused by the magnetometer.

[0069] When the carried component is a sun sensor, the corresponding accumulated number of times is increased by 1 each time any one of a first sun sensor corresponding to a pitch axis and a second sun sensor corresponding to a roll axis has the determined data anomaly that the sun is invisible, and when it is determined that the accumulated number of times is greater than a preset number threshold corresponding to the sun sensor, such as 9000, it is determined that the fault is a fault of the carried component for attitude control, and the fault is caused by the sun sensor. The reason why 9000 is set here is that 9000 times correspond to about 2250 seconds, which exceeds the time length of the earth shadow.

[0070] When the mounted component is a gyroscope, the corresponding cumulative number is incremented by 1 each time the gyroscope output is always zero or a certain data is invalid. When it is determined that the cumulative number is greater than a first preset number threshold corresponding to the gyroscope, such as 40, the gyroscope is actively controlled to be powered off and restarted. After restarting, the cumulative number of faults is continued to be accumulated. When it is determined that the cumulative number is greater than a second preset number threshold corresponding to the gyroscope, it is determined that the fault this time is a fault of the mounted component for attitude control, and the fault mounted component is the gyroscope. It should be further pointed out that when the current target attitude control mode is the fine control mode, the new target attitude control mode can be switched back to the first coarse control mode. When the current target attitude control mode is the second coarse control mode, the new target attitude control mode can be switched back to the sun capture mode. In addition, in the fine control mode, the measurement result of the gyroscope can also be replaced by star sensor data difference to realize the determination of the satellite attitude, which is not particularly limited here.

[0071] It can be seen that the above-mentioned mode can reliably determine whether the fault this time is a fault of the mounted component for attitude control.

[0072] In some embodiments, in the fine control mode, the attitude of the satellite is determined according to the star sensor and the gyroscope mounted on the satellite.

[0073] The coarse control mode includes a first coarse control mode and a second coarse control mode. In the first coarse control mode, the attitude of the satellite is determined according to the sun sensor and the magnetometer mounted on the satellite. In the second coarse control mode, the attitude of the satellite is determined according to the magnetometer and the gyroscope mounted on the satellite.

[0074] In the sun capture mode, the attitude of the satellite is determined according to the sun sensor.

[0075] In the present embodiment, the attitude determination modes are different in different modes. The gyroscope herein can be a three-axis gyroscope, and the magnetometer can be a three-axis magnetometer.

[0076] More specifically, in the fine control mode, the attitude determination is realized according to the star sensor and the gyroscope, that is, the position of the star is observed by the star sensor and the three-axis attitude of the satellite is determined in combination with the output of the gyroscope. The measurement results of the star sensor and the gyroscope are fused through a preset attitude determination algorithm, which is not particularly limited here. Through such a combination, the accuracy of attitude determination is improved.

[0077] The coarse control mode includes a first coarse control mode and a second coarse control mode. The first coarse control mode is based on a sun sensor and a magnetometer to realize attitude determination, i.e., based on a dual-vector attitude determination algorithm combining a sun vector determined by the sun sensor and a geomagnetic vector determined by the magnetometer to realize coarse attitude determination. The second coarse control mode is based on the magnetometer and a gyroscope to realize attitude determination, i.e., fusing the geomagnetic vector determined by the magnetometer and an angular velocity vector determined by the gyroscope to realize coarse attitude determination.

[0078] It can be seen that the above-mentioned mode can reliably switch between different modes according to different states of the satellite, and the adaptability is strong.

[0079] In some embodiments, the determination of the attitude of the satellite according to the sun sensor includes:

[0080] The satellite is rate-damped, and it is determined whether the angular velocities of the satellite on each axis are less than a second preset angular velocity threshold and the duration exceeds a second preset duration.

[0081] If yes, the pitch search is performed by the first sun sensor corresponding to the pitch axis to capture the sun, and the roll search is performed by the second sun sensor corresponding to the roll axis to capture the sun, and then the sun tracking is performed when the first sun sensor and the second sun sensor both capture the sun, so as to determine the attitude of the satellite according to the position of the captured sun.

[0082] In the embodiment, the sun capture mode can be understood as a basic attitude control mode for ensuring safe operation of the satellite. Specifically, the second preset angular velocity can be 0.5 degrees per second, and the second preset duration can be set according to the actual application, which is not particularly limited here.

[0083] In some embodiments, when it is determined that the failure is a failure of the carried component for attitude control, the mode rollback processing is performed according to the type of the failed carried component to redetermine the target attitude control mode, including:

[0084] When it is determined that the star sensor fails in the fine control mode, it is determined whether the sun sensor and the magnetometer both fail.

[0085] If yes, it is determined that the new target attitude control mode is the first coarse control mode.

[0086] If no, when it is determined that the magnetometer fails and the sun sensor does not fail, it is determined that the new target attitude control mode is the sun capture mode.

[0087] When it is determined that the sun sensor fails and the magnetometer does not fail, it is determined that the new target attitude control mode is the second coarse control mode.

[0088] It should be further explained that, in the fine control mode, if the gyroscope fails, it is determined whether the sun sensor and the magnetometer both fail, if yes, the new target attitude control mode is determined as the first coarse control mode, if no, when it is determined that the magnetometer fails and the sun sensor does not fail, the new target attitude control mode is determined as the sun capture mode. When the sun sensor also fails, the star sensor can be used to determine the satellite attitude alone.

[0089] It can be understood that, the new target attitude control mode is determined as the second coarse control mode, which needs to ensure that the gyroscope and the magnetometer both do not fail, and if an extreme case occurs, that is, the star sensor, the gyroscope, the sun sensor and the magnetometer all fail, at this time, a full failure signal can be fed back to the ground remote control module to wait for the fourth remote control signal sent by the ground remote control module, and then the attitude control of the satellite is realized according to the fourth remote control signal.

[0090] In some embodiments, after the new target attitude control mode is determined as the first coarse control mode, the method further comprises:

[0091] When it is determined that the attitude angle determined in the first coarse control mode is greater than a preset safe attitude angle threshold, and the duration that the attitude angle is greater than the preset safe attitude angle threshold exceeds a third preset duration, the new target attitude control mode is determined as the sun capture mode.

[0092] In this embodiment, it is further considered that there may be an attitude angle out-of-tolerance situation in the first coarse control mode, at this time, the new target attitude control mode can be determined as the sun capture mode to ensure reliable control of the satellite attitude, and specifically, the preset safe attitude angle threshold and the third preset duration can be set according to the application actual situation, which is not particularly limited here.

[0093] Please refer to Figure 2 , Figure 2 A structure schematic diagram of a satellite attitude control device based on mode conversion provided by the present application.

[0094] The satellite attitude control device based on mode conversion comprises:

[0095] The memory 21 is used to store a computer program.

[0096] The processor 22 is used to execute the computer program to realize the steps of the satellite attitude control method based on mode conversion as described above.

[0097] For the satellite attitude control device based on mode conversion provided in the present application, please refer to the above embodiments of the satellite attitude control method based on mode conversion, which will not be repeated here.

[0098] The application further provides a satellite, comprising a star sensor, a gyroscope, a sun sensor and a magnetometer, and further comprising the satellite attitude control device based on mode conversion as described above.

[0099] The satellite attitude control device based on mode conversion is connected with the star sensor, the gyroscope, the sun sensor and the magnetometer respectively.

[0100] For the satellite provided in the application, refer to the above-mentioned embodiments of the satellite attitude control method based on mode conversion, which will not be repeated here.

[0101] The embodiments in the specification are described in a progressive manner, and each embodiment focuses on the difference from other embodiments, and the same or similar parts between the embodiments can be referred to each other. For the device disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple, and the related parts can be referred to the method part. The relationship 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 the entities or operations. Moreover, the term "include", "contain" or any other variant thereof is intended to cover non-exclusive inclusion, so that the process, method, article or equipment including a series of elements not only includes those elements, but also includes other elements not explicitly listed or inherent to such process, method, article or equipment. Without more limitations, the element defined by the statement "including one" does not exclude the presence of another identical element in the process, method, article or equipment including the element.

[0102] The above description of the disclosed embodiments enables a person skilled in the art to implement or use the application. Various modifications to the embodiments will be apparent to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the application. Therefore, the application will not be limited to the embodiments shown herein, but will conform to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A method for satellite attitude control based on mode switching, characterized in that, Comprise: When receiving the first remote control signal sent by the ground remote control module, which represents that the satellite starts to work normally, determine that the fine control mode is the current target attitude control mode; Determine whether the satellite has failed; If yes, when determining that the failure is a system out-of-control failure, determine that the satellite enters the stop control mode until receiving the second remote control signal sent by the ground remote control module, and then control the attitude of the satellite according to the second remote control signal; When determining that the failure is a failure of the carried component for attitude control, perform mode rollback processing according to the type of the failed carried component to re-determine the target attitude control mode, so that the satellite is switched from the fine control mode to a new target attitude control mode; wherein the new target attitude control mode is one of the coarse control mode and the sun capture mode, and the attitude determination accuracy of the fine control mode > the attitude determination accuracy of the coarse control mode > the attitude determination accuracy of the sun capture mode; In the fine control mode, the attitude of the satellite is determined according to the star sensor and the gyroscope carried on the satellite; The coarse control mode includes a first coarse control mode and a second coarse control mode, and in the first coarse control mode, the attitude of the satellite is determined according to the sun sensor and the magnetometer carried on the satellite; in the second coarse control mode, the attitude of the satellite is determined according to the magnetometer and the gyroscope carried on the satellite; In the sun capture mode, the attitude of the satellite is determined according to the sun sensor; When determining that the failure is a failure of the carried component for attitude control, performing mode rollback processing according to the type of the failed carried component to re-determine the target attitude control mode, comprising: When determining that the star sensor fails in the fine control mode, determine whether the sun sensor and the magnetometer both fail; If yes, determine that the new target attitude control mode is the first coarse control mode; If not, when determining that the magnetometer fails and the sun sensor does not fail, determine that the new target attitude control mode is the sun capture mode; When determining that the sun sensor fails and the magnetometer does not fail, determine that the new target attitude control mode is the second coarse control mode.

2. The method of claim 1, wherein the satellite attitude control method based on mode transition is characterized by, Determine that the failure is a system out-of-control failure, comprising: When determining that the absolute value of the angular velocity of the satellite in any one of the roll axis, the pitch axis and the yaw axis is greater than the first preset angular velocity threshold, and the duration that the absolute value is greater than the preset safety angular velocity threshold exceeds the first preset time, determine that the failure is a system out-of-control failure.

3. The method of claim 1, wherein the satellite attitude control method based on mode transition is characterized by, Before receiving the first remote control signal sent by the ground remote control module, which represents that the satellite starts to work normally, also comprising: When determining that the satellite has completed the satellite-rocket separation, determine that the sun capture mode is the current target attitude control mode, so as to control the attitude of the satellite according to the sun capture mode.

4. The method of claim 1, wherein the satellite attitude control method based on mode transition is characterized by, Determine that the failure is a failure of the carried component for attitude control, comprising: For each of the satellite-mounted components for attitude control, it is determined whether the cumulative number of times of invalid data of the mounted component is greater than a preset number threshold corresponding to the mounted component; If yes, it is determined that the current fault is a mounted component fault for attitude control.

5. The method of claim 1, wherein the satellite attitude control method based on mode transition is characterized by, The determination of the satellite attitude according to the sun sensor includes: The satellite is rate-damped, and it is determined whether the angular velocity of the satellite on each axis is less than a second preset angular velocity threshold and lasts for more than a second preset time length; If yes, the first sun sensor corresponding to the pitch axis is used for pitch search to capture the sun, and the second sun sensor corresponding to the roll axis is used for roll search to capture the sun, and then the sun is tracked when the first sun sensor and the second sun sensor both capture the sun, so as to determine the satellite attitude according to the position of the captured sun.

6. The method of claim 1, wherein the satellite attitude control method based on mode transition is characterized by, After determining the new target attitude control mode as the first coarse control mode, it further includes: When it is determined that the attitude angle determined in the first coarse control mode is greater than a preset safe attitude angle threshold, and the duration that the attitude angle is greater than the preset safe attitude angle threshold exceeds a third preset time length, the new target attitude control mode is determined as the sun capture mode.

7. A mode-switching-based satellite attitude control apparatus characterized by comprising: It includes: a memory for storing a computer program; a processor for executing the computer program to realize the steps of the satellite attitude control method based on mode conversion according to any one of claims 1 to 6.

8. A satellite, characterized by It includes a star sensor, a gyroscope, a sun sensor and a magnetometer, and further includes the satellite attitude control device based on mode conversion according to claim 7; The satellite attitude control device based on mode conversion is connected with the star sensor, the gyroscope, the sun sensor and the magnetometer respectively.

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