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

Abstract

The invention discloses a satellite attitude control method and device based on mode conversion and a satellite, and relates to the field of satellite control. A precise control mode, a coarse control mode and a sun capture mode which are used for controlling the attitude of the satellite and have different attitude determination precisions are preset; when the first remote control signal is received, determining a precise control mode as a current target attitude control mode; when it is determined that the fault of the satellite is a system out-of-control fault, determining to enter a stop control mode, and regulating and controlling the attitude of the satellite according to a second remote control signal when the second remote control signal sent by the ground remote control module is received; and when it is determined that the fault of the satellite is the fault of the carrying component for attitude control, re-determining the target attitude control mode according to the type of the faulted carrying component, so that the satellite is switched to a new target attitude control mode. According to the scheme, conversion between different modes can be achieved according to different states of the satellite, reliable control over the satellite attitude is achieved, the flexibility is high, the self-adaptive capacity is high, and practical application is facilitated.

Description

Technical Field

[0001] The present invention relates to the field of satellite control technology, and in particular to a satellite attitude control method, device and satellite based on mode conversion. Background Art

[0002] Reliable determination and control of satellite attitude is the basis for executing various increasingly complex space missions. At present, the method of determining satellite attitude is mainly achieved by using a single attitude sensor, but it does not consider combining the attitude control scheme with the actual operating status of the satellite. As a result, the attitude control method for the satellite is relatively single and cannot be flexibly changed according to the actual operating conditions of the satellite. Therefore, how to provide a solution to the above technical problems is a problem that technicians in this field urgently need to solve. Summary of the invention

[0003] In view of this, the present invention provides a satellite attitude control method, device and satellite based on mode conversion. The scheme can convert between different modes according to different states of the satellite to achieve reliable control of the satellite attitude, with high flexibility and strong adaptability.

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

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

[0006] Determining whether the satellite fails;

[0007] If yes, when it is determined that the fault is a system out-of-control fault, the satellite is determined to enter a stop control mode until a second remote control signal sent by the ground remote control module is received, and the attitude of the satellite is controlled according to the second remote control signal;

[0008] When it is determined that the fault is a fault of an onboard component used for attitude control, mode transition processing is performed according to the type of the onboard component that has failed to redefine the target attitude control mode, so that the satellite switches from the fine control mode to a new target attitude control mode; wherein the new target attitude control mode is one of a coarse control mode and a sun capture mode, and the attitude determination accuracy of the fine control mode is greater than the attitude determination accuracy of the coarse control mode and greater than the attitude determination accuracy of the sun capture mode.

[0009] Furthermore, it was determined that the fault was a system out-of-control fault, including:

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

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

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

[0013] Furthermore, it was determined that the fault was a fault of the onboard components used for attitude control, including:

[0014] For each onboard component used for attitude control on the satellite, determining whether the cumulative number of invalid data occurrences of the onboard component is greater than a corresponding preset number threshold;

[0015] If so, it is determined that the fault is a failure of the onboard components used for attitude control.

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

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

[0018] In the sun capture mode, the attitude of the satellite is determined based on the sun sensor.

[0019] Further, determining the attitude of the satellite according to the sun sensor includes:

[0020] Performing rate damping on the satellite, and determining 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;

[0021] If so, a pitch search is performed through the first solar sensor corresponding to the pitch axis to capture the sun, and a roll search is performed through the second solar sensor corresponding to the roll axis to capture the sun, and then a solar cruise is performed when both the first solar sensor and the second solar sensor capture the sun to determine the attitude of the satellite based on the captured position of the sun.

[0022] Further, when it is determined that the fault is a fault of an onboard component used for attitude control, a mode transition process is performed according to the type of the onboard component that has failed to re-determine the target attitude control mode, including:

[0023] When it is determined that the star sensor fails in the precise control mode, determining whether both the sun sensor and the magnetometer are not failed;

[0024] If so, determining that the new target attitude control mode is the first coarse control mode;

[0025] If not, when it is determined that the magnetometer fails and the sun sensor does not fail, determining that the new target attitude control mode is a sun capture mode;

[0026] When it is determined that the sun sensor fails and the magnetometer does not fail, the new target attitude control mode is determined to be the second coarse control mode.

[0027] Further, after determining that the new target posture control mode is the first coarse control mode, the method further includes:

[0028] When it is determined that 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] In order to solve the above technical problems, the present invention also provides a satellite attitude control device based on mode conversion, comprising:

[0030] Memory for storing computer programs;

[0031] A processor is used to implement the steps of the satellite attitude control method based on mode conversion as described above when executing the computer program.

[0032] In order to solve the above technical problems, the present invention also provides a satellite, comprising a star sensor, a gyroscope, a sun sensor and a magnetometer, and also comprising the satellite attitude control device based on mode conversion as described above;

[0033] The mode conversion-based satellite attitude control device is connected to the star sensor, the gyroscope, the sun sensor and the magnetometer respectively.

[0034] The present application provides a satellite attitude control method, device and satellite based on mode conversion, which pre-sets a fine control mode, a coarse control mode and a sun capture mode for controlling the attitude of the satellite and having different attitude determination accuracy. When receiving the first remote control signal sent by the ground remote control module indicating that the satellite starts working normally, the fine control mode is determined to be the current target attitude control mode; it is determined whether the satellite fails. If so, when it is determined that the failure is a system out-of-control failure, it means that the satellite has lost control and cannot adjust the attitude through various preset modes, then the satellite is determined to enter the stop control mode to wait for the control instruction of the ground remote control module, 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; when it is determined that the failure is a failure of the onboard component used for attitude control, the mode conversion and decommissioning processing is performed according to the type of the onboard component that fails to re-determine the target attitude control mode, so that the satellite is transferred from the fine control mode to the new target attitude control mode. It can be seen that the scheme can switch between different modes according to different states of the satellite to achieve reliable control of the satellite attitude, with high flexibility and strong adaptability, which is conducive to practical application.

[0035] The above description is only an overview of the technical solution of the present application. In order to more clearly understand the technical means of the present application, it can be implemented in accordance with the contents of the specification. In order to make the above and other purposes, features and advantages of the present application more obvious and easy to understand, the specific implementation methods of the present application are listed below. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

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

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

[0039] The core of the present invention is to provide a satellite attitude control method, device and satellite based on mode conversion. The scheme can convert between different modes according to different states of the satellite to achieve reliable control of the satellite attitude, with high flexibility and strong adaptability.

[0040] The following will be combined with the drawings in the embodiments of the present application to clearly describe the technical solutions in the embodiments of the present application. Obviously, the described embodiments are part of the embodiments of the present application, rather than all the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments in the present application belong to the scope of protection of this application.

[0041] The terms "first", "second", etc. 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 interchangeable under appropriate circumstances, so that the embodiments of the present application can be implemented in an order other than those illustrated or described here, and the objects distinguished by "first", "second", etc. are generally of one type, and the number of objects is not limited. For example, the first object can be one or more. In addition, "and / or" in the specification and claims represents at least one of the connected objects, and the character " / " generally indicates that the objects associated with each other are in an "or" relationship.

[0042] Please refer to Figure 1 , Figure 1 The present invention provides a flow chart of a satellite attitude control method based on mode conversion.

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

[0044] S11: upon receiving a first remote control signal sent by a ground remote control module indicating that the satellite has started to work normally, determining that the precise control mode is the current target attitude control mode;

[0045] S12: Determine whether the satellite fails; if so, proceed to S13;

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

[0047] S14: When it is determined that the fault is a fault of an onboard component used for attitude control, a mode transition process is performed according to the type of the onboard component that has failed to redefine the target attitude control mode, so that the satellite switches 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 is greater than the attitude determination accuracy of the coarse control mode and the attitude determination accuracy of the sun capture mode.

[0048] In this embodiment, the satellite attitude control method can be applied to the control module on the satellite. The control module is communicated with the ground remote control module and control modes with different attitude determination accuracies are pre-set, namely, fine control mode, coarse control mode and sun capture mode, so that different attitude control modes can be adapted subsequently according to the status of the satellite.

[0049] Specifically, after the satellite completes the separation of the rocket and the satellite, the ground remote control module sends a first remote control signal to instruct the satellite to start working normally to perform various tasks. In order to ensure the reliability of attitude control during task execution, the precision control mode is determined to be the current target attitude control mode, so that the satellite attitude can be controlled according to the precision control mode; then, the satellite is monitored to see if a fault occurs, that is, it is continuously determined whether the satellite has a fault. If not, the satellite attitude can continue to be controlled according to the precision control mode; if so, the type of fault is further determined, and the types of faults include system out-of-control faults and onboard component faults.

[0050] To elaborate, when it is determined that the fault is a system out-of-control fault, it means that the satellite has lost control, such as the planet may be in a tumbling situation. At this time, the attitude can no longer be adjusted through various preset modes. At this time, the satellite is determined to enter the stop control mode to wait for the control instruction of the ground remote control module as the upper-level monitoring, that is, until the second remote control signal sent by the ground remote control module is received, the satellite's attitude is adjusted according to the second remote control signal.

[0051] When it is determined that the fault is a fault of an onboard component used for attitude control, a mode transition process can be performed according to the type of the onboard component that has failed to re-determine the target attitude control mode, thereby causing the satellite to transition 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 onboard component here can be a star sensor, a sun sensor, a magnetometer, etc., and is not particularly limited here.

[0052] It should also be noted that the difference between the three different modes lies in the different accuracy of attitude determination. After the attitude determination is completed, all of them can be combined with the PID control algorithm to output a control signal for controlling the action of the attitude actuator to adjust the actual attitude of the satellite based on the currently determined attitude and the target attitude. Of course, the control signal can also be filtered through a filter to remove noise and smooth the signal before being transmitted to the attitude actuator, which is not specifically limited here; the attitude actuator here can include a magnetic torquer and a flywheel, and more specifically a three-axis magnetic torquer and a zero-momentum flywheel, so as to achieve the purpose of attitude control by adjusting the torque according to the control signal, which is not specifically limited here.

[0053] Furthermore, it is also possible to continuously monitor whether the magnetic torquer and the flywheel have failed. Specifically, when the angular momentum increment of the flywheel per cycle exceeds a preset threshold, such as 0.00125Nms, the corresponding cumulative value is set to +1. If it is determined that the angular momentum increment exceeds the threshold for 120 consecutive seconds, it is determined that the flywheel has failed. At this time, if the remaining flywheel is sufficient to support the control according to the target posture, the remaining flywheel continues to work; if it is not sufficient to support, the three-axis magnetic torquer will perform control according to the target posture. The monitoring method for whether the magnetic torquer has failed is: when the magnetic torquer loses power or communication is unsuccessful, the corresponding cumulative value is set to +1. When the cumulative value is greater than the first preset value, such as 100, the magnetic torquer is actively controlled to lose power and restart. When the cumulative value is greater than the second preset value, such as 240, it is determined that the magnetic torquer has failed. Since the fault is a single-point fault, a prompt signal indicating that the magnetic torquer has failed to be processed is output, so that the technical personnel can grasp the situation in time.

[0054] It should also be noted that when it is determined that 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 be aware of 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 the fault is a failure of an onboard component, the fault can also be recorded in the form of a system log, or the redundant component corresponding to the faulty onboard component on the satellite can be actively switched to take over the function of the faulty onboard component. When the corresponding redundant component also fails, a system alarm is issued to promptly alert the technical staff.

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

[0056] In summary, the present application provides a satellite attitude control method based on mode conversion, which can convert between different modes according to the different states of the satellite to achieve reliable control of the satellite attitude, with high flexibility and strong adaptability, which is conducive to practical application.

[0057] Based on the above embodiments:

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

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

[0060] Specifically, the first preset angular velocity threshold here may be 10 degrees / second, and the first preset time length may be 250 seconds. Thus, the above method can be used to reliably determine whether a system out-of-control failure occurs, so that corresponding measures can be taken in a timely manner.

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

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

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

[0064] In some embodiments, determining that the failure is a failure of an onboard component for attitude control includes:

[0065] For each onboard component used for attitude control on the satellite, determine whether the cumulative number of invalid data occurrences of the onboard component is greater than a corresponding preset number threshold;

[0066] If so, it is determined that the fault is a failure of the onboard components used for attitude control.

[0067] Specifically, when the onboard component is a star sensor, each time the determined data of the star sensor is invalid, the corresponding cumulative number is increased by 1. When it is determined that the cumulative number 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 onboard component for attitude control, and the faulty onboard component is a star sensor.

[0068] When the onboard component is a magnetometer, each time the magnetometer shows invalid data, the corresponding cumulative number is increased by 1. When it is determined that the cumulative number is greater than the preset number threshold corresponding to the magnetometer, such as 400, it is determined that the fault is a fault of the onboard component used for attitude control, and the faulty onboard component is the magnetometer.

[0069] When the onboard component is a sun sensor, each time any one of the first sun sensor corresponding to the pitch axis and the second sun sensor corresponding to the roll axis has data abnormality that the sun is invisible, the corresponding cumulative number is increased by 1. When it is determined that the cumulative number is greater than the preset number threshold corresponding to the sun sensor, such as 9000, it is determined that the fault is a fault of the onboard component used for attitude control, and the faulty onboard component is the sun sensor. The reason why it is set to 9000 here is that 9000 times roughly corresponds to 2250 seconds, which has exceeded the length of the earth's shadow.

[0070] When the onboard component is a gyroscope, each time the gyroscope outputs a constant zero or certain data is invalid, the corresponding cumulative number is increased by 1. 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 number of faults continues 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 is a fault of the onboard component used for attitude control, and the faulty onboard component is the gyroscope; it should also be noted that when the current target attitude control mode is the fine control mode, the new target attitude control mode can be transferred to the first coarse control mode, and when the current target attitude control mode is the second coarse control mode, the new target attitude control mode can be transferred to the sun capture mode; in addition, in the fine control mode, the star-sensitive data differential can be used to replace the measurement result of the gyroscope to determine the satellite attitude, which is not particularly limited here.

[0071] It can be seen that the above method can reliably determine whether the fault is a fault of the onboard components used for attitude control.

[0072] In some embodiments, in the fine control mode, the attitude of the satellite is determined based on a star sensor and a gyroscope carried 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 magnetometer carried on the satellite; in the second coarse control mode, the attitude of the satellite is determined according to the magnetometer and gyroscope carried on the satellite.

[0074] In the sun capture mode, the satellite's attitude is determined based on the sun sensor.

[0075] In this embodiment, the manners for determining the posture are different in different modes. The gyroscope here may specifically be a three-axis gyroscope, and the magnetometer may specifically be a three-axis magnetometer.

[0076] More specifically, in the precise control mode, attitude determination is achieved based on the star sensor and the gyroscope, that is, the three-axis attitude determination of the satellite is achieved by observing the positions of the stars through the star sensor and combining the output of the gyroscope, and the measurement results of the star sensor and the gyroscope are integrated through a preset attitude determination algorithm. No special limitation is made to the preset attitude determination algorithm here. This combination setting is conducive to improving the accuracy of attitude determination.

[0077] The coarse control mode includes a first coarse control mode and a second coarse control mode. The first coarse control mode realizes attitude determination based on a sun sensor and a magnetometer, that is, a dual-vector attitude determination algorithm is used to combine the sun vector determined by the sun sensor and the geomagnetic vector determined by the magnetometer to realize coarse attitude determination; the second coarse control mode realizes attitude determination based on a magnetometer and a gyroscope, that is, the geomagnetic vector determined by the magnetometer is integrated with the angular velocity vector determined by the gyroscope to realize coarse attitude determination.

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

[0079] In some embodiments, determining the attitude of a satellite based on a sun sensor includes:

[0080] Performing rate damping on the satellite, and determining 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;

[0081] If so, a pitch search is performed through the first solar sensor corresponding to the pitch axis to capture the sun, and a roll search is performed through the second solar sensor corresponding to the roll axis to capture the sun, and then a solar cruise is performed when both the first solar sensor and the second solar sensor capture the sun to determine the satellite's attitude based on the captured position of the sun.

[0082] In this embodiment, the sun capture mode can be understood as a basic attitude control mode to ensure the safe operation of the satellite. Specifically, the second preset angular velocity here can be 0.5 degrees / second, and the second preset time duration can be set according to the actual application, and no special limitation is made here.

[0083] In some embodiments, when it is determined that the fault is a fault of an onboard component used for attitude control, a mode transition process is performed according to the type of the onboard component that has failed to re-determine the target attitude control mode, including:

[0084] When the star sensor is judged to be faulty in the precision control mode, determine whether the sun sensor and magnetometer are both not faulty;

[0085] If so, determining the new target attitude control mode as the first coarse control mode;

[0086] If not, when it is determined that the magnetometer fails and the sun sensor does not fail, determining the new target attitude control mode to be the sun capture mode;

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

[0088] It should also be noted that in the fine control mode, if the gyroscope fails, it is determined whether the sun sensor and the magnetometer are both fine. If so, the new target attitude control mode is determined to be the first coarse control mode; if not, when it is determined that the magnetometer fails and the sun sensor does not fail, the new target attitude control mode is determined to be the sun capture mode. When the sun sensor also fails, the star sensor can be used to determine the satellite attitude alone.

[0089] It is understandable that determining the new target attitude control mode as the second coarse control mode requires ensuring that both the gyroscope and the magnetometer are not faulty. If an extreme situation occurs, that is, when the star sensor, gyroscope, sun sensor and magnetometer all fail, a full fault 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 achieved according to the fourth remote control signal.

[0090] In some embodiments, after determining that the new target posture control mode is the first coarse control mode, the method further includes:

[0091] When it is determined that the attitude angle determined in the first coarse control mode is greater than the 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.

[0092] In this embodiment, it is further taken into consideration that there may be an attitude angle error 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. Specifically, the preset safety attitude angle threshold and the third preset time length here can be set according to the actual application, and no special limitation is made here.

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

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

[0095] A memory 21, used for storing computer programs;

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

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

[0098] The present invention also provides a satellite, comprising a star sensor, a gyroscope, a sun sensor and a magnetometer, and also comprising the satellite attitude control device based on mode conversion as described above;

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

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

[0101] In this specification, each embodiment is described in a progressive manner, and each embodiment focuses on the differences from other embodiments, and the same and similar parts between the embodiments can be referred to each other. For the device disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant parts refer to the method part description. Relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply that there is any such actual relationship or order between these entities or operations. Moreover, the term "include", "comprise" 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 that are not clearly listed, or also includes elements inherent to such process, method, article or equipment. In the absence of more restrictions, the elements limited by the sentence "including one" do not exclude the existence of other identical elements in the process, method, article or equipment including the elements.

[0102] The above description of the disclosed embodiments enables those skilled in the art to implement or use the present invention. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention 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 satellite attitude control method based on mode conversion, characterized in that: include: When receiving a first remote control signal sent by a ground remote control module indicating that the satellite has started to work normally, determining that the precise control mode is the current target attitude control mode; Determining whether the satellite fails; If yes, when it is determined that the fault is a system out-of-control fault, the satellite is determined to enter a stop control mode until a second remote control signal sent by the ground remote control module is received, and the attitude of the satellite is controlled according to the second remote control signal; When it is determined that the fault is a fault of an onboard component used for attitude control, mode transition processing is performed according to the type of the onboard component that has failed to redefine the target attitude control mode, so that the satellite switches from the fine control mode to a new target attitude control mode; wherein the new target attitude control mode is one of a coarse control mode and a sun capture mode, and the attitude determination accuracy of the fine control mode is greater than the attitude determination accuracy of the coarse control mode and greater than the attitude determination accuracy of the sun capture mode.

2. The satellite attitude control method based on mode conversion as claimed in claim 1, characterized in that: It is determined that this fault is a system out-of-control fault, including: When it is determined that the absolute value of the angular velocity of the satellite on any one of the roll axis, pitch axis and yaw axis is greater than the first preset angular velocity threshold, and the duration for which the absolute value is greater than the preset safety angular velocity threshold exceeds the first preset duration, the fault is determined to be a system out-of-control fault.

3. The satellite attitude control method based on mode conversion as claimed in claim 1, characterized in that: Before receiving the first remote control signal sent by the ground remote control module indicating that the satellite starts to work normally, the method further includes: When it is determined that the satellite has completed the separation of the satellite and the rocket, the solar capture mode is determined as the current target attitude control mode, so that the attitude control of the satellite is performed according to the solar capture mode.

4. The satellite attitude control method based on mode conversion as claimed in claim 1, characterized in that: The fault was determined to be a fault of the onboard components used for attitude control, including: For each onboard component used for attitude control on the satellite, determining whether the cumulative number of invalid data occurrences of the onboard component is greater than a corresponding preset number threshold; If so, it is determined that the fault is a failure of the onboard components used for attitude control.

5. The satellite attitude control method based on mode conversion according to any one of claims 1 to 4, characterized in that: In the precise control mode, the attitude of the satellite is determined according to a star sensor and a gyroscope carried on the satellite; 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 a sun sensor and a magnetometer carried on the satellite; in the second coarse control mode, the attitude of the satellite is determined according to a magnetometer and a gyroscope carried on the satellite. In the sun capture mode, the attitude of the satellite is determined based on the sun sensor.

6. The satellite attitude control method based on mode conversion as claimed in claim 5, characterized in that: Determining the attitude of the satellite according to the sun sensor includes: Performing rate damping on the satellite, and determining 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; If so, a pitch search is performed through the first solar sensor corresponding to the pitch axis to capture the sun, and a roll search is performed through the second solar sensor corresponding to the roll axis to capture the sun, and then a solar cruise is performed when both the first solar sensor and the second solar sensor capture the sun to determine the attitude of the satellite based on the captured position of the sun.

7. The satellite attitude control method based on mode conversion as claimed in claim 5, characterized in that: When it is determined that the fault is a fault of an onboard component used for attitude control, a mode transition process is performed according to the type of the onboard component that has failed to re-determine the target attitude control mode, including: When it is determined that the star sensor fails in the precise control mode, determining whether both the sun sensor and the magnetometer are not failed; If so, determining that the new target attitude control mode is the first coarse control mode; If not, when it is determined that the magnetometer fails and the sun sensor does not fail, determining that the new target attitude control mode is a sun capture mode; When it is determined that the sun sensor fails and the magnetometer does not fail, the new target attitude control mode is determined to be the second coarse control mode.

8. The satellite attitude control method based on mode conversion as claimed in claim 7, characterized in that: After determining that the new target attitude control mode is the first coarse control mode, the method further includes: When it is determined that 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.

9. A satellite attitude control device based on mode conversion, characterized in that: include: Memory for storing computer programs; A processor is used to implement the steps of the satellite attitude control method based on mode conversion as described in any one of claims 1 to 8 when executing the computer program.

10. A satellite, characterized in that: The invention comprises a star sensor, a gyroscope, a sun sensor and a magnetometer, and further comprises a satellite attitude control device based on mode conversion as claimed in claim 9; The mode conversion-based satellite attitude control device is connected to the star sensor, the gyroscope, the sun sensor and the magnetometer respectively.

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