A method for switching the attitude control mode of a cube satellite

Abstract

This invention discloses a method for switching attitude control modes for a CubeSat, specifically relating to the field of satellite attitude control technology. The steps are as follows: After separation of the satellite from the launch vehicle, the attitude control system is powered on, the pitch axis momentum wheel begins to rotate, and the satellite performs bias momentum angular rate damping. When the orbit is deemed valid, the filter is activated. When the filter converges and attitude is acquired, bias momentum control is performed; if a fault occurs, the system switches to angular rate damping mode. Finally, a command is injected to switch to zero momentum control mode. When the filter converges, zero momentum control is performed; if a fault occurs, the system switches to angular rate damping mode. When one momentum wheel fails, the other three wheels form either bias momentum control or zero momentum control mode to complete the predetermined attitude control task. Therefore, the above attitude control mode switching method makes full use of limited hardware resources, has clear mode switching logic, and is simple to implement in software, effectively improving the reliability of the CubeSat attitude control system.

Description

Technical Field

[0001] This invention belongs to the field of satellite attitude control technology, specifically relating to a method for switching attitude control modes for a cubic satellite. Background Technology

[0002] With the continuous advancement of aerospace technology, satellite applications have evolved from early space science experiments or on-orbit demonstrations to increasingly playing a vital role in people's livelihoods, national defense, and scientific research, such as Earth remote sensing, navigation, deep space exploration, communication relay, space situational awareness, and aerospace engineering education. Compared to large satellites, CubeSats offer a series of advantages, including small size, light weight, high functional density, low development cost, short development cycle, and flexible deployment, making them a research hotspot in the domestic and international aerospace fields. Currently, the functions of CubeSats are becoming increasingly sophisticated, and from an overall development trend perspective, the application missions of CubeSats are becoming increasingly complex, which places higher demands on the performance and reliability of CubeSat attitude control systems.

[0003] Domestic universities and research institutions have conducted extensive research on the hardware components and attitude measurement and control algorithms of CubeSat attitude control systems, significantly improving the performance of these systems. However, research on improving the reliability of CubeSat attitude control systems is limited. Due to limitations in mass, size, and energy consumption, CubeSats are characterized by simple protective measures, low redundancy, and limited onboard resources, resulting in relatively low reliability in mission execution. This is especially true for the attitude control system, which typically only has one momentum gear system. Its reliability and stability in orbit determine whether the CubeSat can successfully complete various flight missions, directly impacting its performance and on-orbit lifespan. Therefore, improving the reliability of the attitude control system using limited hardware resources is a pressing engineering challenge. Researching the switching strategy for CubeSat attitude control modes, enabling it to identify individual unit and mission faults in orbit and autonomously switch between bias momentum control and zero momentum control modes, has significant theoretical and practical implications for improving the reliability and stability of the attitude control system in orbit and enhancing the CubeSat's autonomous on-orbit operation capabilities.

[0004] In view of the above problems, designing a highly reliable attitude control mode switching method for satellites, which can enable the entire satellite's attitude to switch autonomously according to mission requirements and redundant faults, is an urgent problem to be solved by technicians in this field. Summary of the Invention

[0005] The present invention aims to provide a method for switching attitude control modes of a CubeSat, which enables the CubeSat to flexibly and reliably switch attitude control modes to complete a designated flight mission.

[0006] The technical solution for implementing the present invention is as follows: A method for switching cubic star attitude control modes, comprising the following steps:

[0007] Step 1: After the satellite separates from the rocket, the attitude control system is powered on, and all attitude sensors and actuators are powered on. The bias momentum wheel starts to rotate. The bias momentum wheel adopts a uniform rotation strategy. When the rated speed is reached, the speed is maintained. At this time, the satellite has an attitude angular rate greater than 5° / s. It is necessary to perform bias momentum angular rate damping control on the satellite to reduce its attitude angular rate. Proceed to Step 2.

[0008] Step 2: During the bias momentum angular rate damping stage, there is no need to control the attitude angle. The desired goal is to control the attitude angular rate to below 0.2° / s. The B-dot control law is used to gradually reduce the satellite's energy, thus gradually decreasing the satellite's attitude angular rate and achieving the purpose of rate damping. This control law does not require feedback of attitude information. Then, it is determined whether the bias momentum angular rate damping is successful. The orbit is considered valid when the magnetic field change rate is less than 1000nT and the damping time is greater than 10000s. If the bias momentum angular rate damping is successful, the filter is turned on and the process proceeds to step 3. Otherwise, step 2 is repeated to continue the bias momentum angular rate damping.

[0009] Step 3: The filters include a single magnetic pitch filter and a magnetometer gyroscope filter. During the bias momentum angular rate damping stage, only the attitude angular rate is controlled. Therefore, after the angular rate damping is completed, there will be a large attitude angle greater than 100°. It is necessary to perform large-angle attitude acquisition to establish a ground pointing reference. During the attitude acquisition stage, both the attitude angle and the attitude angular rate need to be controlled. When the single magnetic pitch filter converges and the attitude acquisition time is greater than 2000s, step 4 is entered to perform bias momentum triaxial stabilization control. Otherwise, step 3 is re-executed to enable the single magnetic pitch filter for large-angle attitude acquisition.

[0010] Step 4: During the bias momentum triaxial stabilization control phase, determine the attitude angular rate of the bias momentum triaxial stabilization control mode. If the attitude angular rate is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and Step 4 is re-executed to perform bias momentum triaxial stabilization control. If the attitude angular rate is less than 1° / s, the count is set to 0. If the fault count is greater than 60, it is considered that the control has failed, and Step 2 is re-executed to perform bias momentum angular rate damping.

[0011] According to mission requirements, the ground-based injection command switches from the bias momentum triaxial stabilization control mode in step 4 to the zero momentum control mode in step 5 and selects the corresponding zero momentum control mode. When the magnetometer gyroscope filter converges, it enters step 5 to start zero momentum control; otherwise, it re-executes step 4 to continue bias momentum triaxial stabilization control.

[0012] Step 5: In the zero momentum control mode, there are two modes, a and b. When entering the zero momentum control mode, the zero momentum mode a is executed first. The speed of the three orthogonal momentum wheels increases uniformly by 3000 revolutions, and the speed of the inclined momentum wheel decreases uniformly by 5196 revolutions. The zero momentum wheel system needs to cross zero.

[0013] When zero momentum mode a fails, it switches to zero momentum mode b. The speed of the three orthogonal momentum wheels is directly adjusted to 3000 rpm, and -5196 rpm for skewed mounting. Sudden changes in the speed of the momentum wheels in this mode may cause attitude roll. Therefore, zero momentum angular rate damping is required when entering zero momentum mode b. In this mode, when the rate of change of the magnetic field is less than 1000 nT, the damping time is greater than 10000 s and the filter converges, zero momentum triaxial stabilization control is performed; otherwise, zero momentum angular rate damping continues to be executed.

[0014] When the satellite enters zero-momentum three-axis stabilization control, it is necessary to determine the attitude angular rate at this time. If it is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and step 5 is executed again to perform zero-momentum three-axis stabilization control. If it is less than 1° / s, the count is set to 0.

[0015] When the attitude angular rate is greater than 1° / s, it is also necessary to determine the zero momentum mode at this time. When the satellite is in zero momentum mode a, if the fault count is greater than 200, it is switched to zero momentum mode b. When the satellite is in zero momentum mode b, if the fault count is greater than 2000, it is switched to step 2 to perform bias momentum angular rate damping.

[0016] Compared to existing attitude control mode switching strategies, the advantages of this invention are:

[0017] (1) It has high reliability and high fault tolerance. It can switch freely between bias momentum mode and zero momentum mode. Since the momentum wheel is three orthogonal and one oblique, its fault tolerance is high. When the zero momentum mode fails, it can switch to bias momentum. It can also use a magnetic torque device for auxiliary control. The attitude control modes are rigorous and comprehensive.

[0018] (2) It has good applicability and high flexibility. It can manually switch modes according to task requirements. When the control accuracy requirement is not high, it adopts bias momentum control. At this time, the control accuracy is low but the power consumption is also low. When the control accuracy requirement is high, it switches to zero momentum control to meet the task requirements.

[0019] (3) It has good stability and stable mission mode switching. Its momentum wheel speed is stable, ensuring that the satellite attitude will not change during the process, and can better control the attitude to complete the predetermined flight mission. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of the framework of a cubic star attitude control mode switching method of the present invention.

[0021] Figure 2 This is a flowchart illustrating the specific implementation of a cubic star attitude control mode switching method according to the present invention.

[0022] Figure 3 This is a simulation result diagram of a cubic star attitude control mode switching method according to the present invention. Detailed Implementation

[0023] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0024] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative positional relationship and movement of each component in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indication will also change accordingly.

[0025] Combination Figure 1 The cubic star attitude control mode switching method described in this invention requires that when the satellite's attitude angle is greater than 40° and its attitude angular rate is greater than 5° / s during separation from the launch vehicle, it needs to enter the bias momentum angular rate damping mode to reduce its angular rate. Subsequently, it enters the bias momentum triaxial stabilization control mode to reduce its attitude angle and determines whether the filter has converged. If it has converged, the zero momentum mode is selected, and finally, it enters the zero momentum triaxial stabilization control state. In case of a fault, it can switch to the bias momentum angular rate damping mode.

[0026] Combination Figure 2 The working principle and process of the cubic star attitude control mode switching method described in this invention are as follows:

[0027] Step 1: After separation of the satellite from the launch vehicle, the attitude control system is powered on, as are all attitude sensors and actuators. The bias momentum wheel begins to rotate, employing a uniform rotation strategy. Once the rated speed is reached, the speed is maintained constant. At this point, the satellite has an attitude angular rate greater than 5° / s, requiring bias momentum angular rate damping control to reduce its attitude angular rate and provide the necessary conditions for subsequent attitude control.

[0028] Step 2: During the bias momentum angular rate damping stage, attitude angle control is unnecessary; the desired goal is to control the attitude angular rate to below 0.2° / s. A B-dot control law can be used to gradually reduce satellite energy, thus decreasing the satellite's attitude angular rate and achieving rate damping. This control law does not require attitude feedback and is highly reliable. Further, it is necessary to determine if the bias momentum angular rate damping is successful. The orbit is considered valid if the magnetic field change rate is less than 1000 nT and the damping time is greater than 10000 s, indicating successful bias momentum angular rate damping and activation of the filter, proceeding to Step 3. Otherwise, Step 2 is re-executed to continue bias momentum angular rate damping.

[0029] Step 3: The filters include a single magnetic pitch filter and a magnetometer-based gyroscope filter. During the bias momentum angular rate damping stage, only the attitude angular rate is controlled. Therefore, after angular rate damping, there will be a large attitude angle greater than 100°, requiring large-angle attitude capture to establish a ground pointing reference. During the attitude capture stage, both the attitude angle and attitude angular rate need to be controlled. The system is set to proceed to step 4 for bias momentum triaxial stabilization control when the single magnetic pitch filter converges and the attitude capture time is greater than 2000s; otherwise, step 3 is re-executed, and the single magnetic pitch filter is activated for large-angle attitude capture.

[0030] Step 4: During the bias momentum triaxial stabilization control phase, the attitude angular rate of the bias momentum triaxial stabilization control mode needs to be determined. When the attitude angular rate is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and Step 4 is re-executed to perform bias momentum triaxial stabilization control. When the attitude angular rate is less than 1° / s, the count is set to 0. When the fault count is greater than 60, it is considered that the control has failed, and Step 2 is re-executed to perform bias momentum angular rate damping.

[0031] Furthermore, in performing bias momentum three-axis stabilization control, the pitch axis bias momentum wheel is primarily used for control. When the pitch axis bias momentum wheel fails, the other three axis momentum wheels combine to form a bias momentum wheel for control. Conversely, if the pitch axis bias momentum wheel does not fail, but control still fails, step 4 is re-executed for bias momentum three-axis stabilization control. In this mode, the cubic star attitude control system is simple, reliable, and has a high fault tolerance.

[0032] According to mission requirements, the ground-based injection command switches from the bias momentum triaxial stabilization control mode in step 4 to the zero momentum control mode in step 5 and selects the corresponding zero momentum control mode. When the magnetometer gyroscope filter converges, it enters step 5 to start zero momentum control; otherwise, it re-executes step 4 to continue bias momentum triaxial stabilization control.

[0033] Step 5: In zero momentum control mode, there are two modes, a and b. When entering zero momentum control mode, mode a is executed first. The speed of the three orthogonal momentum wheels increases uniformly by 3000 rpm, and the speed of the skewed momentum wheel decreases uniformly by 5196 rpm. At this time, since there is no large attitude disturbance, zero momentum three-axis stable control can be directly performed. Generally, the zero momentum wheel system needs to cross zero. When the angular momentum crosses zero, it will cause disturbance to the control system due to dry friction. This method can avoid the interference caused by crossing zero.

[0034] Furthermore, when zero momentum mode a fails, it switches to zero momentum mode b. The speeds of the three orthogonal momentum wheels are directly adjusted to 3000 rpm, and -5196 rpm for skewed mounting. Sudden changes in momentum wheel speed in this mode may cause attitude roll. Therefore, zero momentum angular rate damping is required when entering zero momentum mode b. In this mode, zero momentum triaxial stabilization control is performed when the magnetic field change rate is less than 1000 nT, the damping time is greater than 10000 s, and the filter converges; otherwise, zero momentum angular rate damping continues.

[0035] Furthermore, when the satellite enters zero-momentum three-axis stabilization control, it is necessary to determine the attitude angular rate at this time. If it is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and step 5 is executed again to perform zero-momentum three-axis stabilization control. If it is less than 1° / s, the count is set to 0.

[0036] Furthermore, when the attitude angular rate is greater than 1° / s, it is also necessary to determine the zero momentum mode at this time. When the satellite is in zero momentum mode a, if the fault count is greater than 200, it switches to zero momentum mode b. When the satellite is in zero momentum mode b, if the fault count is greater than 2000, it switches to step 2 to perform bias momentum angular rate damping.

[0037] The momentum wheel mounted on a CubeSat is relatively small, with a diameter of less than 2 cm, resulting in weak control capabilities and a maximum output torque of less than 2 mNm (million Newton-meters). This places high demands on the attitude control algorithm, primarily relying on bias momentum control with zero momentum control as a secondary measure. The method described in this invention mainly involves switching between bias momentum control and zero momentum control, making full use of limited hardware resources. The mode switching logic is clear, and the software implementation is simple, effectively improving the reliability of the CubeSat attitude control system.

[0038] Figure 3 This is a simulation result diagram of a cubic star attitude control mode switching method of the present invention. The simulation first performs a bias momentum angular rate damping mode, and it can be seen that the attitude quaternion vector part begins to converge. When switching to the bias momentum triaxial stabilization mode, it can be seen that the attitude quaternion vector part tends to 0. When switching to the zero momentum mode, it can be seen that the attitude quaternion vector part jitters near 0. The simulation results show that the present invention can well achieve the predetermined functional requirements and solve practical problems.

[0039] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for switching cubic star attitude control modes, characterized in that, The steps are as follows: Step 1: After the star and rocket separate, the attitude control system is powered on, and all attitude sensors and actuators are powered on. The bias momentum wheel starts to rotate. The bias momentum wheel adopts a uniform speed rotation strategy. After reaching the rated speed, the speed is maintained at a constant speed. At this point, the satellite has an attitude angular rate greater than 5° / s, and it is necessary to perform bias momentum angular rate damping control on the satellite to reduce its attitude angular rate, and proceed to step 2. Step 2: In the bias momentum angular rate damping stage, the desired goal is to control the attitude angular rate to below 0.2° / s. The B-dot control law is used to gradually reduce the satellite's energy and thus gradually decrease the satellite's attitude angular rate. Then, it is determined whether the bias momentum angular rate damping is successful. The orbit is considered valid when the magnetic field change rate is less than 1000nT and the damping time is greater than 10000s. If the bias momentum angular rate damping is successful, the filter is activated and the process proceeds to Step 3. Otherwise, Step 2 is re-executed to continue the bias momentum angular rate damping. Step 3: The filters include a single magnetic pitch filter and a magnetometer gyroscope filter. During the bias momentum angular rate damping stage, only the attitude angular rate is controlled. Therefore, after the angular rate damping is completed, there will be a large attitude angle greater than 100°. It is necessary to perform large-angle attitude capture to establish a ground pointing reference. During the attitude capture stage, both the attitude angle and the attitude angular rate need to be controlled. When the single magnetic pitch filter converges and the attitude capture time is greater than 2000s, step 4 is entered to perform bias momentum triaxial stabilization control. Otherwise, step 3 is re-executed to enable the single magnetic pitch filter for large-angle attitude capture. Step 4: During the bias momentum triaxial stabilization control phase, determine the attitude angular rate of the bias momentum triaxial stabilization control mode. If the attitude angular rate is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and Step 4 is re-executed to perform bias momentum triaxial stabilization control; if the attitude angular rate is less than 1° / s, the count is reset to 0; if the fault count is greater than 60, it is considered that the control has failed, and Step 2 is re-executed to perform bias momentum angular rate damping. According to mission requirements, the ground-based injection command switches from the bias momentum triaxial stabilization control mode in step 4 to the zero momentum control mode in step 5 and selects the corresponding zero momentum control mode. When the magnetometer gyroscope filter converges, it enters step 5 to start zero momentum control; otherwise, it re-executes step 4 to continue bias momentum triaxial stabilization control. Step 5: In the zero momentum control mode, there are two modes, a and b. When entering the zero momentum control mode, the zero momentum mode a is executed first. The speed of the three orthogonal momentum wheels increases uniformly by 3000 revolutions, and the speed of the inclined momentum wheel decreases uniformly by 5196 revolutions. The zero momentum wheel system needs to cross zero. When zero momentum mode a fails, it switches to zero momentum mode b, and the speed of the three orthogonal momentum wheels is directly adjusted to 3000 rpm, and the speed of the oblique mounting is -5196 rpm. When entering zero momentum mode b, zero momentum angular rate damping is required. In this mode, when the rate of change of the magnetic field is less than 1000nT, the damping time is greater than 10000s and the filter converges, zero momentum triaxial stabilization control is performed; otherwise, zero momentum angular rate damping continues to be executed. When the satellite enters zero-momentum three-axis stabilization control, it is necessary to determine the attitude angular rate at this time. If it is greater than 1° / s, it is considered that the control error is large, the fault count is incremented by 1, and step 5 is executed again to perform zero-momentum three-axis stabilization control. If it is less than 1° / s, the count is set to 0. When the attitude angular rate is greater than 1° / s, it is also necessary to determine the zero momentum mode at this time. When the satellite is in zero momentum mode a, if the fault count is greater than 200, it is switched to zero momentum mode b. When the satellite is in zero momentum mode b, if the fault count is greater than 2000, it is switched to step 2 to perform bias momentum angular rate damping.

2. The cubic star attitude control mode switching method according to claim 1, characterized in that, In step 4, when performing bias momentum triaxial stabilization control, the pitch axis bias momentum wheel is used for control. When the pitch axis bias momentum wheel fails, the other three axis momentum wheels are combined to form a bias momentum wheel for control. On the other hand, if the pitch axis bias momentum wheel does not fail, but the control still fails, step 4 is executed again to perform bias momentum triaxial stabilization control.

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