Satellite attitude control method and system when gyro data cannot be used, and satellite

Abstract

The invention discloses a satellite attitude control method and system when gyro data cannot be used and a satellite, and relates to the technical field of satellite attitude control, the method comprises the following steps: when it is determined that the gyro data of the satellite cannot be used, judging whether star sensing data collected by a star sensor configured by the satellite in a current control period is valid, if yes, executing the step 2, and if not, executing the step 3; determining the attitude data of the satellite at the current moment according to the star sensor data collected by the star sensor in the current control period; and when a preset condition is met, performing attitude control on the satellite according to the attitude data of the satellite at the current moment. In the invention, when the gyro data of the satellite cannot be used, the attitude data of the satellite at the current moment can be determined by using the star sensor data in the current control period acquired by the star sensor configured on the satellite, so that the attitude control can be effectively and timely performed on the satellite, the task implementation can be guaranteed to the greatest extent, and the satellite attitude control accuracy is improved. The engineering implementation operability is high.

Description

Background Art

[0002] A satellite attitude control system usually consists of sensors, actuators, and control software, which controls the satellite attitude to be in the desired position relative to a reference datum, and thus provides an accurately guaranteed and reliable attitude pointing for the satellite to complete its tasks.

[0003] Satellite attitude control includes attitude determination, attitude stabilization, and attitude maneuver control. Attitude determination is to determine the orientation or pointing of the satellite relative to a reference datum, and thus obtain attitude angle parameters; attitude stabilization control is to keep the satellite attitude at the expected specified direction and specified value, and attitude maneuver is the reorientation process for the satellite to transition from one attitude to another.

[0004] Commonly used sensors in a satellite attitude control system include: gyroscopes, star sensors, sun sensors, earth sensors, magnetometers, etc. Among them, gyroscopes are used to measure the angular motion of the satellite attitude, star sensors are used to achieve inertial attitude determination at discrete moments, sun sensors are used to measure sun vector information, earth sensors are used to measure earth vector information, and magnetometers are used to sense the geomagnetic field information at the position where the satellite is located.

[0005] Commonly used actuators in a satellite attitude control system include: flywheels, magnetic torque actuators, jet propulsion systems, electric propulsion systems, etc. Among them, low-earth orbit satellites are often equipped with flywheels and magnetic torque actuators.

[0006] The "star sensor + gyroscope" scheme is a commonly used high-precision attitude determination scheme for satellites. By using the measurement information of star sensors and gyroscopes and passing through a combined filter, high-precision attitude determination can be achieved. When the measurement information of sun sensors, earth sensors, and magnetometers is valid, the information of any two of them can be used to implement satellite attitude determination by using the double-vector attitude determination algorithm.

[0007] Currently, satellites with relatively low requirements for attitude stability are often equipped with low-precision fiber optic gyroscopes or MEMS gyroscopes, while satellites with higher precision requirements are equipped with multiple sets of three-axis fiber optic gyroscopes or one set of fiber optic gyroscopes with multi-axis redundant configuration.

[0008] The working principle of a fiber optic gyroscope is based on the Sagnac Effect, which is a general effect of light propagating in a closed optical path. When two beams of light with equal characteristics emitted from the same light source propagate in opposite directions in the same closed optical path, if the plane where the closed optical path is located has an angular velocity of rotation relative to inertial space, then the optical paths traveled by the forward and backward propagating light beams will be different, thus generating an optical path difference. This optical path difference is proportional to the angular velocity of rotation. By detecting this optical path difference and its corresponding phase difference, the angular velocity of rotation can be determined. Currently, the mainstream fiber optic gyroscope adopts a closed-loop demodulation scheme with a PID control algorithm. The fiber optic gyroscope has an obvious startup characteristic at the moment of power-on. The startup time is usually between 5 s and 15 s. The characteristic of this startup is that the output is completely irregular within the startup time after power-on and the angular velocity measurement cannot be achieved. The fiber optic gyroscope also has failure modes such as open-loop, constant value output, and abnormal output data.

[0009] The working principle of a MEMS gyroscope (Micro-Electro-Mechanical Systems gyroscope) is based on the Coriolis force. The Coriolis force is a description of the linear motion deviation of a particle in linear motion in a rotating system due to inertia relative to the rotating system. The MEMS gyroscope measures the angular velocity through the Coriolis force. The MEMS gyroscope also has an obvious startup characteristic at the moment of power-on. The startup time is usually between 1 s and 5 s. The characteristic of this startup is that the output data within the startup time after power-on cannot meet the requirements of the steady-state index and cannot be used for attitude closed-loop control.

[0010] The prerequisite for a spacecraft to perform attitude control is to accurately describe the attitude of the spacecraft. Currently, the common attitude description methods mainly include: rotation matrix, Euler angles, Euler quaternions, and modified Rodriguez parameters. Among them, Euler angles and MRPs have singularities, and the description of attitude by Euler quaternions has the disadvantage of non-uniqueness. The rotation matrix can globally and uniquely describe the attitude of a rigid body. In early research work, the three-dimensional special orthogonal matrix group SO(3) composed of rotation matrices is a non-Euclidean manifold. Due to factors such as the imperfect development of its model stability theory, once control design is involved, other parameterized models are usually selected for use. Currently, the research on Lie groups is gradually developing, and the control method based on SO(3) shows control performance with higher control accuracy and faster speed.

[0011] Due to the obvious start-up characteristics of current satellite fiber optic gyroscopes and MEMS gyroscopes when powered on, the satellite attitude control system has no effective gyro measurement angular velocity data for satellite attitude closed-loop control during gyroscope failures or main backup gyroscope switches. It is extremely important for the satellite to adopt effective control logic and attitude control methods in this state, especially to ensure the stable and controllable satellite attitude and the smooth implementation of tasks when high-value satellites encounter gyroscope failures or main backup gyroscope switches during mission execution. Summary of the Invention

[0012] The technical problem to be solved by the present invention is to provide a satellite attitude control method, system, and satellite when gyro data is unavailable, specifically as follows:

[0013] 1) In the first aspect, the present invention provides a satellite attitude control method when gyro data is unavailable, and the specific technical solution is as follows:

[0014] When it is determined that the gyro data of the satellite is unavailable, it is judged whether the star sensor data collected by the star sensors configured on the satellite within the current control cycle is valid, and a first judgment result is obtained;

[0015] When the first judgment result is yes, the attitude data of the satellite at the current moment is determined according to the star sensor data collected by the star sensors within the current control cycle;

[0016] When the preset conditions are met, attitude control is performed on the satellite according to the attitude data of the satellite at the current moment.

[0017] The beneficial effects of the satellite attitude control method provided by the present invention when gyro data is unavailable are as follows:

[0018] When the gyro data of the satellite is unavailable, the attitude data of the satellite at the current moment can be determined by using the star sensor data collected by the star sensors carried on the satellite within the first preset historical time period, and then the satellite can be effectively and timely controlled in terms of attitude, which can maximize the guarantee of mission implementation, and the engineering implementation of the present invention has strong operability.

[0019] On the basis of the above solution, the satellite attitude control method of the present invention when gyro data is unavailable can be further improved as follows.

[0020] Furthermore, it further includes:

[0021] When the first judgment result is no, it is judged whether the data collected by the earth sensors configured on the satellite and the data collected by the sun sensors configured on the satellite within the current control cycle are both valid, and a second judgment result is obtained;

[0022] When the second judgment result is yes, determine the attitude data of the satellite at the current moment according to the data collected by the earth sensor in the current control cycle and the data collected by the sun sensor configured on the satellite in the current control cycle.

[0023] Further, it further includes:

[0024] When the second judgment result is no, determine the attitude data of the satellite at the current moment according to the satellite attitude data cached within the preset historical time period.

[0025] Further, it further includes: Determine whether the preset conditions are met according to the fault diagnosis result and the fault duration that cause the gyro data of the satellite to be unusable.

[0026] 2) In a second aspect, the present invention further provides a satellite attitude control system when gyro data is unusable, and the specific technical solution is as follows:

[0027] It includes a judgment module, an attitude data acquisition module, and an attitude control module;

[0028] The judgment module is used for: When it is determined that the gyro data of the satellite is unusable, judge whether the star sensor data collected by the star sensor configured on the satellite in the current control cycle is valid, and obtain a first judgment result;

[0029] The attitude data acquisition module is used for: When the first judgment result is yes, determine the attitude data of the satellite at the current moment according to the star sensor data collected by the star sensor in the current control cycle;

[0030] The attitude control module is used for: When the preset conditions are met, perform attitude control on the satellite according to the attitude data of the satellite at the current moment.

[0031] On the basis of the above solution, a satellite attitude control system when gyro data is unusable according to the present invention can be further improved as follows.

[0032] Further, the judgment module is further used for: When the first judgment result is no, judge whether both the data collected by the earth sensor configured on the satellite in the current control cycle and the data collected by the sun sensor configured on the satellite in the current control cycle are valid, and obtain a second judgment result;

[0033] The attitude data acquisition module is further used for: When the second judgment result is yes, determine the attitude data of the satellite at the current moment according to the data collected by the earth sensor in the current control cycle and the data collected by the sun sensor configured on the satellite in the current control cycle.

[0034] Further, the attitude data acquisition module is further used for:

[0035] When the second judgment result is negative, determine the satellite attitude data at the current moment according to the satellite attitude data cached by the satellite within a preset historical time period.

[0036] Furthermore, the judgment module is further configured to: determine whether the preset conditions are met according to the fault diagnosis result and the fault duration that cause the gyro data of the satellite to be unusable.

[0037] 3) In a third aspect, the present invention further provides a satellite, including an on-board processor, and the on-board processor is configured to execute the satellite attitude control method when the gyro data is unusable according to any one of the above.

[0038] 4) In a fourth aspect, the present invention further provides an electronic device, which includes a processor. The processor is coupled to a memory, and at least one computer program is stored in the memory. The at least one computer program is loaded and executed by the processor to enable the electronic device to implement the satellite attitude control method when the gyro data is unusable according to any one of the above.

[0039] 5) In a fifth aspect, the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the satellite attitude control method when the gyro data is unusable according to any one of the above is implemented.

[0040] It should be noted that for the beneficial effects obtained by the technical solutions of the second to fifth aspects of the present invention and the corresponding possible implementation manners, reference may be made to the technical effects of the first aspect and its corresponding possible implementation manners described above, and details are not described herein again. Description of the Drawings

[0041] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for description in the embodiments of the present invention:

[0042] Figure 1 It is a schematic flowchart of a satellite attitude control method when the gyro data is unusable according to an embodiment of the present invention;

[0043] Figure 2 It is a schematic diagram of zero-attitude and zero-angular velocity feedback attitude stabilization control simulation in an orbital system;

[0044] Figure 3 It is a schematic diagram of small-angle maneuver and zero-angular velocity feedback attitude stabilization control simulation in an orbital system;

[0045] Figure 4 It is a schematic diagram of the error-axis angle of small-angle maneuver and zero-angular velocity feedback attitude stabilization control simulation in an orbital system;

[0046] Figure 5Schematic diagram of a satellite attitude control system when gyro data cannot be used according to an embodiment of the present invention;

[0047] Figure 6 Schematic diagram of an electronic device according to an embodiment of the present invention. Detailed implementation manners

[0048] The principles and features of the present invention are described below. The examples given are only used to explain the present invention and are not intended to limit the scope of the present invention.

[0049] The technical solutions of the present invention and how the technical solutions of the present invention solve the above technical problems are described in detail below with specific embodiments. These several specific embodiments can be combined with each other. For the same or similar concepts or processes, they may not be repeated in some embodiments. The embodiments of the present invention will be described below with reference to the accompanying drawings.

[0050] As Figure 1 shown, a satellite attitude control method when gyro data cannot be used according to an embodiment of the present invention includes the following steps:

[0051] S1. When it is determined that the gyro data of the satellite cannot be used, it is judged whether the star sensor data collected by the star sensors configured on the satellite in the current control period is valid, and a first judgment result is obtained;

[0052] Among them, the current control period refers to: the control period to which the moment when it is determined that the gyro data of the satellite cannot be used belongs.

[0053] Among them, when the star sensor outputs star sensor data, it will carry an identifier indicating whether the star sensor data is valid. Whether the star sensor data of the star sensors configured on the satellite in the current control period collected is valid is judged through this identifier.

[0054] Among them, the sensors equipped in the attitude control system include star sensors, gyro sensors, earth sensors, sun sensors, etc., and the satellite actuators are configured with flywheels, magnetometers, etc.

[0055] Among them, the fault conditions of the gyro sensors of the satellite attitude control system are diagnosed to determine the fault diagnosis results, and a fault category identifier is set for each fault diagnosis result. Specifically:

[0056] 1) If all the gyro sensors on the satellite fail, the fault category identifier is recorded as: the first type of fault.

[0057] Among them, the process of judging that any gyro sensor fails is as follows:

[0058] When the angular velocity output by any gyro sensor exceeds the preset angular velocity threshold and the duration exceeds the first preset duration, it is determined that the gyro sensor has failed. For the sake of distinction, the duration here is denoted as the first preset duration. The preset angular velocity threshold and the first duration can be determined according to the actual situation and ground tests. Or, when the change amount between every two adjacent angular velocities output by any gyro sensor does not exceed the preset change amount threshold and the duration exceeds the second preset duration, it is determined that the gyro sensor has failed. For the sake of distinction, the duration here is denoted as the second preset duration. The preset change amount threshold and the second duration can be determined according to the actual situation and ground tests.

[0059] 2) When some of the gyro sensors of the satellite fail and there are still normal gyro sensors (gyro sensors that can operate normally) that have not been powered on, after powering on the normal gyro sensors that have not been powered on, during the startup process, the data output by the gyro sensor is unstable and cannot be directly used for attitude determination and closed-loop control. The fault category identifier of this fault is denoted as: the second type of fault.

[0060] 3) If all the gyro sensors on the satellite are normal, but during the power-on startup process of each gyro sensor, the data output by the gyro sensor is unstable and cannot be directly used for attitude determination and closed-loop control. The fault category identifier of this fault is denoted as: the third type of fault.

[0061] When the above-mentioned first type of fault, second type of fault, or third type of fault occurs, it is determined that the gyro data of the satellite cannot be used. If none of the first type of fault, second type of fault, and third type of fault occur, it is determined that the gyro data of the satellite can be used.

[0062] S2. When the first judgment result is yes, according to the star sensor data collected in the current control cycle, determine the attitude data of the satellite at the current moment;

[0063] S3. When the preset conditions are met, perform attitude control on the satellite according to the attitude data of the satellite at the current moment.

[0064] Optionally, in the above technical solution, it further includes:

[0065] When the first judgment result is no, judge whether the data collected by the earth sensor configured on the satellite and the data collected by the sun sensor configured on the satellite in the current control cycle are both valid, and obtain the second judgment result;

[0066] When the earth sensor outputs data, it will carry an identifier indicating whether the output data is valid. Through this identifier, it is possible to determine whether the data of the current control cycle of the earth sensor configured on the satellite is valid. Similarly, when the sun sensor outputs data, it will carry an identifier indicating whether the output data is valid. Through this identifier, it is possible to determine whether the data of the current control cycle of the sun sensor configured on the satellite is valid.

[0067] When the second judgment result is yes, based on the data of the current control cycle collected by the earth sensor and the data of the current control cycle collected by the sun sensor configured on the satellite, determine the attitude data of the satellite at the current moment.

[0068] Optionally, in the above technical solution, it further includes:

[0069] When the second judgment result is no, based on the satellite attitude data cached by the satellite within a preset historical time period, determine the attitude data of the satellite at the current moment.

[0070] Among them, the preset historical time period is: the time period between the start time and the end time. The end time is: the moment when it is determined that the gyro data of the satellite cannot be used. The duration between the start time and the end time is the preset duration, such as 1 minute or 5 minutes, etc., which can be set according to the actual situation. Table 1 lists the satellite attitude data cached by the satellite within the preset historical time period, specifically including the roll angle, pitch angle, and yaw angle at different star times.

[0071] Table 1:

[0072]

[0073]

[0074] Based on the satellite attitude data (roll angle, pitch angle, and yaw angle) cached by the satellite within the preset historical time period, determine the attitude data of the satellite at the current moment. The specific implementation process is as follows:

[0075] Based on the satellite attitude data cached by the satellite within the preset historical time period, determine the corresponding relationships between the roll angle, pitch angle, and yaw angle and the star time respectively:

[0076]

[0077] Among them, φ t represents: the roll angle of the satellite at star time t, θ t represents: the pitch angle of the satellite at star time t, ψ t represents: the yaw angle of the satellite at star time t, and a3, a2, a1, a0, b3, b2, b1, b0, c3, c2, c1, and c0 are all coefficients.

[0078] ① The acquisition methods of a3, a2, a1, and a0 are as follows:

[0079] Construct a set of formulas: Denote this set of formulas as Ang_φ = M φ *X, where Use the least squares method to fit Ang_φ = M φ *X to obtain a3, a2, a1, and a0, and further obtain the roll angle φ n+1 at the current moment t n+1 , φ n+1 = a3*(t n + t step ) 3 + a2*(t n + t step ) 2 + a1*(t n + t step ) + a0, t n+1 = t n + t step , t step represents: the time step of control, which can be set according to the actual situation.

[0080] ② The acquisition process of b3, b2, b1, and b0 is as follows:

[0081] Construct a set of formulas: Denote this set of formulas as: Ang_θ = M θ *Y, where Use the least squares method to fit Ang_θ = M θ *Y to obtain b3, b2, b1, and b0, and further obtain the roll angle θ n+1 at the current moment t n+1 , θ n+1 = b3*(t n + t step ) 3 + b2*(t n + t step ) 2 + b1*(t n + t step ) + b0.

[0082] ③ The acquisition process of c3, c2, c1, and c0 is as follows:

[0083] Construct a set of formulas: Denote this set of formulas as: Ang_ψ = M ψ *Z, where Use the least squares method to fit Ang_ψ = M ψPerform fitting on *Z*, calculate to obtain c3, c2, c1, and c0, and then obtain the current time t n+1 of the pitch angle ψ n+1 , ψ n+1 = c3*(t n + t step ) 3 + c2*(t n + t step ) 2 + c1*(t n + t step ) + c0

[0084] Optionally, in the above technical solution, it further includes:

[0085] According to the fault diagnosis result and the fault duration that cause the gyro data of the satellite to be unusable, determine whether it meets the preset conditions. When it meets the preset conditions, perform attitude control on the satellite according to the attitude data of the satellite at the current moment. When it does not meet the preset conditions, during the process of performing attitude control on the satellite, the attitude data of the satellite at the current moment is not required. Specifically:

[0086] 1) When the fault diagnosis result is not a type-I fault, judge whether the fault duration exceeds the first preset duration threshold to obtain a third judgment result. Then:

[0087] ① When the third judgment result is yes, use a PID controller with angular rate participation for closed-loop attitude control, calculate the control torque, so that the actuator of the satellite outputs the control torque to perform closed-loop attitude control on the satellite;

[0088] ② When the third judgment result is no, based on the calculated attitude data of the satellite at the current moment, use an SO(3) controller without angular rate participation for closed-loop attitude control, calculate the control torque, so that the actuator of the satellite outputs the control torque to perform closed-loop attitude control on the satellite;

[0089] Based on the calculated attitude data of the satellite at the current moment, use an SO(3) controller without angular rate participation for closed-loop attitude control, calculate the control torque. The specific implementation process is as follows:

[0090] a. Define the vector calculation method: Define any vector x, Define: (x × ) ∨ = x;

[0091] b. Calculate the control torque using the following formula:

[0092] τ = -κ Q e Q - κ R Re

[0093] Among them, Ω = Γe Q , τ represents the control torque, κ Q and κ R are both coefficients of the SO(3) controller. κ Q and κ R are determined through simulation tests and ground tests according to the requirements of the satellite for the rapidity, stability and other indicators of the control system. Q and Γ are both intermediate variables. The initial values of Q and Γ can be set according to the actual situation. Q belongs to the three-dimensional special orthogonal group. R is the attitude matrix at the current moment calculated according to the attitude angles φ n+1 , θ n+1 , ψ n+1 at the current moment. R d is the target attitude matrix determined according to the attitude control target of the satellite. tr(R T Q) represents calculating the trace of R T Q.

[0094] 2) When the fault that occurs is a type-I fault, judge whether the fault duration exceeds the second preset duration threshold to obtain a fourth judgment result. The first preset duration threshold is less than the second preset duration threshold;

[0095] ① When the fourth judgment result is yes, the control method for satellite attitude control is: using the actuator for body magnetic damping control.

[0096] ② When the fourth judgment result is no, based on the calculated attitude data of the satellite at the current moment, use the SO(3) controller without angular rate participation for closed-loop attitude control, calculate the control torque, so that the actuator of the satellite outputs the control torque for closed-loop attitude control of the satellite.

[0097] Among them, meeting the preset conditions means: meeting "the fault diagnosis result is not a type-I fault and the third judgment result is no", or meeting "the fault that occurs is a type-I fault and the fourth judgment result is no". When not meeting "the fault diagnosis result is not a type-I fault and the third judgment result is no" or "the fault that occurs is a type-I fault and the fourth judgment result is no", it is determined that the preset conditions are not met.

[0098] When it is determined that the gyro data of the satellite cannot be used, according to the time step t step of the control time, obtain the attitude data of the satellite at multiple moments in real time, and perform attitude control on the satellite according to the fault diagnosis result, fault duration and the latest attitude data of the satellite that cause the gyro data of the satellite to be unusable.

[0099] The simulation results of attitude control without angular rate feedback using the attitude handling control method of the present invention are as follows Figure 2 、 Figure 3 and Figure 4 shown.

[0100] In the above embodiments, although the steps are numbered S1, S2, etc., these are only specific embodiments given by the present invention. Those skilled in the art can adjust the execution order of S1, S2, etc. according to the actual situation, and this is also within the protection scope of the present invention. It can be understood that in some embodiments, it may include some or all of the above embodiments.

[0101] As Figure 5 shown, a satellite attitude control system 200 when gyro data cannot be used according to an embodiment of the present invention includes a judgment module 201, an attitude data acquisition module 202, and an attitude control module 203;

[0102] The judgment module 201 is configured to: when it is determined that the gyro data of the satellite cannot be used, judge whether the star sensor data collected by the star sensors configured on the satellite in the current control period is valid, and obtain a first judgment result;

[0103] The attitude data acquisition module 202 is configured to: when the first judgment result is yes, determine the attitude data of the satellite at the current moment according to the star sensor data collected by the star sensors in the current control period;

[0104] The attitude control module 203 is configured to: when the preset conditions are met, perform attitude control on the satellite according to the attitude data of the satellite at the current moment.

[0105] Optionally, in the above technical solution, the judgment module 201 is further configured to: when the first judgment result is no, judge whether the data collected by the earth sensors configured on the satellite in the current control period and the data collected by the sun sensors configured on the satellite in the current control period are both valid, and obtain a second judgment result;

[0106] The attitude data acquisition module 202 is further configured to: when the second judgment result is yes, determine the attitude data of the satellite at the current moment according to the data collected by the earth sensors in the current control period and the data collected by the sun sensors configured on the satellite in the current control period.

[0107] Optionally, in the above technical solution, the attitude data acquisition module 202 is further configured to:

[0108] when the second judgment result is no, determine the attitude data of the satellite at the current moment according to the satellite attitude data cached by the satellite within a preset historical time period.

[0109] Optionally, in the above technical solution, the determination module 201 is further configured to: determine whether a preset condition is met according to a fault diagnosis result that causes the gyro data of the satellite to be unusable and the duration of the fault.

[0110] It should be noted that the beneficial effects of the satellite attitude control system 200 when the gyro data is unusable provided in the above embodiment are the same as those of the satellite attitude control method when the gyro data is unusable, which will not be elaborated here. In addition, when the system provided in the above embodiment implements its functions, only the division of the above function modules is used as an example. In practical applications, the above functions can be allocated to different function modules according to needs, that is, the system can be divided into different function modules according to the actual situation to complete all or part of the functions described above. In addition, the system provided in the above embodiment and the method embodiment belong to the same concept, and the specific implementation process can be seen in the method embodiment, which will not be elaborated here.

[0111] Among them, the satellite attitude control system of the present invention when the gyro data is unusable can be a computer program (including program code) running in a computer device. For example, the satellite attitude control system of the present invention when the gyro data is unusable is an application software, which can be used to execute the corresponding steps in the satellite attitude control method of the present invention when the gyro data is unusable.

[0112] In some embodiments, the satellite attitude control system of the present invention when the gyro data is unusable can be implemented in a combination of software and hardware. As an example, the satellite attitude control system of the present invention when the gyro data is unusable can be a processor in the form of a hardware decoding processor, which is programmed to execute the satellite attitude control method of the present invention when the gyro data is unusable. For example, the processor in the form of a hardware decoding processor can use one or more application specific integrated circuits (ASICs), DSPs, programmable logic devices (PLDs), complex programmable logic devices (CPLDs), field programmable gate arrays (FPGAs) or other electronic components.

[0113] Among them, the modules described in the embodiments of the present invention can be implemented by software or by hardware. Among them, the name of the module does not constitute a limitation to the module itself in some cases.

[0114] The present invention also provides a satellite, including an on-board processor for executing the satellite attitude control method according to any one of the above when gyro data is unavailable.

[0115] An electronic device according to an embodiment of the present invention includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the satellite attitude control method according to any one of the above when gyro data is unavailable. That is to say, an electronic device according to an embodiment of the present invention may include, but is not limited to, a processor and a memory; the memory is used for storing a computer program; the processor is used for executing the satellite attitude control method according to any embodiment of the present invention shown when gyro data is unavailable by calling the computer program.

[0116] In an optional embodiment, an electronic device is provided, as Figure 6 shown, Figure 6 The electronic device 4000 shown includes a processor 4001 and a memory 4003. Among them, the processor 4001 and the memory 4003 are connected, such as connected through a bus 4002. Optionally, the electronic device 4000 may further include a transceiver 4004, and the transceiver 4004 may be used for data interaction between the electronic device and other electronic devices, such as data sending and / or data receiving, etc. It should be noted that in practical applications, the transceiver 4004 is not limited to one, and the structure of the electronic device 4000 does not constitute a limitation to the embodiment of the present invention.

[0117] The processor 4001 may be a CPU (Central Processing Unit, central processor), a general-purpose processor, a DSP (Digital Signal Processor, data signal processor), an ASIC (Application Specific Integrated Circuit, application-specific integrated circuit), an FPGA (Field Programmable Gate Array, field programmable gate array), or other programmable logic devices, transistor logic devices, hardware components, or any combination thereof. It can implement or execute various exemplary logical blocks, modules, and circuits described in connection with the disclosure of the present invention. The processor 4001 may also be a combination for implementing a computing function, such as a combination including one or more microprocessors, a combination of a DSP and a microprocessor, etc.

[0118] The bus 4002 may include a path for transmitting information among the above components. The bus 4002 may be a PCI (Peripheral Component Interconnect) bus, an EISA (Extended Industry Standard Architecture) bus, or the like. The bus 4002 may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Figure 6 only a thick line is used to represent the bus 4002 in Figure 6 , but it does not mean that there is only one bus or one type of bus.

[0119] The memory 4003 may be a ROM (Read Only Memory) or other type of static storage device that can store static information and instructions, a RAM (Random Access Memory) or other type of dynamic storage device that can store information and instructions, or it may also be an EEPROM (Electrically Erasable Programmable Read Only Memory), a CD-ROM (Compact Disc Read Only Memory), or other optical disc storage, optical disc storage (including compact discs, laser discs, optical discs, digital versatile discs, Blu-ray discs, etc.), magnetic disk storage media, or other magnetic storage devices, or any other medium that can be used to carry or store the desired program code in the form of instructions or data structures and can be accessed by a computer, but is not limited thereto.

[0120] The memory 4003 is used to store the application program code (computer program) for implementing the solution of the present invention and is controlled by the processor 4001 for execution. The processor 4001 is used to execute the application program code stored in the memory 4003 to implement the content shown in the foregoing method embodiments.

[0121] Among them, the electronic device may also be a terminal device, and the terminal device may be any device on which an application can be installed, including at least one of a smart phone, a tablet computer, a notebook computer, a desktop computer, a smart speaker, a smart watch, a smart TV, and a smart vehicle device.

[0122] It should be noted that Figure 6 the electronic device shown is only an example and should not impose any limitations on the functions and usage scope of the embodiments of the present invention.

[0123] A computer-readable storage medium according to an embodiment of the present invention has a computer program stored thereon. When the computer program is executed by a processor, the satellite attitude control method when gyro data cannot be used as described in any one of the above is implemented.

[0124] Optionally, the computer-readable storage medium may be a read-only memory (ROM), a random access memory (RAM), a compact disc read-only memory (CD-ROM), magnetic tape, a floppy disk, an optical data storage device, etc.

[0125] In an exemplary embodiment, a computer program product or a computer program is further provided. The computer program product or the computer program includes computer instructions, and the computer instructions are stored in a computer-readable storage medium. The processor of the electronic device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the electronic device executes the satellite attitude control method when gyro data cannot be used as described in any one of the above.

[0126] Computer program code for performing the operations of the present invention may be written in one or more programming languages or combinations thereof. The above programming languages include object-oriented programming languages such as Java, Smalltalk, C++, and also include conventional procedural programming languages such as the "C" language or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, executed as a stand-alone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computer (for example, by using an Internet service provider to connect through the Internet).

[0127] It should be understood that the flowcharts and block diagrams in the accompanying drawings illustrate the possible architectures, functions, and operations of methods and computer program products according to various embodiments of the present invention. In this regard, each block in the flowchart or block diagram may represent a module, a program segment, or a part of code that contains one or more executable instructions for implementing a specified logical function. It should also be noted that in some alternative implementations, the functions marked in the blocks may occur in a different order than that marked in the accompanying drawings. For example, two consecutive blocks shown may actually be executed substantially in parallel, and they may sometimes be executed in the reverse order, depending on the functions involved. It should also be noted that each block in the block diagram and / or flowchart, as well as the combinations of blocks in the block diagram and / or flowchart, may be implemented by a dedicated hardware-based system that performs the specified functions or operations, or may be implemented by a combination of dedicated hardware and computer instructions.

[0128] The computer-readable storage medium provided by the embodiments of the present invention may be, but is not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the above. More specific examples of the computer-readable storage medium may include, but are not limited to: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EEPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present invention, the computer-readable storage medium may be any tangible medium that contains or stores a program that can be used by or in conjunction with an instruction execution system, apparatus, or device.

[0129] The above computer-readable storage medium carries one or more programs that, when executed by the electronic device, cause the electronic device to perform the methods shown in the above embodiments.

[0130] The above description is only a preferred embodiment of the present invention and an explanation of the applied technical principles. Those skilled in the art should understand that the scope of disclosure involved in the present invention is not limited to the technical solutions formed by the specific combination of the above technical features, and should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the above disclosure concept. For example, the technical solutions formed by mutually replacing the above features with the technical features (but not limited to) having similar functions disclosed in the present invention.

[0131] It should be noted that the terms "first", "second", etc. in the description and claims of this application are used to distinguish similar objects, rather than to limit a specific order or sequence. Under appropriate circumstances, the order of use of similar objects can be interchanged so that the embodiments of the present application described herein can be implemented in an order other than the order shown or described.

[0132] Those skilled in the art know that the present invention can be implemented as a system, a method, or a computer program product. Therefore, the present invention can be specifically implemented in the following forms, that is, it can be completely hardware, or completely software (including firmware, resident software, microcode, etc.), or a combination of hardware and software, which is generally referred to as "circuit", "module", or "system" herein. In addition, in some embodiments, the present invention can also be implemented in the form of a computer program product in one or more computer-readable media, which contain computer-readable program codes.

[0133] Although the embodiments of the present invention have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention.

Claims

1. A satellite attitude control method when gyro data is unavailable, characterized in that: include: When it is determined that the gyro data of the satellite is unusable, determining whether the star-sensing data collected by the star sensor configured by the satellite in the current control cycle is valid, and obtaining a first determination result; When the first judgment result is yes, determining the attitude data of the satellite at the current moment according to the star-sensing data in the current control period collected by the star sensor; When the preset conditions are met, the satellite is attitude controlled according to the attitude data of the satellite at the current moment.

2. The satellite attitude control method when gyro data is unavailable according to claim 1, characterized in that: Also includes: When the first judgment result is no, judging whether the data of the current control period collected by the earth sensor configured by the satellite and the data of the current control period collected by the sun sensor configured by the satellite are both valid, and obtaining a second judgment result; When the second judgment result is yes, the attitude data of the satellite at the current moment is determined according to the data of the current control period collected by the earth sensor and the data of the current control period collected by the sun sensor configured by the satellite.

3. The satellite attitude control method when gyro data is unavailable according to claim 2, characterized in that: Also includes: When the second judgment result is no, the attitude data of the satellite at the current moment is determined according to the satellite attitude data within a preset historical time period cached by the satellite.

4. A satellite attitude control method when gyro data is unavailable according to any one of claims 1 to 3, characterized in that: Also includes: Whether the preset condition is met is determined based on the fault diagnosis result that causes the satellite's gyro data to be unusable and the duration of the fault.

5. A satellite attitude control system when gyro data is unavailable, characterized in that: It includes a judgment module, a posture data acquisition module and a posture control module; The judgment module is used to: when it is determined that the gyro data of the satellite is unusable, judge whether the star-sensing data collected by the star sensor configured by the satellite in the current control cycle is valid, and obtain a first judgment result; The attitude data acquisition module is used to: when the first judgment result is yes, determine the attitude data of the satellite at the current moment according to the star-sensing data in the current control period collected by the star sensor; The attitude control module is used to: when a preset condition is met, perform attitude control on the satellite according to the attitude data of the satellite at the current moment.

6. A satellite attitude control system when gyro data is unavailable according to claim 5, characterized in that: The judgment module is also used to: when the first judgment result is no, judge whether the data of the current control period collected by the earth sensor configured by the satellite and the data of the current control period collected by the sun sensor configured by the satellite are both valid, to obtain a second judgment result; The attitude data acquisition module is also used to: when the second judgment result is yes, determine the attitude data of the satellite at the current moment according to the data of the current control cycle collected by the earth sensor and the data of the current control cycle collected by the sun sensor configured by the satellite.

7. A satellite attitude control system when gyro data is unavailable according to claim 6, characterized in that: The posture data acquisition module is also used for: When the second judgment result is no, the attitude data of the satellite at the current moment is determined according to the satellite attitude data within a preset historical time period cached by the satellite.

8. A satellite, characterized in that: The invention comprises an onboard processor, wherein the onboard processor is used to execute a satellite attitude control method when gyro data is unavailable as described in any one of claims 1 to 4.

9. An electronic device, characterized in that: The invention comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein when the processor executes the computer program, a satellite attitude control method when gyroscope data is unavailable as described in any one of claims 1 to 4 is implemented.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the method for controlling a satellite attitude when gyroscope data is unavailable as described in any one of claims 1 to 4 is implemented.

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