28 results about "Satellite attitude control system" patented technology

The invention provides a satellite attitude control system fault diagnosis and fault-tolerant reconstruction method and system, and relates to the technical field of satellites, and the method comprises the steps: obtaining attitude data through a Kalman filtering state observer, building a fault diagnosis model, generating a residual sequence, and carrying out the adaptive dynamic threshold calculation to obtain a fault feature vector; and inputting the feature vector into a pre-trained deep belief network for fault diagnosis, and determining the fault type and degree. According to a diagnosis result, a mixed integer programming model considering attitude control precision and actuator power constraint is established, an improved genetic algorithm is adopted for solving, an actuator reconstruction scheme and controller parameters are obtained, and fault-tolerant control is achieved. According to the invention, faults can be diagnosed rapidly and accurately, fault-tolerant reconstruction is carried out according to fault types and degrees, the reliability and stability of a satellite attitude control system are effectively improved, and the satellite attitude control precision is ensured.

The invention discloses a satellite attitude control system fault detection method and device and a medium, and the method comprises the steps: introducing a diffusion model to carry out the preprocessing of an original data set before a fault diagnosis task is executed through employing a neural network, carrying out the focusing training of an efficient diffusion model, generating scarce minority class samples through employing the model, and carrying out the fault diagnosis of a satellite attitude control system. The problem of category imbalance generally existing in satellite attitude control system data is effectively solved, and the strong data processing capacity of the diffusion model is also utilized. Based on the balanced training set, the adaptive one-dimensional convolutional neural network is further combined, complex time sequence data is efficiently processed, key fault information is mined, and the accuracy and reliability of fault diagnosis of the satellite attitude control system are improved.

The invention provides an attitude control system for a short-wave satellite, and relates to the technical field of data processing, and the attitude control system comprises the steps: inputting a three-axis synthesis control torque instruction into a reaction flywheel controller, converting the instruction into a flywheel motor rotating speed adjustment instruction, and driving a reaction flywheel to execute; adjusting and analyzing the continuously updated attitude deviation angle based on the actual attitude change after the execution of the driving reaction flywheel, obtaining the updated gain parameter of the sliding mode controller, and feeding back the parameter of the updated sliding mode controller; performing control based on the updated parameters of the sliding mode controller, performing direction consistency verification on a newly output sun direction vector measurement value, and verifying an actual included angle; if the actual included angle is smaller than the preset precision threshold value, a new round of data acquisition cycle is triggered, and closed-loop control is achieved. According to the invention, the timeliness and effectiveness of attitude control are ensured, and the control efficiency and reliability are improved.

The present invention relates to the technical field of spacecraft fault diagnosis, and specifically discloses a neural adaptive sliding mode observer design method for satellite attitude control system fault estimation, including taking into account concurrent actuator, sensor faults and constraints, performing equivalent transformation on the attitude control system; establishing a continuous time constrained equivalent uncertainty description system; adding a fuzzification layer and a fuzzy reasoning layer to a hyperbasis function neural network to obtain an expression for uncertainty obtained by the fuzzy hyperbasis function neural network; designing a neural adaptive sliding mode observer based on an improved fuzzy hyperbasis function neural network, solving the observer gain matrix using linear matrix inequalities, and achieving better concurrent fault estimation based on better uncertainty estimation. The present invention adopts the above-mentioned neural adaptive sliding mode observer design method for satellite attitude control system fault estimation, solves the concurrent fault estimation problem of the satellite attitude control system under complex constraints, and improves the accuracy of fault estimation.

The invention provides a satellite attitude control system fault prediction method based on data driving. The satellite attitude control system fault prediction method comprises the steps of collecting historical fault time sequence data of key parameters of a satellite attitude control system; dividing the data into to-be-predicted parameter historical data and other influence parameter historical data; compressing the historical data of the to-be-predicted parameter by using a mean value down-sampling method and performing preprocessing; historical data of other influence parameters are preprocessed, and feature data, closely related to the to-be-predicted parameters, of the historical data are extracted through an ECA-COV1D module; and inputting the mean value down-sampling processing and the parameter sequence extracted by the ECA-COV1D module into the LSTM network to predict the future state of the satellite attitude control system. According to the method, the amount of data input into the LSTM network can be effectively reduced, meanwhile, correlation and differences among other sensor parameters can be better reflected, the prediction precision is improved by adding the auxiliary loss function, and therefore the purpose of short-term fault prediction is achieved.

The invention discloses a spacecraft attitude control single-machine model selection method and system and computer equipment. The spacecraft attitude control single-machine model selection method comprises the steps that disturbance torque is calculated based on a disturbance torque model; torque data of the disturbance torque and the suppression capability of the attitude control single machine on the disturbance torque are obtained, wherein the torque data comprise the torque distribution condition, the torque size and the torque direction; selecting a corresponding attitude control single machine combination and single machine control parameters by utilizing an optimization algorithm on the basis of spacecraft attitude control requirements, torque data and the suppression capability of attitude control single machines on disturbance torque; and based on the attitude control single-machine combination, the corresponding single-machine control parameters and numerical simulation, simulating the operation attitude of the spacecraft in different orbits and space environments, and selecting the attitude control single-machine combination according to a simulation result. According to the method, the precision and the stability of the satellite attitude control system can be remarkably improved, and the possible uncertainty in a traditional model selection method is reduced.

The application belongs to the technical field of storage batteries, and discloses a residual magnetic moment eliminating structure of a storage battery pack, which comprises a first battery pack and a second battery pack symmetrically arranged with an XY plane as a mirror plane, so that the current paths on the first printed circuit board and the second printed circuit board are consistent, the magnetic ring areas are equal, the current sizes are equal, and the current directions are opposite, therefore, the dynamic residual magnetic moments in the XY plane generated by the currents on the first printed circuit board and the second printed circuit board can be offset, so that the residual magnetic moment of the storage battery pack is effectively reduced, the external magnetic interference of the storage battery pack is reduced, the continuous disturbance of the residual magnetic moment on a satellite attitude control system is avoided, the satellite attitude control precision is significantly improved, the frequency of high-frequency magnetic unloading operation of a magnetic moment device is reduced, the satellite energy loss is reduced, and the in-orbit service life of the satellite is effectively prolonged.

This invention provides a novel integrated attitude sensor for satellites, relating to the fields of onboard attitude control and electronics. It includes: a star sensor body; a solar sensor body mounted on the star sensor body; a magnetometer body within the star sensor body; and a gyroscope body within the star sensor body. By integrating traditional discrete structures into a single device, the size and weight are reduced, and costs are significantly lowered. Magnetometer and gyroscope data are fused in real-time within an FPGA. This device features rapid adaptability, modular design supporting plug-and-play functionality, and integrated design improving EMC performance. It solves the problems of existing traditional satellite attitude control systems, where various attitude sensors are independently configured and equipped with dedicated processors, resulting in high hardware redundancy, large overall size and weight, high power consumption and cost, hindering further improvements in attitude measurement accuracy and dynamic performance, and leading to high overall design complexity.

The invention discloses a satellite sensor fault diagnosis method based on EMD (Empirical Mode Decomposition), which comprises the following steps: a data acquisition stage: acquiring historical operation time sequence data of an attitude control system in satellite equipment; in the signal decomposition stage, an empirical mode decomposition algorithm is adopted to perform multi-scale decomposition on the time sequence signals, and a plurality of IMF components are extracted; in the feature visualization stage, energy features of all IMF components are calculated and converted into a two-dimensional energy distribution map; in the intelligent diagnosis stage, a pre-trained deep convolutional neural network model is used for conducting feature extraction and mode recognition on the energy spectrum, and an equipment state diagnosis conclusion is output. According to the invention, the fault diagnosis accuracy of the satellite attitude control system can be improved. The method can be widely applied to the field of satellite equipment fault diagnosis.

The present disclosure provides a satellite attitude control system fault diagnosis method and device based on time-frequency domain feature fusion, belonging to the technical field of satellite fault diagnosis. The method can include: obtaining target satellite data and converting the target satellite data into satellite image data using a Gram angle field; performing feature extraction on the satellite image data to obtain a satellite feature image; extracting global time sequence features of the satellite feature image, wherein the global time sequence features refer to image sequence features extracted from the entire time sequence of the satellite and capable of representing the characteristics of the entire sequence; inputting the global time sequence features into a dense connection model to obtain a fusion time sequence feature map corresponding to the satellite feature image; and obtaining a target fault type corresponding to the target satellite data according to the fusion time sequence feature map. This technical solution can improve the efficiency and accuracy of satellite fault diagnosis.

The application discloses a satellite attitude control system event-triggered STF fault detection method, and belongs to the technical field of detection, which comprises the following steps: establishing a fault model of a satellite attitude control system; introducing an event-triggered mechanism; adopting an STF method to design an event-triggered fault detection filter; obtaining an estimation error dynamic equation and a residual signal through linearization processing of the fault detection filter, realizing complete decoupling of the residual signal and the event-triggered error; converting a design problem of a residual generator into solving a gain matrix of the fault detection filter; calculating a residual evaluation function and comparing the residual evaluation function with a threshold value to judge whether a fault occurs or not. The application utilizes the STF method to design the fault detection filter, can enable the system to keep the tracking ability to a sudden state when reaching a stable state, adopts a brand-new event-triggered framework, realizes complete decoupling of the residual signal and the event-triggered error, and effectively improves the resource utilization rate under the premise of guaranteeing the system performance.

The present application relates to a kind of maximum energy efficiency control method of reaction flywheel group and satellite attitude control system.The method includes providing satellite attitude control system equipped with reaction flywheel group;Reaction flywheel speed saturation protection threshold and reaction flywheel speed upper limit are set, the saturation protection threshold is less than the reaction flywheel speed upper limit, the saturation protection threshold is used to limit the actual speed of reaction flywheel before reaching the reaction flywheel speed upper limit, to extend the time of reaction flywheel group speed to reach the reaction flywheel speed upper limit.The method of the present application can extend the time of reaction flywheel speed to reach upper limit to the maximum, the angular momentum that reaction flywheel needs to compensate is more fully distributed to each reaction flywheel, the angular momentum compensation capacity of reaction flywheel is exerted as far as possible, the maximum angular momentum that reaction flywheel can provide is fully exerted.

The invention provides a reaction wheel reliability prediction method in a degradation process, and the method comprises the steps: collecting the historical degradation data of the bearing temperature of a satellite reaction wheel, and carrying out the preprocessing of the data; establishing a reaction wheel temperature performance degradation model by using a Wiener process and performing parameter estimation; establishing a reaction wheel reliability model and calculating the reliability of the reaction wheel; a Kalman filtering algorithm and the reaction wheel temperature performance degradation model are combined to establish a prediction model of the reaction wheel; and predicting the reliability of the reaction wheel by using the prediction model. According to the invention, a more accurate and reliable fault prediction result can be obtained, so that the purpose of fault prediction of the satellite attitude control system is achieved, and improvement of reliability and performance of the satellite attitude control system is promoted; according to the method, the random item is added in the Kalman filtering algorithm, the advantages of each model and algorithm are fully played, the accuracy and stability of overall prediction are improved, and the problem that the real-time performance is poor in the method for predicting the reliability by using the random degradation process is solved.

The utility model discloses an integrated satellite attitude control system, and relates to the technical field of satellite structures. Comprising a main body support, star sensors, a terahertz sensor, a mechanical gyroscope, a hydraulic device and a connecting rod, the support is of an octahedral hollow structure, three star sensor supports are rotationally connected to the top of the support, the three star sensors are fixedly connected to the three star sensor supports respectively and located at the three corners of an equilateral triangle respectively, and the three star sensors are connected with the hydraulic device. The three star sensors are installed on the main body support, the sensor is fixedly connected to the outer top surface of the main body support and located in the center of the three star sensors, the mechanical gyroscope is installed on the inner top surface of the main body support and located in the center of the three star sensors, the hydraulic device is fixedly connected to the inner bottom surface of the main body support, and the top end of a hydraulic rod of the hydraulic device is rotationally connected with the bottom ends of the three connecting rods respectively. And the top end of each connecting rod is rotationally connected with a star sensor bracket. The star sensor is simple in structure, small in occupied space and capable of effectively saving installation space, and the installation angle of the star sensor is adjustable.

The invention relates to the technical field of spaceflight, in particular to a cubesat attitude control support assembly and a cubesat attitude control system.The cubesat attitude control support assembly comprises a first support and a second support, and the first support comprises a first base plate and a first mounting plate perpendicularly fixed to the first base plate; the first support comprises a first substrate, the second support comprises a second substrate parallel to the first substrate, the first mounting plate is fixedly connected with the second support so as to form an attitude control assembly accommodating space between the first substrate and the second substrate, and the first substrate, the first mounting plate and the second substrate are all used for mounting attitude control assemblies. The invention aims to provide a cubesat attitude control bracket assembly and a cubesat attitude control system aiming at at least one technical problem related in the background technology.

The application provides a reaction wheel reliability prediction method under degradation process, comprising collecting satellite reaction wheel bearing temperature history degradation data, and preprocessing the data; using Wiener process to establish a reaction wheel temperature performance degradation model and performing parameter estimation; establishing a reaction wheel reliability model and calculating the reaction wheel reliability; combining Kalman filtering algorithm and the reaction wheel temperature performance degradation model to establish a prediction model of the reaction wheel; and using the prediction model to predict the reliability of the reaction wheel. The application can more accurately and reliably predict the fault result, so as to achieve the purpose of satellite attitude control system fault prediction, promote the reliability and performance improvement of the satellite attitude control system; the application adds a random term in the Kalman filtering algorithm, fully gives play to the advantages of each model and algorithm, improves the accuracy and stability of the overall prediction, and improves the poor real-time problem in the reliability prediction method using the random degradation process.

The invention discloses a dual-redundancy commutation control method of a reaction flywheel and a control system of a reaction flywheel system, relates to the field of aviation, effectively overcomes the technical bottleneck of'failure of a Hall element, namely downtime 'in an existing flywheel system, and does not need to add hardware modification and test. And dual-redundancy commutation control of the Hall sensor and the incremental encoder can be realized only through calibration of software. If the Hall sensor operates normally, the reaction flywheel is controlled according to the commutation logic of the Hall sensor; if the Hall sensor fails, whether the reaction flywheel is restarted or not is further judged, and if the reaction flywheel is not restarted, the reaction flywheel is controlled according to phase change logic of an incremental encoder; and if restarting, executing the zero-searching operation until the zero-searching is finished, quitting the zero-searching operation, and controlling the reaction flywheel according to the commutation logic of the incremental encoder. The method is suitable for processing the reaction flywheel of the core execution mechanism of the satellite attitude control system.

The utility model discloses an attitude control integrated structure based on 1U size, and relates to the technical field of satellite structures. Comprising a structure unit, an attitude control subsystem, an attitude monitoring subsystem, a data processing subsystem, a data transmission communication subsystem and a power supply subsystem, the structure of the unit is a hollow hexahedron structure formed by a front mounting plate, a rear mounting plate, a left mounting plate, a right mounting plate, an upper mounting plate and a lower mounting plate, and the six mounting plates are all hollow structures. The front installation plate, the left installation plate and the lower installation plate form a first integrated installation plate, the upper installation plate and the right installation plate form a second integrated installation plate, and the first integrated installation plate, the second integrated installation plate and the rear installation plate are connected through bolts. The attitude monitoring subsystem, the attitude control subsystem, the data processing subsystem, the data transmission communication subsystem and the power supply subsystem are all attached to the unit structure. The satellite attitude control system is simple in structure, low in cost and capable of adapting to integrated attitude control of most satellites.

The invention relates to a maximum energy efficiency control method for a reaction flywheel set and a satellite attitude control system. The method comprises the following steps: providing a satellite attitude control system provided with a reaction flywheel group; a saturation protection threshold and a reaction flywheel rotation speed upper limit of the reaction flywheel rotation speed are set, the saturation protection threshold is smaller than the reaction flywheel rotation speed upper limit, and the saturation protection threshold is used for limiting the actual rotation speed of the reaction flywheel before the reaction flywheel reaches the reaction flywheel rotation speed upper limit; therefore, the time that the rotating speed of the reaction flywheel set reaches the upper limit of the rotating speed of the reaction flywheel is prolonged. According to the method, the time for the rotating speed of the reaction flywheels to reach the upper limit can be prolonged to the maximum extent, the angular momentum, needing to be compensated, of the reaction flywheels is more fully distributed to each reaction flywheel, the angular momentum compensation capacity of the reaction flywheels is fully played as much as possible, and the maximum angular momentum capable of being provided by the reaction flywheels is fully played.

This invention discloses a method, system, and computer equipment for selecting individual attitude control units for spacecraft. The method includes: calculating disturbance torque based on a disturbance torque model; acquiring torque data of the disturbance torque and the suppression capability of the attitude control unit against the disturbance torque, with the torque data including torque distribution, magnitude, and direction; selecting the corresponding attitude control unit combination and unit control parameters using an optimization algorithm based on the spacecraft's attitude control requirements, torque data, and the suppression capability of the attitude control unit against the disturbance torque; simulating the spacecraft's operating attitude under different orbits and space environments based on the attitude control unit combination, corresponding unit control parameters, and numerical simulation, and selecting the attitude control unit combination based on the simulation results. This invention can significantly improve the accuracy and stability of satellite attitude control systems and reduce the uncertainties that may exist in traditional selection methods.

The invention discloses a microsatellite attitude control system joint simulation method based on a control moment gyroscope, and relates to the field of satellite attitude control system simulation. The method specifically comprises the following steps: building a three-dimensional model of a control moment gyroscope and a satellite shell in SolidWorks; building a solar wing flexible model in the HyperMesh; all the models are imported into Adams to build an Adams satellite virtual prototype model; building a satellite attitude control system model in the Simulink; and finally, through Adams and Simulink data interaction, realizing joint simulation. The method has the advantages that integration of mechanical-control system simulation is achieved, the dynamic performance of the control moment gyroscope is accurately verified, model parameters can be adjusted at will, and the product research and development period is shortened; and meanwhile, expansion, iteration and optimization of a control strategy of the satellite attitude control system can be realized, and simulation support is provided for development of the system.

This invention discloses a hybrid control configuration method for highly maneuverable micro-nano satellites, which uses three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism. This allows the control moment gyroscopes to exert high maneuverability on a single axis while also serving as backup for the three reaction wheels, thereby improving the reliability of the satellite attitude control system.

The invention belongs to the technical field of satellite structure dynamics and vibration measurement, and aims to provide an in-orbit solar wing low-frequency vibration frequency measurement device and measurement method based on video feature point tracking so as to solve the problem of coupling of the low-frequency vibration frequency of a solar wing and the frequency of a satellite attitude control system when a satellite is in orbit. The measuring device comprises a reaction flywheel used for applying vibration excitation to the solar wing, a reflective target installed at the free end of the solar wing, a solar wing monitoring camera used for shooting, a near-infrared light source and a comprehensive electronic system used for data processing. The measuring method comprises the following steps: estimating the amplitude of the solar wing through kinetic analysis, and designing measuring device parameters; afterwards, a flywheel is utilized to excite the solar wing to vibrate in orbit, and a target motion video is shot through a camera; vibration signals are extracted through image processing, and finally the modal frequency is recognized through frequency domain analysis. According to the invention, non-contact accurate measurement of low-frequency vibration of an in-orbit solar wing is realized, and a key basis is provided for formulating a satellite attitude control strategy.

The invention discloses a satellite attitude control system fault diagnosis method based on a dual confrontation mechanism, and relates to the field of fault diagnosis of satellite attitude control systems. A dynamic fusion generator in a double adversarial network established by the method expands domain invariant features and domain specific features of each group of sample images extracted by a variable generation module to generate target domain reconstruction data; the target domain data and target domain sample data jointly form target domain data, fault diagnosis is carried out through a feature extraction module and a state classifier, and adversarial training is carried out on a variable generation module by a domain discriminator to optimize data cross-platform consistency; the generation discriminator carries out adversarial training on the dynamic fusion generator to optimize the data authenticity of the target domain reconstruction data, so that source domain sample data can be utilized to assist in making up the defect of limited in-orbit satellite fault data to carry out network training, and automatic fault diagnosis of the satellite attitude control system of an online satellite is realized. And the method has relatively high accuracy and robustness.

The invention discloses a hybrid control configuration method applied to a high-maneuverability micro-nano satellite, which takes three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, so that the control moment gyroscopes give consideration to the backup function of the three reaction flywheels while exerting high maneuverability in a single axis, and the reliability of a satellite attitude control system is improved.

The present application relates to a kind of flexible solar wing for sun synchronous orbit satellite constellation and on-orbit control method, solar wing is flexible stacked structure, stack on satellite body in retracted state, solar wing includes oppositely arranged single wing, the single wing includes: substrate assembly, it is configured to carry out photoelectric conversion, the substrate assembly includes: rigid substrate and several flexible substrates;Unfolding mechanism, it is configured to provide unfolding stiffness for solar wing, to ensure that the flexible substrate keeps planar form;And unfolding motor, it is configured to drive the unfolding mechanism unfolding or folding, to realize the retraction or unfolding of solar wing;First solar wing driving mechanism and lifting mechanism are provided between oppositely arranged single wing.The present application is overall used flexible stacked structure form retraction envelope size small, light in weight, does not need additional use satellite attitude control system to carry out attitude adjustment and realizes sun tracking.

The application provides a short-wave satellite attitude control system and relates to the technical field of data processing, and comprises the following steps: inputting a three-axis combined control torque instruction into a reaction flywheel controller, converting the three-axis combined control torque instruction into a flywheel motor speed regulation instruction and driving the reaction flywheel to execute; based on actual attitude changes after driving the reaction flywheel to execute, adjusting and analyzing a continuously updated attitude deviation angle to obtain updated sliding mode controller gain parameters and feeding back the parameters of the updated sliding mode controller; based on the parameters of the updated sliding mode controller, controlling, verifying the direction consistency of newly output sun direction vector measurement values, and verifying the actual included angle; if the actual included angle is less than a preset precision threshold, triggering a new round of data acquisition cycle to realize closed-loop control. The application guarantees the timeliness and effectiveness of attitude control and improves the efficiency and reliability of control.

The invention discloses a semi-physical simulation gyro measurement simulation method for a satellite attitude control system, a compensation interface does not need to be added in gyro measurement simulation, and a frame angle of a three-axis turntable can directly correspond to a satellite attitude angle. The invention specifically relates to three aspects. The method comprises the following steps: (1) establishing a three-axis turntable motion equation according to a required attitude rotation sequence; and (2) calculating the angular velocity of the three-axis turntable frame according to the relative inertial angular velocity of the satellite. And (3) performing earth rotation angular velocity compensation calculation in the gyro measured value, and calculating and compensating earth rotation through software to obtain gyro output. According to the method provided by the invention, a hardware compensation interface does not need to be added in gyro measurement simulation, and the adaptability of a satellite semi-physical simulation system to different hardware equipment is improved. And a frame angle of the three-axis turntable can directly correspond to a satellite attitude angle, so that the simulation effect of the system on satellite attitude motion simulation is improved.