18 results about "Satellite attitude control system" patented technology

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.

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 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 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.