2600results about "Spacecraft guiding apparatus" patented technology

The invention belongs to the technical field of aircraft navigation, guidance and control, and provides a satellite exploration control system and method based on deep reinforcement learning, and the system comprises a sensing layer, a decision layer, and an execution layer. The sensing layer is used for integrating equipment data and performing primary processing; the decision-making layer is used for deploying a deep reinforcement learning model and regularly carrying out data training and model updating; the execution layer is used for controlling satellite attitude and orbit adjustment and performing autonomous fault diagnosis and recovery operation; according to the method, effective control under different environments and emergencies is realized through continuous learning and optimization, the adaptive capacity and the autonomous control capacity are relatively high, and an efficient and accurate control strategy can be realized in a complex and changeable space environment through gradual optimization and iteration, so that the method is suitable for large-scale popularization and application. And the autonomous operation capability and the task execution efficiency of the satellite are greatly improved.

The invention discloses a satellite group orbit pursuit decision-making method based on multi-near-end reinforcement learning. The method comprises the following steps: S1, updating a satellite state by adopting an eight-order Runge-Kutta method; s2, all satellites in the satellite group collect state information of each other, a state vector is constructed, and a strategy and value network is formed according to the state vector; selecting and executing an action through a policy network, and updating a state vector after receiving an environment reward; calculating a time sequence difference error and a dominant function, and updating the network by using a near-end strategy optimization algorithm; meanwhile, fusing the strategy networks of the satellites by adopting a federal near-end algorithm; and S3, carrying out comprehensive performance evaluation on the satellite group, dynamically adjusting an updating strategy of each satellite strategy network parameter, and continuously optimizing the overall efficiency of the satellite group. According to the method, efficient cooperation of the satellite groups in a complex environment is ensured, the overall response speed and decision accuracy are improved through continuous learning and adjustment, and efficient and stable task execution is ensured.

The invention provides a lunar satellite formation autonomous navigation method based on inter-satellite ranging information. The lunar satellite formation autonomous navigation method comprises three parts of orbit dynamics model establishment, collaborative observation model establishment and navigation filtering algorithm data fusion. The invention provides an autonomous navigation method based on inter-satellite distance measurement and angle measurement information, which establishes a high-precision dynamic model containing 165-order moon non-spherical perturbation for moon satellite formation. The system can operate independently of a ground station, and high-precision autonomous navigation of the lunar satellite formation is realized only by using on-orbit observation data.

An example method executed by a controller onboard a spacecraft generates a mission plan in real-time that guides the spacecraft along a space autonomous mission to rendezvous with two or more orbiting target objects. The method includes establishing potential permutations for all possible unique maneuver sequences in which to visit the two or more orbiting target objects, and for each potential permutation, determining a collision-free maneuver plan for defining orbit trajectories to intercept each of the two or more orbiting target objects for each of the possible unique maneuver sequences. An optimal permutation is determined that meets viewing constraints of the two or more orbiting target objects, a viewing priority, and fuel constraints of the spacecraft. A mission visit plan is generated using the optimal permutation, and the spacecraft executes the plan to rendezvous with and inspect the two or more orbiting target objects.

Intelligent optimization of mission planning and execution for spacecraft through intelligent analysis of mission objectives and payload operation is disclosed. Values for mission objectives, payload operation, and / or vehicle conditions may be defined during given time periods, including variable values for repeat operations or other factors. These values may then be used through a vehicle / mission simulation agent to optimize the mission plan through maximizing the mission score according to the values. The simulation agent may identify possible mission plans based on vehicle capability, and then score the mission plan based on the total objective score.

The invention provides a lunar satellite formation orbit determination method and system. The method comprises the following steps: S1, establishing a lunar satellite high-precision dynamic model; s2, establishing a lunar satellite laser ranging and inter-satellite ranging mathematical model; s3, measuring the distance, namely the absolute position, of the A-type satellite relative to the ground station through laser ranging, and measuring the relative position of the B-type satellite relative to the A-type satellite through inter-satellite ranging; s4, preprocessing the absolute position and relative position data of the satellite measured by laser ranging and inter-satellite ranging, and generating a format data file required by orbit calculation; and S5, resolving the position coordinates of each satellite in the monthly formation through the format data file. According to the method, laser ranging and inter-satellite ranging are combined for orbit determination, precise orbit determination of the lunar satellite formation is achieved, compared with a traditional orbit determination mode, the precision can be improved from the hundred-meter level to the meter level, the requirements for the arrangement number of ground stations and set distribution are reduced, and the measurement and control pressure of existing ground stations is relieved.

The invention relates to the field of satellite operation, and discloses an in-orbit collision avoidance method and device for a constellation system, and the method comprises the steps: enabling a ground system to obtain the ephemeris information of a constellation satellite and space debris; the ground system evaluates the first collision risk of the satellite in the first period time in the future according to the ephemeris information of the constellation satellite and the space debris; when the first collision risk is greater than a first threshold value, the ground system uploads ephemeris information of the space debris to a satellite; the satellite updates a second collision risk in a second period time in the future in real time based on the received space debris ephemeris information and the satellite positioning information; when the second collision risk is greater than a second threshold value, the satellite formulates a corresponding collision avoidance strategy, and orbital maneuver is carried out according to the collision avoidance strategy; and after the avoidance is completed, the satellite actively maneuvers and returns to a nominal position corresponding to satellite ephemeris information, and the connection with the ground system is recovered. According to the invention, the accuracy of collision risk assessment is improved, and the implementability is improved.

The invention discloses a service spacecraft multi-orbit maneuvering strategy approaching observation method based on J2 perturbation compensation, solves the problem of how to quickly improve orbit maneuvering accuracy, and belongs to the field of service spacecraft multi-orbit maneuvering strategy approaching observation. Comprising the following steps: constructing a motion equation under J2 item perturbation compensation under a target orbit coordinate system; constructing a service spacecraft multi-orbit maneuvering strategy model through the relative motion propagation relational expression and the motion equation; inputting the termination moment position conforming to the termination moment distance constraint into the model to output a single speed increment, and reserving the single speed increment conforming to the termination moment illumination angle constraint and the single speed increment constraint at the same time to obtain a speed increment solution set; constructing a circular region which takes the space target as a center and has a radius meeting a termination moment distance constraint; when the service spacecraft arrives at the circular area at a specified moment, marking the task successfully, and obtaining a service spacecraft multi-orbit maneuvering strategy for approaching observation; according to the invention, accurate and rapid track maneuvering is realized.

The invention relates to a propulsion system, in particular to an air-breathing radio frequency ion electric propulsion system with mixed working medium power and a multi-mode working method thereof, and aims to solve the problems that an existing air-breathing electric propulsion system only takes atmosphere as a single working medium, the atmosphere trapping efficiency is low, and the performance of a spacecraft is affected. The device comprises a propelling and gas trapping unit, a radio frequency neutralizer, a power and control unit and a xenon storage and supply unit, the propelling and gas trapping unit comprises a radio frequency ion thruster and an atmosphere trapping gas inlet channel; the power and control unit comprises a direct current module, a control module and a radio frequency module; the xenon storage and supply unit comprises a xenon bottle and a valve and sensor assembly; the control module and the valve are electrically connected with the sensor assembly. The radio frequency ion thruster is electrically connected with the radio frequency neutralizer through the radio frequency module. The first output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency neutralizer, and the second output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency ion thruster.

The invention discloses an agile satellite attitude fusion method and system considering multi-source reference difference, and belongs to the field of remote sensing satellite ground data processing, and the method comprises the steps: calculating a thermal environment factor based on satellite measurement data and satellite sensor installation parameters; extracting a space-time low-frequency error based on observation data of the star camera and the star sensor; performing normalization processing on the thermal environment factor, and constructing a relation model of the space-time low-frequency error and the thermal environment factor by adopting a multi-layer perceptron network to obtain compensated attitude data of the star sensor and the star camera; establishing a multi-source absolute attitude sensor data fusion model based on the compensated attitude data of the star sensor and the star camera; and carrying out absolute and relative attitude data fusion filtering processing on the processed gyroscope data and the fusion result of the star sensor and the star camera by adopting a bidirectional dynamic unscented Kalman filter to obtain a high-precision attitude data result. According to the method, the attitude determination precision of the agile satellite is improved, and the requirement of a high-precision remote sensing task is met.

The invention relates to a system and method for autonomous navigation of a host satellite equipped with moving and orienting means, a unit for controlling these means, and at least one on-board image-acquisition camera, said method comprising the following steps: acquiring (E1) a plurality of images; default processing of said images, referred to as long-range processing (E2), configured to detect and identify space objects and to calculate their relative orbits; conditional processing of said acquired images, referred to as short-range processing (E3), configured to estimate the attitude of at least one of said space objects, referred to as target object, detected during the long-range processing, this step being implemented when said long-range step detects at least one space object located at a distance estimated to be less than a predetermined threshold distance; determining (E4) a possible rendezvous between at least said target object and the host satellite; preparing and transmitting instructions (E5) to said control unit of said moving means based on at least one rendezvous and / or risk of collision determined in the previous step.

The invention discloses an ultra-low orbit guarantee method, system and equipment for a near-earth satellite. Comprising the following steps: determining a satellite orbit element at the current moment based on an atmospheric resistance perturbation item, a conventional perturbation item and a small thrust perturbation item, and inputting the satellite orbit element at the current moment into an orbit mechanical model; based on the orbit mechanical model and the satellite orbit elements, determining a first velocity position parameter of the satellite at the prediction moment under the ICRS coordinate system, and converting the first velocity position parameter into an ECEF coordinate system to obtain a second velocity position parameter at the prediction moment; inputting the time characteristic, the satellite state characteristic, the space environment characteristic and the second speed position parameter of the current moment into an orbit correction model to obtain a target error prediction value; adjusting the second speed position parameter based on the target error prediction value to obtain a target speed position parameter of the satellite at the prediction moment; and performing orbit attenuation risk early warning based on the target speed and position parameters of the satellite at the prediction moment.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

Disclosed are systems and methods for a distributed space transportation network. Satellite launches to orbit are more efficiently performed by large rockets. Modern satellites are in smaller form factor, leaving the large launch rockets with excess capacity. Small satellite operators can use ride-shares, but do not have efficient options for delivering their satellites to their desired destination and may be forced to operate their satellites in compromise orbits. The disclosed distributed space transportation network maintains a fleet of space tugs, which can dock with satellites in space at an initial arrival destination and transport the satellites to their final destinations. In one embodiment, the space tugs can dock with satellite depots to obtain fuel and repairs.

The invention provides a maintenance-free dual-mode attitude control power system which is used for being installed on an aerospace vehicle and comprises a positive feedback circulation system, a thruster set, a fuel gas generator, a fuel storage box and an oxidizing agent storage box. The positive feedback circulation system is used for self-pressurization of the attitude control power system; the positive feedback circulation system comprises a gunpowder starter, a differential storage tank, a liquid path pressure stabilizing valve, a fuel gas generator and a pressurizing one-way valve; the thruster group comprises a gas thruster, a single-component thruster and a double-component thruster; an inlet of the fuel gas generator is communicated with the positive feedback circulation system; the outlet is communicated with the gas thruster, the fuel storage tank and the oxidant storage tank; the fuel tank and the oxidant tank are in communication with the thruster set. The advantages of a single-component engine and a double-component engine are organically combined, the overall performance of the power system is improved, the use and maintenance requirements of the attitude control power system before rocket launching are lowered, and the launching preparation time is shortened.

The present disclosure relates to a MEMS inertial sensor device in a semiconductor chip package that includes an integrated circuit configured to process inertial sensor data. Preferred implementations utilize inertial sensors having different sensitivity ranges to adjust operation of dynamic system control such as motion and or attitude control of autonomous vehicles.

A modular satellite system is provided comprising a main body member, one or more controllers, a communication system, a datastore, a power unit, and one or more orbital cameras. The controller(s) may be operable to perform a mission instruction. The communication system may be in communication with the controller(s) and with a client terminal. The datastore may store data accessible by the controller(s). The orbital camera(s) may be operable to capture one or more image(s) associated with the mission instruction, which may be defined as captured image(s). The controller(s) may be operable to detect and identify one or more predetermined object(s) in the one or more captured image(s). The controller(s) may generate a mission analytics packet, and the controller(s) may be operable to transmit the mission analytics packet to the client terminal.

Systems, apparatus, and method for momentum management for a spacecraft are disclosed. An example spacecraft includes a body; an appendage moveably coupled to the body; a thruster carried by the body; machine-readable instructions; and processor circuitry to execute the machine-readable instructions to detect a change in angular momentum at the spacecraft during firing of the thruster; and in response to the detection, cause the appendage to move relative to the body to adjust a center of mass of the spacecraft relative to a net thrust vector associated with the thruster.

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

The invention provides a satellite orbital transfer control method and device, electronic equipment and a storage medium, and relates to the technical field of satellite control. The satellite orbital transfer control method comprises the following steps: constructing a three-dimensional centralized monitoring area corresponding to a target satellite based on acquired first operation information of the target satellite; based on the three-dimensional centralized monitoring area, generating a monitoring instruction, and based on the monitoring instruction, controlling a monitoring terminal corresponding to the target satellite to perform centralized monitoring on the three-dimensional centralized monitoring area; when it is monitored that the target terminal exists in the three-dimensional centralized monitoring area, second operation information corresponding to the target terminal in a unit time period is obtained, and a control mode corresponding to the target satellite is determined based on the first operation information and the second operation information; and based on the control mode, controlling the target satellite to perform orbital transfer operation. According to the invention, the orbital transfer capability of the target satellite can be improved and the collision probability can be reduced.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapid accompanying flying, fly-around flying and sweeping flying of an elliptical orbit coplanar target, and relates to a fuel-saving control on-satellite autonomous solving method for rapid approaching and high-precision accompanying flying / fly-around flying / sweeping flying from a long-distance position on the same orbit surface. Comprising the following steps: providing a three-time planning and five-time maneuvering long-distance approaching accompanying flight / fly-around / grazing flight configuration forming method; the efficient iteration method is suitable for solving the initial value of a minimum fuel control analytical solution and the final value of an iterative precise solution of a high-precision orbit extrapolation model through a satellite autonomous solution two-body model. The method is suitable for a near-earth space random elliptical orbit satellite to quickly approach a long-distance target on an approximate same orbit to carry out accompanying flight / fly-around / grazing flight, and the control precision can reach a ten-meter level; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The invention belongs to the technical field of aerospace, and relates to an unscented batch processing orbit determination method and system for a continuous low-thrust maneuvering satellite. The method comprises the following steps: acquiring a radar observation vector according to radar measured data or satellite ephemeris data in combination with radar site coordinates; obtaining an initial orbit state at a to-be-estimated moment by adopting a Lambert algorithm according to the radar observation vector; generating an initial expansion state quantity according to the initial orbit state at the to-be-estimated moment; a simplified orbit kinetic equation containing earth non-spherical perturbation, atmospheric resistance perturbation and constant small thrust acceleration under an RTN coordinate system is constructed under the ECI coordinate system; based on the simplified orbit kinetic equation, obtaining an expansion state quantity precision value of the to-be-estimated moment through the initial expansion state quantity, the radar observation vector and an unscented batch processing filtering method; and substituting the precision value of the expansion state quantity of the to-be-estimated moment into the simplified orbit kinetic equation to carry out orbit forecasting at other moments. According to the method, the orbit determination precision is remarkably improved, and the influence of single-point observation errors is reduced.

According to the double-pulse co-switching optimal orbital transfer method for the lunar large elliptical orbit, firstly, a double-pulse co-switching orbital transfer control strategy based on a two-body nonlinear model is calculated according to the known initial state of the lunar orbit and the state of a target orbit, and through two times of pulse control in the tangential direction, the optimal orbital transfer control strategy is obtained; meanwhile, the average semi-major axis, the average eccentricity rate and the average latitude argument of the large elliptical orbit are adjusted, the speed increment minimum problem is converted into the one-dimensional nonlinear function minimum problem to solve the optimal orbital transfer strategy, and accurate transfer between any two elliptical orbits is achieved; then, a two-body orbit model is designed as an initial value, numerical integration is carried out on the average orbit elements under the influence of perturbation force, the two-body model is corrected, and then a real orbital transfer control strategy of the average orbit parameters under the influence of perturbation force is obtained; the problem that the problem becomes a time-varying problem and is difficult to solve due to the fact that the short-cycle term introduced by directly adopting instantaneous orbit element description is solved, and the engineering practical value of the algorithm is improved.

The present disclosure is a device and methods for satellite trajectory control and collision avoidance. Embodiments of the disclosure are comprised of a smart satellite device performing a process three steps. First, sensors collect data about the satellite's physical landing environment, passing information to satellite's database and processors. Second, the processors manipulate the information with a deep reinforcement learning program to produce instructions. Third, the instructions steer the satellite body by manipulating the satellite's panels for optimal trajectory and collision avoidance. The purpose for the present disclosure is to help solve the space debris problem.

The invention provides a low-orbit navigation constellation autonomous orbit determination method and system under a weak inter-satellite topological structure, and relates to the technical field of satellite orbit determination, and the method comprises the steps: obtaining various observation data of an inter-satellite link on a predetermined orbit, and carrying out the preprocessing of the various observation data; establishing a synchronous observation equation of orbit and clock correction parameters by adopting a segmented linear polynomial satellite clock correction modeling method, generating orbit, clock correction and related error parameters combining satellite-borne GNSS and inter-satellite link data, and forecasting; based on an error analysis result, multiple key features are extracted from the generated orbit, clock error and related error parameters to serve as input of a parameter prediction model, and a parameter time sequence of four types of prediction including relative weight, inter-satellite link hardware delay, low-orbit satellite clock error and unmodeled error is obtained through output; and uploading the predicted parameter time sequence and the broadcast ephemeris to a main satellite together, and transmitting to each sub-satellite by the main satellite through an inter-satellite link to realize distributed autonomous orbit determination of the sub-satellites.

The invention discloses a method for constructing and screening orbital transfer speed increment required by on-orbit service of a pulse thrust spacecraft based on a CW equation, belongs to the field of on-orbit service of spacecrafts, and solves the problem of low precision of a maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft. Comprising the following steps: constructing a relative motion equation of a pulsed thrust spacecraft relative to space debris based on a CW equation; inputting the orbital transfer state transfer matrix and the speed transfer matrix into the relative motion equation at the same time to obtain an on-orbit service orbital transfer model; setting a termination state distance constraint, a termination state illumination angle constraint and an orbital transfer speed increment constraint according to task requirements; and inputting all the termination moment position vectors conforming to the termination state distance constraint into the on-orbit service orbital transfer model, and reserving orbital transfer speed increments conforming to the termination state illumination angle constraint and the orbital transfer speed increment constraint to obtain a screened orbital transfer speed increment solution set. According to the invention, the precision of the maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft is improved.

The invention discloses a spacecraft executing mechanism fault diagnosis and online calibration device and method based on double sensors, and belongs to the technical field of aircraft control and ground simulation. In order to carry out detection and off-line compensation on errors, the device comprises an absolute grating encoder, an actuating mechanism expansion shaft, an inductosyn, a control moment gyroscope, a photoelectric reading head and a signal amplification circuit, the absolute grating encoder, the inductosyn and the control moment gyroscope are coaxially installed through the executing mechanism expansion shaft, one end of the executing mechanism expansion shaft is connected with the absolute grating encoder, the other end of the executing mechanism expansion shaft is connected with the inductosyn, and the inductosyn is connected with the control moment gyroscope. The absolute grating encoder is connected with a first processing circuit box through a photoelectric reading head wire, and the first processing circuit box is connected with an upper computer. The inductosyn is connected with the signal amplification circuit and the excitation power supply through wires, the signal amplification circuit is connected with the second processing circuit box through a wire, and the second processing circuit box is connected with the upper computer.

The invention discloses an electric propulsion GEO satellite orbit transfer and point fixing method, and belongs to the technical field of satellite orbit dynamics, and the method comprises the following steps: S1, building a low-thrust satellite orbit dynamics model based on multi-perturbation and earth shadow constraints; s2, according to the low-thrust satellite orbit dynamic model, a Lyyunov function is constructed, and an RQ-Law feedback control model is constructed by using the Lyyunov function; s3, weight parameters of the RQ-Law feedback control model are optimized; and S4, using the optimized RQ-Law feedback control model to complete orbit transfer and point fixing. Aiming at the problem of long time of current satellite orbit transfer, the invention provides an innovative method, the traditional orbit transfer and fixed point staged optimization are integrated into one stage, and the integration of orbit transfer and fixed point is realized.

The invention discloses a full-process closed-loop quick response satellite collision avoidance method, which comprises the following steps of: 1, observing and acquiring an orbit state of a space object, and then carrying out multi-source data fusion so as to construct a space object database; 2, screening out high-risk collision sources which have the possibility of collision with the own satellite; 3, obtaining a determined high-risk collision source according to the track state historical record of the possible high-risk collision source and the own track state; 4, generating an initial evasion maneuvering scheme, and optimizing the initial evasion maneuvering scheme through a multi-objective optimization algorithm; and obtaining a final evading maneuvering scheme for subsequent execution. And 5, according to the engine startup and shutdown time and the thrust vector contained in the final evasion maneuver scheme, the satellite engine is controlled to work with the specified thrust vector within the given startup and shutdown time, and collision evasion maneuver is executed. Compared with a conventional method, the method has the advantage that the response speed is higher.

Techniques for maneuvering a space vehicle are presented. The techniques can include: obtaining a representation, in a computer, of a multi-segment planned position, planned velocity, and planned acceleration of the space vehicle along a planned continuous trajectory; tracking, using at least one navigation sensor, an indication of an actual position, actual velocity, and actual acceleration of the space vehicle; and maneuvering the space vehicle, using a closed-loop controller, and based on the tracking, to return to the planned position, planned velocity, and planned acceleration.