131results about "Spacecraft guiding apparatus" patented technology

A thrust generating device according to the present invention is a thrust generating device for irradiating a target with a laser beam to generate thrust for the target, the thrust generating device including: a laser beam generating device that generates a first laser beam having a first wavelength and a second laser beam having a second wavelength different from the first wavelength; and an irradiation device that simultaneously irradiates the target with the first laser beam and the second laser beam. The second wavelength may be a wavelength having a high absorptivity in the target than the first wavelength, and the intensity of the second laser beam may be lower than that of the first laser beam. Further, the second laser beam may be generated by converting the wavelength of the first laser beam.

The application discloses an aircraft constraint attitude control method and system based on explicit reference management, belongs to the field of aerospace, and solves the aircraft attitude maneuver control problem existing multiple constraints (including attitude pointing constraint, angular velocity constraint and control constraint). The application separates constraint processing from control law design, designs a reference manager to generate a reference attitude trajectory, can provide a feasible maneuver trajectory satisfying the constraints and reaching the target attitude from the initial attitude, and designs a feedback control law based on an immersion and invariant disturbance observer to ensure the stability and robustness of the system and realize asymptotic tracking of the reference attitude trajectory. For an aircraft attitude control system with a flywheel as an actuator, the method can realize attitude maneuver control satisfying the pointing constraint, angular velocity constraint and control constraint in the presence of external disturbance. The numerical simulation result of the attitude control method proves the effectiveness and feasibility of the application.

The present application relates to the field of aerospace technology, in particular to a kind of spacecraft autonomous control system, including autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, by setting autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, to realize the full-automatic intelligent navigation of spacecraft in the process of space flight and the autonomous docking of spacecraft and space station, further not need manual intervention, by the autonomous control of spacecraft in the process of space flight, to greatly shorten the working time of astronauts in the small space of spacecraft, improve the work efficiency of astronauts, further solve the technical problems that the control system of existing spacecraft cannot realize automatic control, need manual intervention, to increase the working time of astronauts in the small space of spacecraft, reduce the work efficiency of astronauts.

This invention discloses a Tube model predictive control method for electromagnetic docking in on-orbit refueling, comprising: S1, establishing a spacecraft orbit-electromagnetic-liquid coupled dynamic model and converting it into a discrete state-space form; S2, designing a constraint tightening method based on adaptive robust invariant sets to address disturbance terms in the dynamic model, ensuring recursive feasibility under bounded disturbance conditions; S3, constructing an energy consumption-accuracy multi-objective optimization function, and using a rolling time-domain optimization framework to achieve a dynamic trade-off between energy consumption and docking accuracy; S4, constructing a control Lyapunov function, and using robust invariant sets to ensure the stability of the algorithm under constraints, thereby obtaining the optimal controller; S5, substituting the Tube model predictive controller into the closed-loop system to achieve a stable relative position of the spacecraft, realizing coupled control. This invention can achieve precise electromagnetic docking control under conditions of system time-varying dynamics and coexistence of external and tank liquid disturbances, with low control energy consumption and wide applicability.

The application discloses an anti-unwinding flexible spacecraft attitude tracking composite control method, first, based on the quaternion method, a flexible spacecraft attitude tracking error kinematics and dynamics model is established; then, the flexible accessory vibration mode, external environmental disturbance and model uncertainty are regarded as a lumped disturbance, a generalized proportional integral observer is designed, and the lumped disturbance is observed and estimated; finally, a non-singular terminal sliding mode surface is constructed, the lumped disturbance observation and estimation compensation of the generalized proportional integral observer is taken as a feedforward signal, and a feedforward-feedback anti-unwinding attitude tracking composite controller is designed. Through determining appropriate control gain, the closed-loop attitude tracking error system is globally finite time stable, and the anti-unwinding effect can be realized. The application can realize finite time attitude tracking of the flexible spacecraft, has high attitude tracking precision and strong robustness.

This invention discloses a high-precision periodic orbit maintenance method based on maintenance point rolling optimization, relating to the field of spacecraft orbit control and orbit maintenance technology. Based on a high-precision ephemeris dynamics model, it constructs a maintenance point rolling prediction mechanism, a navigation and execution error model, and a least-squares orbit correction strategy under multi-path point constraints. This application uses the above method to achieve high-precision, low-energy-consumption, and highly robust long-term maintenance of a spacecraft's periodic orbit in a realistic multi-body perturbation environment.

The application relates to a single-axis inertia space orientation attitude control method using magnetic control, which comprises the following steps: performing optimization control according to the current attitude deviation of a satellite; calculating an optimal magnetic control torque under the current attitude; and calculating an output magnetic torque of a magnetic torque device. The single-axis inertia space orientation attitude control method using magnetic control can be used for single-axis inertia space orientation attitude control of a satellite using magnetic control, can calculate an optimal output torque vector of the magnetic control according to the current expected control torque of the satellite and the direction of the current magnetic field vector, and can guarantee that the actual output torque of two control axes is the same as the expected control torque when the satellite is controlled by the magnetic torque device, so that high-precision attitude control can be realized. Meanwhile, an angular velocity damping strategy is designed for a non-control axis, so that the angular velocity of the axis will not be excessively large.

A method and system for energy harvesting for low-inclination orbit satellites, belonging to the field of spacecraft energy management technology. The solar panels are obliquely mounted at a fixed angle θ around the X-axis of the satellite body. Based on a comparison between the real-time orbital angle β and a preset threshold βi, the operating mode is intelligently switched: when |β|>βi, the satellite maintains its Earth-oriented attitude, passively compensating for the oblique angle θ; when |β|≤βi, the satellite performs one-dimensional solar tracking; when θ>0, the roll angle is adjusted by (-θ+|β|)°; when θ<0, the roll angle is adjusted by (-θ-|β|)°; the yaw angle is set to 0° or 180° depending on the polarity of θ and β to compensate for changes in β polarity during the satellite's lifetime, ensuring that the solar panel normal vector and the solar vector are always on the same side of the orbital plane. This invention eliminates the need for a solar panel rotation mechanism. By combining passive compensation with active optimization, it achieves stable and efficient energy supply throughout the entire orbital cycle through small-angle rolling of the satellite body. This significantly reduces satellite costs and control system complexity, and decreases the satellite's windward area, providing significant advantages for long-term on-orbit operation of ultra-low orbit or low orbit satellites.

This invention provides a method and system for attitude maneuver planning with constrained solar vector paths. The method includes: planning a target maneuver path for the current satellite; decomposing the target maneuver path into a first virtual rotation and a second virtual rotation; determining a first rotation angle and a first rotation axis corresponding to the first virtual rotation, and a second rotation angle and a second rotation axis corresponding to the second virtual rotation; optimizing the first and second rotation angles; and controlling the current satellite to complete a maneuver from an initial mission-pointing attitude maneuver to a target mission-pointing attitude maneuver by alternately executing the first and second virtual rotations, while satisfying the constrained solar vector path. This invention effectively achieves the simultaneous satisfaction of solar vector pointing constraints during large-angle attitude maneuvers of the satellite.

The application discloses a kind of asymmetric configuration satellite attitude maneuvering process coupling compensation method, comprising: establishing the satellite attitude dynamics model including the non-diagonal item of moment of inertia matrix;Read current attitude sensor data, and calculate deviation quaternion and deviation angular velocity;According to the attitude dynamics model and current angular velocity data, calculate coupling compensation feedforward control torque;Based on the coupling compensation feedforward control torque, deviation quaternion and deviation angular velocity, the control torque after compensation is calculated;The control torque after compensation is limited and priority torque distribution, and the final control torque is obtained;Final control torque is sent to actuator drive satellite and carries out attitude maneuvering.Solve the problem that coupling effect caused by inertia product cannot be effectively compensated when asymmetric configuration satellite is in large-angle attitude maneuvering in prior art, and then affect control precision and stability.

The present application relates to the field of space technology, in particular to a spacecraft sound wave suspension method, which generates five sound waves through a specific device, controls the frequency, intensity and duration of the sound waves, realizes accurate control of the spacecraft, changes the rotation speed, direction or height of the spacecraft by changing the frequency or intensity of the sound waves, uses five sound waves for navigation, realizes accurate navigation of the spacecraft, monitors the propagation characteristics of the sound waves, the state and position of the spacecraft and other information in real time during the flight of the spacecraft, compares the difference between the expected trajectory and the actual trajectory, adjusts the frequency, intensity and other parameters of the sound waves in time to ensure that the spacecraft can accurately reach the destination, is not limited by the traditional propulsion method, and can realize flexible suspension and movement in various environments. The problem that the fuel weight of the spacecraft occupies a large part of the effective load and limits the voyage and load capacity of the spacecraft is solved.

The application discloses an ultra-low orbit adsorption type air-breathing pressurization system, which is used for an ultra-low orbit electric propulsion system and comprises an adsorption pressurizer on one side and a gas storage on the other side; the adsorption pressurizer is used for adsorbing rare thin gas molecules under low-temperature conditions and releasing the gas molecules under high-temperature conditions, and generates a pressurization effect; and the gas storage is used for storing the pressurized gas molecules and releasing the gas molecules on demand. The application discards the traditional mechanical compression mode and innovatively adopts an adsorption type pressurization principle. Through integrated design of a rotating staggered valve and an adsorption plate group, gas guiding, flow channel switching and adsorption pressurization functions are integrated into one, forming a highly compact overall structure. The application discards the traditional mechanical compression mode and innovatively adopts the adsorption type pressurization principle. The problems of a complex structure, a large size, extremely high power consumption and prominent energy contradiction of the prior art ultra-low orbit electric propulsion system are solved.

The application discloses a kind of distributed spacecraft in-orbit pursuit game control method and system based on vision, which comprises the following steps: S1, the orbit target image sequence is collected using the forward camera of multiple pursuit spacecraft, and the effective target area is screened out;S2, the feature points are extracted using SuperPoint, and the high-confidence matching point pair is output by SuperGlue structure semantic matching;S3, the intrinsic matrix is calculated and the nine-dimensional state vector is generated by fusing the estimation results of multiple spacecraft;S4, the nine-dimensional state vector is used to establish an orbit pursuit game model, and the optimal control law is solved by combining the relative motion dynamics equation to generate the expert strategy trajectory;S5, the trajectory is used to train the control strategy network;S6, the trained control strategy network is deployed to the attitude and orbit control system, and the orbit control instruction is output to form a perception-decision-control closed loop;The application realizes the efficient cooperative control of distributed spacecraft in-orbit pursuit through spacecraft visual perception, feature matching, game modeling and generative adversarial imitation learning.

The application discloses a spacecraft autonomous avoidance of space debris based on a deep deterministic policy gradient algorithm, classifies a scene, constructs different training environments according to different types of orbit threats, constructs a constrained Markov decision process, makes avoidance decisions while satisfying multiple constraints of the spacecraft, in an offline training stage, makes the intelligent agent interact with different types of environments continuously, and learns and trains corresponding neural networks, in an online application stage, first, threat identification is performed, corresponding trained neural networks are extracted, avoidance actions are quickly generated, the spacecraft can learn to make corresponding optimal decisions in different scenes and environment states, and the avoidance of space debris is realized online. The technology can enable the spacecraft to face space debris threats, autonomously formulate avoidance strategies, autonomously complete avoidance actions, autonomously grasp the timing and strength of space debris avoidance, and effectively respond to the space debris threats at the minimum cost (minimize interference with task execution).

This application provides a method, apparatus, electronic device, and storage medium for adjusting the solar array's attitude towards the sun, relating to the field of satellite control technology. The method for adjusting the solar array's attitude towards the sun includes: when the energy output corresponding to the target satellite's position at a preset orbital point is not within a preset energy output range, acquiring the target solar array's attitude information towards the sun and historical attitude information from the previous orbital period adjacent to the current orbital period; determining initial attitude adjustment information based on the target solar array's attitude information and the historical attitude information; determining multiple historical attitude adjustment information corresponding to the initial attitude adjustment information, and determining the target attitude adjustment information based on the multiple historical attitude adjustment information; determining the final solar array's attitude information based on the target attitude adjustment information and the target solar array's attitude information, and controlling the solar array to adjust its attitude based on the final solar array's attitude information.

Attitude control system for satellites, comprising at least two orthogonally arranged, ring-shaped magnetic belt systems, consisting of several segmented electromagnets that can selectively generate electromagnetic fields.

This invention relates to the field of remote sensing satellite orbit control technology, and discloses a satellite autonomous pipeline control method based on differential interferometric SAR imaging missions. The method includes: acquiring initial satellite orbit parameters; employing a multi-objective optimization algorithm for orbit optimization under a high-order gravity field; sampling and constructing a reference orbit point set; calculating orbit control deviations based on the actual orbit and the reference orbit; generating orbit control triggering rules and avoidance strategies through ground system simulation; executing orbit control strategy generation and control onboard system; and evaluating orbit control effectiveness and the evolution trend of orbit control deviations through the ground system. Existing technologies mainly rely on ground command injection, making it difficult for satellites to achieve efficient autonomous orbit control when performing heavy-orbit differential interferometric SAR imaging missions, and hindering the decoupling of orbit control and imaging tasks. This invention improves the efficiency and autonomy of orbit control, the execution efficiency of imaging missions, and the safety of on-orbit operation by constructing a closed-loop control system encompassing satellite-ground collaboration, strategy formulation, command execution, and effect evaluation.

A pipeline deviation full orbit fitting method of a rigorous regression orbit satellite can determine single-point pipeline deviation based on reference trajectory coordinates and cutting track plane alignment according to the characteristics of the rigorous regression orbit, can calculate and curve fit data of at least one track, and use the maximum deviation after the fitting as the pipeline deviation of the satellite, so as to comprehensively analyze the rigorous regression accuracy of the satellite, effectively overcome the limitations of single-point pipeline deviation calculation, and suppress the influence of single-point error data on pipeline deviation calculation by using the curve fitting method, so that the calculation result is more accurate and reliable.

The application relates to a flywheel speed control method and device based on a polynomial trajectory planning, belongs to the flywheel speed control method and device based on the polynomial trajectory planning, and solves the problem of how to reduce the impact disturbance introduced when a flywheel tracks an expected speed instruction of attitude control and improve the attitude stability of a satellite. The method comprises the following steps: receiving an expected speed instruction sent by an attitude control system; for each attitude control period of the attitude control system, a speed response curve between two adjacent expected speed instructions is constructed by adopting an expected speed sequence based on the polynomial trajectory planning, so that the acceleration is smooth and the jerk is continuous. The flywheel speed control method and device based on the polynomial trajectory planning are suitable for flywheel speed control in satellite attitude control.

The application provides a satellite damage reconstruction phase adjustment method, device, equipment and medium, and relates to the technical field of satellites, and the method comprises the following steps: initializing a preset simulation platform according to the basic information of an on-orbit constellation, and mapping the running state data of the on-orbit constellation to the preset simulation platform; when it is detected that the on-orbit satellite satisfies a preset damage condition, the corresponding on-orbit satellite is marked as a damaged satellite, and the corresponding satellite model is set as a damaged satellite model; the simulation phase set of each satellite model after the damaged satellite model is removed is obtained, and a target adjustment phase set is determined with the simulation phase of any satellite model in the simulation phase set as a reference; the phases of at least part of the satellite models are adjusted in sequence, and when the constellation model satisfies a preset running condition, the on-orbit satellite is controlled according to the current phase interval. The application realizes efficient local optimization and feasible closed-loop verification of phase reconstruction after satellite damage, and comprehensively ensures the reliability of the on-orbit constellation running.

The ever increasing number of orbiting bodies in low Earth orbit has made it infeasible to calculate potential conjunctions between orbiting bodies more than a few days in advance, even with the aid of supercomputers. Disclosed embodiments utilize machine learning to predict potential conjunctions between orbiting bodies faster than state-of-the art systems by orders of magnitude. This enables potential conjunctions to be identified well in advance (e.g., 30 days or more), so that they may be prioritized (e.g., for fine calculations), visualized, and remediated (e.g., via control of the impacted satellites).

The present invention provides an orbit design system for a global carbon inventory satellite, comprising: a long residence unit in the northern hemisphere configured to enable the global carbon inventory satellite operates in a mid-orbit elliptical orbit, and enable the global carbon inventory satellite to be located above the latitude of the human activity intensive region when it operates to its apogee.

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.

Spacecraft servicing devices and related methods may include a propellant tank configured to store a propellant and to be placed into fluid communication with a portion of the target spacecraft.

A method for controlling a set of gyroscopic actuators of a satellite, distributed into two coplanar subsets, is described, the method comprising: Receiving a setpoint angular momentum, Determining an angular momentum to be provided by each coplanar subset (, ), from a distribution parameter, between the two subsets, of a component of the setpoint angular momentum along an axis () common to the two planes, r being a scalar comprising the lowest admissible value and the highest admissible value of the contribution of one of the coplanar subsets to the component of the setpoint angular momentum along the axis common to the two planes, determining a tilt angle value of each gyroscopic actuator of a subset, from the angular momentum to be provided by said subset, and Issuing a command to each gyroscopic actuator comprising the respective tilt angle value determined.Figure from the summary: Figure 2.

An orbital analysis unit (431) of a collision avoidance assist business device identifies a mega-constellation satellite group formed in an orbital altitude which the mega-constellation satellite group is anticipated to pass during a flight of an unsteady-operation space object. An announcement unit (432) of the collision avoidance assist business device announces a danger alarm and orbital information of the unsteady-operation space object, to a mega-constellation business device which manages the mega-constellation satellite group. A collision analysis unit (411) of a satellite mega-constellation business device analyzes collision of the unsteady-operation space object with an individual satellite constituting the mega-constellation satellite group. A countermeasure formulating unit (412) of the satellite constellation business device formulates a collision avoidance countermeasure when collision is predicted.

This invention relates to a non-contact, split-type satellite relative position and attitude calculation method using a magnetic levitation stabilization mechanism, belonging to the field of aerospace technology; it establishes the dynamic equations for the position and attitude of the payload compartment; it establishes the relative dynamic equations between the payload compartment and the platform compartment; and it uses matrix F... L Represent 8 magnetic levitation forces; calculate the resultant force F formed by the magnetic levitation forces at the geometric center of the magnetic levitation system. A According to the resultant force F A Calculate the magnetic levitation force acting on the center of mass of the payload compartment. D F 磁上 Establish magnetic levitation force f Li The calculation equation is used to obtain the magnetic levitation force f. Li The relationship between the payload compartment and the magnetic levitation gap δ0 is established; the conversion relationship between the relative position of the payload compartment and the platform compartment and the magnetic levitation gap δ0 is established; the relationship between the relative position of the payload compartment and the platform compartment and the magnetic levitation force f is obtained. Li The relationship, that is, based on the magnetic levitation force f Li The relative position and attitude of the payload compartment relative to the platform compartment are calculated. This invention solves the problem of calculating the relative position and attitude of the payload compartment following the platform compartment based on the magnetic levitation stabilization mechanism, and provides a theoretical model for non-contact split-type satellite following control and vibration suppression system.

The application provides a kind of based on wind shear environment self-adapting aircraft dynamic soaring control method, system and medium, the method includes: constructing the aircraft dynamics simulation environment containing wind shear model, obtains the real-time state information of aircraft;Real-time state information of aircraft is input into deep neural network policy model, and the control action information of aircraft is output;According to the state transition information of aircraft at the current time step, calculate the comprehensive reward value;Using reinforcement learning algorithm, based on real-time state information, control action information and comprehensive reward value, update the parameters of deep neural network policy model;Real-time monitoring of the task success rate of aircraft in the preset time window, dynamically adjust the wind field parameters and task target parameters of simulation environment according to the task success rate;The application effectively solves the problem of training convergence difficulty and strategy collapse caused by sparse reward in dynamic soaring task, realizes the efficient and robust control of aircraft in complex wind field environment.