74 results about "Orbital control" patented technology

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.

This invention relates to the field of spacecraft orbit control technology, and particularly to a spacecraft approach control method based on the fusion of imitation learning and reinforcement learning. The method includes: S1: generating initial and target states, solving for the optimal fuel transfer trajectory offline using a direct method, and constructing an expert dataset; S2: based on the expert dataset, training a neural network controller through imitation learning to obtain a pre-trained model; S3: loading the weights of the pre-trained model, constructing a composite reward function, and training the policy network using a near-end policy optimization algorithm to obtain an optimized policy network model; S4: acquiring relative state data, optimizing the policy network model, and outputting the control quantity for the optimal fuel consumption trajectory transfer in real time. This invention rapidly masters an approximately optimal policy through imitation learning and surpasses the performance of expert data through reinforcement learning, completing control command solutions in milliseconds. It balances terminal control accuracy and fuel consumption optimization, offering advantages such as high solution efficiency, excellent control accuracy, and strong generalization ability.

The application discloses a method and device for orbit transfer control of a lunar spacecraft, and realizes orbit transfer of the spacecraft into a full-parameter target orbit. The method comprises the following steps: a three-impulse control strategy of a far-lunar maneuver, phase control and period control is designed; based on an orbit plane, an arch line and a perigee height, control time and omnidirectional control impulses for orbit transfer from an initial orbit to a first intermediate orbit are determined; based on an orbit phase, a first tangential impulse for orbit transfer from the first intermediate orbit to a second intermediate orbit is determined; and based on an orbit period, a second tangential impulse for orbit transfer from the second intermediate orbit to the target orbit is determined. The first tangential impulse and the second tangential impulse are taken as variables, and a far-lunar maneuver position vector deviation is taken as a fitness value, and a genetic optimization algorithm is used to solve initial values of the impulses. The method solves the problem of orbit control solution caused by nonlinear enhancement of orbit evolution under perturbation influence and complexity of omnidirectional impulse control parameters.

The present application relates to a kind of satellite load observation attitude control and data transmission, orbit control autonomous intelligent planning system and method, the system includes mission guide module, attitude control module, orbit control module and data transmission module;The attitude control module includes several thrusters for satellite attitude control, the orbit control module is used for satellite autonomous orbit control calculation, and carries out satellite orbit control, the data transmission module is used for satellite autonomous data transmission calculation, and satellite data is transmitted to ground;The mission guide module is in system control core, executes load observation task implementation strategy, according to priority reasonably arrange satellite's working time sequence, so that the satellite is intelligently switched between attitude control and data transmission to ground, orbit control, controls satellite autonomous intelligent operation in orbit.Autonomous intelligent planning, autonomous data transmission calculation, autonomous orbit control calculation are beneficial.

The application discloses a kind of distributed multi-curved surface array full airspace beam continuous control methods, comprising: setting satellite TT&C relay trigger event, trigger according to including satellite and TT&C station uplink and downlink state monitoring, orbit control instruction and pre-configured trigger instruction of switching condition;With satellite TT&C relay as the control core of random access node, utilize curved surface array tracking receiver state parameters and the partial satellite orbit information cached, based on TT&C relay trigger event, adaptively trigger curved surface array inter-beam switching process, ensure the continuous tracking of curved surface array to satellite.The present application can realize low-latency and low TT&C resource overhead distributed multi-curved surface array full airspace beam continuous control, support the low-latency TT&C relay demand of satellite in future large-scale constellation, realize on-demand switching without pre-planning, real-time capacity optimization, improve the adaptability to station network resources and the dynamic change of satellite motion state.

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 discloses a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control and belongs to the technical field of space orbit control. In order to solve the problem of modeling and simulation of relative motion between a mobile satellite and a target satellite under the condition that the parameters of the target satellite are unknown, the application calculates initial position and velocity components in the first geodetic coordinate system of the earth center according to prior orbit parameters of the mobile satellite and the target satellite; a dynamic and kinematic model of the mobile satellite and the target satellite is established; a state equation of relative motion between the mobile satellite and the target satellite is constructed by taking relative motion position, velocity and disturbance acceleration as independent variables; an extended state observer is adopted to estimate state variables; a mobile satellite orbit motion control model is established based on a proportional and differential control framework; and disturbance acceleration is compensated by adopting active disturbance rejection design to realize the rendezvous orbit design of the non-cooperative target satellite. The application does not need to accurately obtain orbit motion characteristics of the target satellite, and the method has strong adaptability.

A method, system, product, equipment, and storage medium for autonomous control of the satellite's semi-major axis based on energy constraints are disclosed. Belonging to the field of on-orbit satellite orbit control, this invention solves the problem that existing technologies rely on manual ground planning and command uploading, making it difficult to meet the needs of efficient autonomous orbit control for large-scale constellations. The method involves acquiring the satellite's real-time position and velocity, calculating the corresponding orbit control discrimination parameters, combining a bilateral semi-major axis threshold orbit control strategy to obtain the required orbit control quantities, acquiring the current solar phase angle and satellite orbit parameters, calculating the shadow area initiation phase, and the start time of control. Based on the required orbit control quantities and the energy constraints of a single orbit control operation, the planned orbit control duration parameters are calculated. The start time and planned orbit control duration parameters are used as orbit control execution parameters to achieve autonomous control of the satellite's semi-major axis.

The invention relates to an autonomous handling method suitable for orbit control faults of aircraft timing and fixed-point intersection, and belongs to the technical field of guidance control. An orbit control mode of aircraft autonomous timing and fixed-point intersection comprises element guidance and tail end correction. The method is characterized in that the following steps are executed in each control period in the orbit control process: S1, based on attitude information and a current orbit control execution state, autonomously judging whether the aircraft needs to be stopped and controlled, if so, autonomously executing orbit stopping and controlling processing, and after orbit stopping and controlling, adopting a corresponding attitude stabilization recovery measure according to an orbit stopping and controlling reason, and ending; s2, based on the attitude information orbit control execution state, autonomously judging whether the orbit control meets the condition of recovering the orbit control, if yes, entering step S3, and if not, entering step S3; otherwise, ending; and S3, according to different track control modes and track control stopping durations before control stopping, a track control strategy is determined in a corresponding fault recovery processing mode, and the track control strategy comprises the track control starting moment, the air injection direction and the execution speed increment needed by track control.

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.

The present application relates to a kind of low-orbit giant constellation networking orbit control strategy methods, comprising the following steps: S1: according to last group of orbit control ignition timing, GNSS telemetry and ground measured orbit data, constellation key parameter calibration and forecast are carried out;S2: according to the calibration result calculated in step S1 and nominal constellation configuration parameter, constellation networking stage is independently judged, and orbit control strategy is generated;S3: according to the control strategy generated in step S2, the constellation configuration evolution of a period of time in the future is calculated, and the control strategy is verified.The advantage is that it is suitable for the low-orbit giant constellation networking task configured with electric propulsion, and meets the existing giant constellation orbit control task demand.

The invention relates to the technical field of spacecraft orbit control and game decision, in particular to a reinforcement learning-based information waiting strategy forming method in an orbit pursuit scene, which comprises the following steps of: establishing a pulse round system game environment for describing relative movement of an orbit based on an orbit mechanical equation; based on the relative distance between the chasing party and the target, calculating the real-time observation confidence of the chasing party for observing the authenticity of the target, and updating the observation belief of the chasing party in real time by adopting Beta distribution; designing an expert strategy, and generating a state action database containing states and expert actions in a simulation environment; constructing a strategy network with an observation belief feature embedding layer, and pre-training the strategy network by using a state action database; and defining a reward function which aims at minimizing the capture time and contains distance penalty when the capture is not carried out, refining by maximizing a soft Q value of the sum of the reward and the strategy entropy, and finally outputting an optimal chasing party information waiting strategy with the high-entropy characteristic.

This invention discloses a camera imaging control method and apparatus for midnight sunlight avoidance, belonging to the field of spacecraft attitude and orbit control. The method includes: determining whether the satellite is currently in the midnight sunlight avoidance phase; if the satellite is not in the midnight sunlight avoidance phase, and the conditions for entering the midnight sunlight avoidance phase are met, then the satellite is initialized to enter the midnight sunlight avoidance phase, and the satellite is set to rapidly maneuver to a preset target avoidance trajectory; if the satellite is in the midnight sunlight avoidance phase, and the satellite receives a satisfactory imaging command, then the satellite is maneuvered to the imaging point for imaging; otherwise, the satellite remains on the target avoidance trajectory; if the satellite is in the imaging state during the midnight sunlight avoidance phase and is exposed to sunlight, then the satellite is rapidly maneuvered to the target avoidance trajectory; otherwise, the satellite's imaging state is maintained. This invention can ensure the imaging safety of the camera during midnight sunlight avoidance, improve the utilization efficiency of the camera payload, and enhance the user experience.

This utility model relates to a satellite electric propulsion system, addressing the problems of high cost, poor control precision, and low efficiency and reliability of existing chemical thrusters for micro and nanosatellites, as well as the issues of low efficiency and low reliability of resistance-heated thrusters. The system includes a propellant tank, a power supply and control module, a dual-path pressure regulation module, and a dual-thrust module. The propellant tank uses a full TC4 titanium alloy spherical structure to store liquid ammonia. The power supply and control module controls the high-voltage arc heating of the working fluid via a CAN bus. The dual-path pressure regulation module achieves redundant pressure reduction through solenoid valves, preheaters, and pressure reducing valves. The dual thrusters are arranged vertically at their centers of mass, providing mutual cold backup. This design reduces size and weight through modular integration, improves electrothermal conversion efficiency, orbit climbing mission success rate, and working fluid utilization efficiency, effectively solving the technical bottlenecks of existing propulsion systems and is suitable for high-precision orbit control of micro and nanosatellites.

This invention provides a position control method and apparatus based on attitude control error feedforward compensation, relating to the field of rendezvous and docking technology. The method includes: calculating the orbital control error based on the attitude control error; determining the position error at the end of the long-range guidance segment under the influence of the orbital control error; calculating a corresponding compensation amount based on the position error at the end of the long-range guidance segment; and correcting the nominal position value based on the compensation amount, so as to use the corrected nominal position value for position control. This solution can achieve high-precision position control under large-pulse orbital control.

This invention provides a ground hot-fire test method for space orbital control engines, used to assess whether the product can reliably complete on-orbit flight. The method comprises the following steps: Step 1, developing acceptance hot-fire test procedures and operating conditions; Step 2, conducting product acceptance hot-fire tests; Step 3, selecting products for random hot-fire test inspection; Step 4, developing qualification-level life assessment hot-fire test procedures and operating conditions; Step 5, conducting qualification-level life assessment hot-fire tests on randomly selected products; Step 6, evaluating whether the product passes the assessment. The method provided by this invention aims to identify orbital control engine defects in advance through hot-fire test inspections, assess the product's adaptability to flight mission profiles, and ensure the reliable on-orbit operation of long-life space orbital control engines.

The application discloses a multifunctional electric propulsion system of a space vehicle, and relates to the technical field of space propulsion systems.The multifunctional electric propulsion system comprises a gas cylinder module, a fluid management module and a thruster module; the gas cylinder module is connected downstream to the fluid management module, and the fluid management module is connected downstream to the thruster module; and the thruster module comprises a Hall thruster, an electric heating type thruster and a cold gas thruster.The multifunctional electric propulsion system utilizes low-cost working gas, integrates the attitude control and orbit control systems of a satellite through modular design and heat flow design, reduces the difficulty of design and manufacture of the satellite, reduces the launch cost of the satellite, and simultaneously realizes intelligent control of the satellite through the integrated control system combined with orbit control algorithms.

The invention discloses an autonomous control method for a sun synchronous regression orbit, and belongs to the technical field of spacecraft orbit design and autonomous control, and the method comprises the steps: executing an in-plane and out-plane instruction calculation step based on a current instantaneous orbit element vector when determining that a satellite passes through a descending node based on a real-time position vector and a real-time adjacent moment position vector; obtaining in-plane maneuvering final execution time and an out-plane maneuvering instruction; and executing in-plane maneuver based on the in-plane maneuver final execution time, and executing out-plane maneuver when the satellite passes through the ascending node based on the out-plane maneuver instruction. The invention provides an autonomous control method for a sun-synchronous regression orbit, and solves the technical problem that an existing high-precision autonomous orbit control method for the sun-synchronous strict regression orbit is not suitable for the sun-synchronous strict regression orbit at a non-morning-night non-midnight intersection point. And the long-term stability and robustness of the control are enhanced while the high precision of the autonomous control of the sun-synchronous orbit is ensured.

A method for autonomous satellite orbit control using a separable independent propulsion module is proposed. This method does not rely on a ground-based telemetry and control system. By autonomously assessing the remaining propellant, it improves the efficiency of chemical propellant use in the satellite propulsion module while ensuring the safety of the satellite's chemical propellant. Furthermore, by calculating the satellite's orbit control parameters and formulating an orbit control ignition strategy based on orbit change constraints and all injection parameters, autonomous satellite orbit control can be achieved without relying on a ground-based telemetry and control system.

The invention discloses a carrier rocket satellite-rocket separation section variable target orbit control method and a system thereof, and the carrier rocket satellite-rocket separation section variable target orbit control method comprises the following steps: determining an attitude required by orbit change according to an orbit change demand; adjusting the attitude of the rocket body according to the required attitude; after the rocket body attitude adjustment is completed, stabilizing within specified time, and after stabilization is completed, carrying out orbital transfer until a target orbit is reached; and after the target track is reached, the final repair power system is closed, and target track adjustment is completed. The orbital transfer can be calculated in real time according to the target orbital parameters and the current flight state parameters, the design is simple, implementation is convenient, and practicability is achieved.

A rhombic structure orbital transfer aerial mobile charging robot system comprises a suspension type charging robot, a cross orbit, a control unit, a robot positioning system and a call response system, the suspension type charging robot comprises an advancing steering module, and the cross orbit comprises a cross through groove. The advancing steering module comprises an advancing wheel steering unit and a plurality of advancing wheels, the advancing states of the advancing steering module comprise the longitudinal advancing state and the transverse advancing state, and the arrangement of the advancing wheels meets the condition that the advancing wheels are arranged in the advancing direction of the advancing steering module; and the wheel bottom center horizontal distance between any two of the at least four traveling wheels is greater than or equal to the width of the crossed through groove. The invention relates to the field of new energy vehicle charging, and can flexibly walk on a crossed track to change the track so as to realize the purpose of finding a vehicle by a pile, and the traveling wheels with the rhombic structure layout can greatly improve the stability when passing through a crossed through groove, greatly reduce noise and vibration, and improve the charging experience of a user.

The invention discloses a geosynchronous orbit satellite position protection method with an unconstrained thruster installation direction, and belongs to the field of satellite attitude and orbit control. The method comprises the following steps: calculating an orbit parameter estimation value of a satellite according to a future orbit position obtained by recursion of a current orbit position; generating a corresponding satellite orbit position keeping strategy according to orbit parameters which do not meet a preset threshold range in the orbit parameter estimation values; and determining whether the satellite orbit position keeping strategy meets the sunlight avoidance constraint and the thruster layout constraint, and if not, adjusting the satellite orbit position keeping strategy according to the sun vector position, the sunlight avoidance demand and the thruster layout at the future position keeping moment. The method can greatly improve the position protection operation efficiency and guarantee the satellite load safety.

This application provides a method, apparatus, device, and medium for maintaining the configuration of low-Earth orbit satellite formations, which can solve problems such as frequent control, high fuel consumption, and shortened on-orbit life caused by ignoring the differences in aerodynamic drag between satellites in related technologies. The method includes: acquiring the difference in the semi-major axis decay rate between a reference satellite and a companion satellite in the formation; predicting the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction based on the difference in the semi-major axis decay rate and the initial semi-major axis difference; determining an orbital control strategy based on the relative drift trajectory and a preset configuration maintenance threshold; executing the orbital control strategy to adjust the orbital altitude of the companion satellite when the drift reaches or exceeds the configuration maintenance threshold; resetting the semi-major axis difference after adjusting the orbital altitude of the companion satellite, and returning to the step of acquiring the difference in the semi-major axis decay rate.

This invention belongs to the field of track dynamic testing of braking systems, and particularly relates to a method and test bench for track dynamic testing of braking systems, aiming to improve the accuracy of joint testing of braking systems. The test bench includes: a circular runway (1), a power trolley (2), and a control center (3); a track for placing the power trolley (2) is set on the circular runway (1); the control center (3) controls the movement of the power trolley (2) and the action of the braking system, and receives braking test data fed back by the power trolley (2); a universal mounting platform is set on the power trolley (2), on which the braking system for testing is placed; wheels are installed on the power trolley (2) or the braking system, which are normally raised, and the power trolley (2) contacts the circular runway through the wheels. When braking is tested, the wheels are lowered and raised, and the braking performance of the wheels is tested. This invention can be applied to the field of aviation and vehicle braking testing.

This application provides a method and apparatus for orbit design and control of differential interferometry for constellations. The method includes: designing a high-precision reference orbit at the initial epoch based on design constraints, and describing it as a time-scaled point set in a geostationary coordinate system; determining the number of constellations' extraterrestrial stars and the spacing between their positions using the reference orbit as a reference; correcting the position parameters of the first star based on the actual on-orbit epoch of the extraterrestrial stars, and reconstructing the positions of other extraterrestrial stars; calculating velocity vectors based on the reconstructed positions to collectively constitute the orbital states of the extraterrestrial stars; converting the orbital states of the extraterrestrial stars in the geostationary coordinate system into orbital parameters in an inertial coordinate system; and generating a control sequence based on the deviation between the actual orbit determination results and the target orbital parameters to complete coordinated orbit control. This application solves the technical problem of the lack of a systematic common reference constellation design, on-orbit deployment, and coordinated control method in the prior art. Through the integration of theoretical design and engineering practice, it achieves precise constellation configuration and long-term coordinated control, ensuring high performance of differential interferometry missions.

This invention relates to a ground flight control and operation support system for a very low Earth orbit (UEO) satellite constellation. It incorporates predicted space environment information, satellite attitude change information, and orbit control information into the high-precision satellite orbit prediction process, effectively ensuring higher orbit prediction accuracy for the UEO constellation and maximizing the assistance to ground control stations during satellite transits in capturing and tracking the satellite. The satellite attitude and orbit control system simulation module performs real-time simulations of the impact of attitude control and orbit control on orbital maneuvers, thereby achieving higher orbit control accuracy and providing better assurance for the precise execution of missions by the UEO constellation. Automatic scheduling of Earth remote sensing and data transmission tasks, along with ground system fault identification and handling, and collision warning, provide more intelligent support for the operation of the UEO constellation.

The present application relates to the technical field of spacecraft attitude orbit control, and particularly relates to a joint position keeping control method and device under the condition of mass center deviation. The method comprises the following steps: determining the direction of satellite drift caused by the tangential thrust of coplanarly installed electric thrusters when the mass center of the whole satellite deviates by 1mm, and the semi-major axis variation of the satellite in the position keeping control period, determining the target deviation direction and target deviation amount of the mass center of the whole satellite based on the drift direction of the satellite prime meridian caused by the perturbation acceleration and the semi-major axis variation of the satellite in the position keeping control period under the action of the earth perturbation force; adjusting the mass center position of the satellite based on the target deviation direction and target deviation amount; after the satellite is fixed in orbit, the coplanarly installed electric thrusters are used to autonomously perform north-south position keeping control every operation day, and since the deviation of the mass center of the whole satellite can simultaneously generate tangential thrust, the north-south position keeping control is completed, and the east-west position keeping control task is also completed.

The objective is to propose a thrust control algorithm for maintaining a constant argument of perigee in an elliptical orbit where the perigee is in an extremely low orbit, and to provide an orbit control device and orbit control method for an aircraft that enables efficient orbit maintenance and high-resolution observation. [Solution] We propose a control method that keeps the variations of each orbital parameter—major axis, eccentricity, and argument of perigee—within a threshold. Based on the differences in thrust sensitivity of each orbital parameter, priority is given to correcting the eccentricity and argument of perigee, which have higher sensitivity, and then the major axis is corrected. By appropriately adjusting the thrust application region at an orbital position where the correction efficiency of the perigee and argument of perigee is maximized, both maintaining the perigee and minimizing propellant consumption are achieved.

The invention belongs to the technical field of agricultural irrigation, and provides a device and method for testing characteristics of a track control rotary valve, and the device comprises a power supply unit which consists of a variable frequency water pump, a water supply pipe and a water tank and is used for supplying water to the valve and driving a valve core to move up and down; the plurality of water outlet pipes are connected with different water outlets of the valve; the multi-dimensional monitoring unit comprises a first pressure sensor arranged on the water supply pipe, and a second pressure sensor and a flow sensor corresponding to each water outlet pipe; and the controller is in communication connection with all the sensors and the water pump to realize data acquisition and cooperative control. Through the scheme, the pressure, the flow and the on-off state of the valve under different water pressures can be obtained in real time, automatic testing and judgment of dynamic performance are completed, the testing precision and efficiency are improved, and the method is suitable for performance verification and optimization of the hydraulic drive valve in the fields of agricultural irrigation and the like.

The invention relates to the technical field of spaceflight orbit control, and provides a method and system for autonomously generating a satellite orbit control instruction and orbit control activity supervision table in order to solve the problems that a traditional satellite orbit control process depends on manual operation and is low in efficiency and prone to errors. According to the method, non-orbit-control-period orbit data are obtained, an apogee moment sequence is automatically calculated according to set parameters such as task time, thruster working duration and an orbit transfer mode, intelligent grouping is carried out according to a continuous or intermittent orbit transfer strategy, and the orbital transfer efficiency is improved. A complete orbit control plan including a specific ignition time sequence, an orbit lifting instruction, a hexadecimal source code capable of being directly injected and related physical parameters is generated, finally, a reported ephemeris file and an orbit control activity supervision record table which meet the specification are output in a one-key mode, and automation and standardization of the whole process from task planning to instruction generation and document compiling are achieved. The method is especially suitable for a task scene in which a low-orbit satellite climbs an orbit by using a thruster after satellite-rocket separation.