845results about "Artificial satellites" patented technology

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The invention discloses a super-large plane array type space computing power center, and belongs to the field of space technologies and space infrastructures. The space computing power center is deployed in a morning and evening orbit and comprises a plurality of computing power satellites which are physically docked in orbit through a docking mechanism. Each computing power satellite comprises a satellite platform, and a computing power load, a condensation type energy system and a space pump drive fluid loop system which are arranged on the satellite platform, and a docking mechanism for inter-satellite in-orbit physical docking is also assembled on the satellite platform. A plurality of computing power satellites form a planar array type rigid whole body through the docking mechanism, and energy scheduling among the light condensation type energy systems is realized based on the docking mechanism. According to the space computing power center, through multi-satellite physical docking and energy collaborative scheduling, the energy utilization efficiency and the power supply reliability of the system are improved.

A multi-satellite task scheduling hybrid control method driven by a large language model comprises the steps that a natural language task instruction input by a user is subjected to semantic analysis through an adaptive large language model, and structured scheduling parameters, heuristic scheduling algorithm control parameters and multi-target optimization weights are automatically generated; the method is used for guiding dynamic allocation and resource scheduling of satellite tasks. The method is combined with a heuristic algorithm to execute multi-objective optimization calculation, and a search strategy and a convergence behavior are adjusted according to task requirements and constraint conditions, so that the optimization effect of a scheduling scheme is ensured. Furthermore, the method supports a user to perform semantic-driven re-optimization based on natural language feedback of an initial scheduling result, realizes iterative update of a scheduling scheme, forms a closed-loop control process from natural language input to optimization result to feedback adjustment, and gives consideration to scheduling intelligence, controllability and multi-objective optimization performance. The method solves the problems that a traditional method is insufficient in environment dynamic adaptation and high in initial parameter and weight dependence, efficient multi-satellite task scheduling with controllable semantics and man-machine cooperation is achieved, and the method is suitable for scenes such as satellite group cooperation observation and real-time task planning.

Space debris capture systems include an outer assembly and an inner assembly. The Outer assembly includes an outer membrane and an inner membrane, where, in a deployed state, the outer membrane is separated from the inner membrane by a separation gap. The inner assembly is arranged within the inner membrane of the outer assembly. The inner assembly is formed of a plurality of sheets of material arranged at a plurality of orientations relative to the inner membrane and define a plurality of cells within the outer assembly when in the deployed state. A first group of the cells are defined within surfaces of the plurality of sheets of material and a second group of cells are defined between the surfaces of the inner assembly and an inner surface of the inner membrane.

In a space traffic management system (500), space traffic management devices (100) individually mounted in a plurality of mega-constellation business devices and in a debris removal business device (45) are connected to each other via a communication line (200). The debris removal device (45) performs Active Debris Removal (ADR) against debris formed by a satellite managed by a first mega-constellation business operator. The debris removal device (45) acquires real-time high-accuracy orbital information of a satellite group of a second mega-constellation business operator in a timeframe in which a debris removal satellite, during orbital descent, passes through an orbital altitude region where the satellite group of the second mega-constellation flies, the debris removal satellite passing through the satellite group while ensuring flight safety.

The invention provides an in-orbit energy management method and system for a damaged satellite, network equipment, a computer readable storage medium and a computer program product. The method comprises the following steps: autonomously determining whether to execute the reasoning of the SADA rotation angle based on the current state data of a satellite, and selecting a reasoning model under the current energy limitation condition; simulating a regression relationship between the SADA rotation angle and the predicted current of the solar panel according to the selected reasoning model and the current state data of the satellite; and searching the corresponding SADA target rotation angle capable of enabling the predicted current of the solar panel to be maximum from the regression relationship based on a Bayesian optimization strategy, and converting the SADA target rotation angle into an SADA control instruction. According to the method, a damaged satellite energy management framework is constructed, real-time and accurate prediction of the optimal sun-pointing angle of the current peak value of the solar panel in the current damaged state is realized, energy fluctuation caused by satellite attitude and sun vector change is avoided, and an important basis is provided for autonomous maintenance of energy supply on a satellite.

The invention provides a satellite power supply system topology architecture and a maximum power point tracking control method thereof, and aims to improve the energy utilization efficiency of a solar cell array of a spinning body spacecraft under a complex illumination condition. The architecture adopts a combined topology architecture of a preceding-stage sequence switch maximum power point shunt regulation (S3MPR) circuit and a backward-stage four-switch Buck-Boost (FSBB) converter, and is suitable for a special operation environment of a spinning body spacecraft power supply system. The system adopts a bus full adjustment mode, enables the DC bus voltage to be stabilized near the maximum power point voltage of the photovoltaic array through a double closed-loop control strategy, and carries out real-time shunting adjustment on an S3MPR circuit in combination with a multi-peak maximum power point tracking (MPPT) algorithm, thereby ensuring the efficient operation of the photovoltaic array. According to the invention, high-efficiency energy tracking and stable charging can be realized under a complex illumination condition, and the reliability and the energy utilization rate of a spinning body spaceflight power supply system are remarkably improved.

A tiered spacecraft docking station is adapted to facilitate docking of spacecraft within outer space. A first tier includes a first frame enclosing a first area. A first net-like mesh is coupled to the first frame and fills the first area enclosed by the first frame. A second tier includes a second frame enclosing a second area. A second net-like mesh is coupled to the second frame and fills the second area enclosed by the second frame. A plurality of support beams attach the first frame to the second frame. A lander is used to slow and stop a spacecraft on a celestial body. The lander includes a first and a second webbed structure and a decelerator coupled to the first webbed structure and / or the second webbed structure. The decelerator maintains a tension in the first and / or second webbed structure below a predetermined threshold.

The invention discloses a flat plate type satellite assembly line, which belongs to the technical field of satellite production, and comprises a transfer line, and a first assembly line, a second assembly line and a final assembly line which are arranged in parallel. Through the parallel design of the first assembly line, the second assembly line and the final assembly line, the tail ends of the first assembly line and the second assembly line are connected with the starting end of the final assembly line through the transfer line, so that a U-shaped layout is formed, the technological process of double-cabin parallel split charging and testing and refitting after combined charging is achieved, the station arrangement is consistent with the technological process, and the working efficiency is improved. The process flow path is greatly shortened, and cross operation is avoided.

Described herein is a method for wireless transmission of energy in space, wherein a constellation of transmitting satellites orbiting around a celestial body, such as the planet (P) Earth or the Moon (L), are equipped with a laser apparatus for sending a beam of coherent electromagnetic waves to an orbiting satellite or another receiving target, such as a mobile vehicle or a fixed base on said generic celestial body, in order to transmit energy thereto. The total power transmitted is divided among the orbiting satellites in the constellation, resulting in less contraindications than in the case of a single transmitting system with equivalent total power. The advantage of such division among a constellation of satellites mainly consists of shorter periods of solar eclipse, since sunlight is necessary for the operation of the transmission system. The receiving target or satellite converts the beam of electromagnetic waves into electric energy.

Various embodiments comprise a multi-agent spacecraft system for rendezvous and proximity operations. The spacecraft system comprises a chief spacecraft and a deputy spacecraft. The chief spacecraft maneuvers to the orbit of a Resident Space Object (RSO) and deploys a deputy spacecraft to rendezvous with the RSO. The deputy spacecraft maneuvers to the RSO and senses the RSO during the rendezvous and proximity operations maneuver. The deputy spacecraft generates a Three-Dimensional (3D) surface reconstruction of the RSO based on the sensing. The deputy spacecraft determines a roll rate and a relative trajectory of the RSO based on the sensing. The deputy spacecraft returns to the chief spacecraft.

The invention discloses an energy system of a morning orbit computing power satellite. The energy system comprises a power generation unit, an energy storage unit and a power supply control unit, the power generation unit comprises a main battery array and an auxiliary battery array, the energy storage unit comprises a load and platform storage battery pack, and the power supply control unit comprises a load and platform power supply control unit. The main battery array adopts an expandable concentrating solar battery array, the output of the main battery array is connected with the load power supply control unit, and the main battery array supplies power to the computing power load and charges the load storage battery pack; and the output of the auxiliary battery array is connected with the platform power supply control unit to supply power to platform equipment and charge the platform storage battery pack. And the discharge ends of the double storage battery packs supply power to respective loads through the corresponding power supply control units respectively. And the dual-power control unit dynamically adjusts a charging and discharging strategy and power distribution according to the rail phase. According to the invention, through a physically isolated dual-channel power supply architecture, the problems of periodic interruption of terminator track energy and load interference are solved, and independent and stable energy supply of high-power computing power load and high-reliability platform equipment is realized.

A satellite constellation forming system (100) forms a satellite constellation (20). The satellite constellation (20) is composed of a satellite group (300). In the satellite constellation (20), the satellite group (300) provides a service cooperatively. The satellite constellation (20) has a plurality of orbital planes in which a plurality of satellites (30) fly at the same orbital altitude in each orbital plane (21). A satellite constellation forming unit (110) forms the satellite constellation (20) in which orbital altitudes of the plurality of orbital planes (21) are mutually different.

The invention discloses an optical satellite drift angle accurate calculation and attitude correction method and system, and relates to the technical field of satellite attitude control, and the method comprises the steps: calculating a prime unitary circle curvature radius according to a WGS84 ellipsoid parameter, and converting the geodetic coordinate of a ground target point into an ECEF rectangular coordinate; calculating Julian days based on UTC time, determining an age difference angle and a nutation angle, and converting an ECEF coordinate into a J2000 coordinate; calculating the projection of the earth rotation angular velocity in a J2000 coordinate system, and determining a relative velocity vector; determining an orthogonalized body coordinate system in combination with a unit vector from the target satellite to the ground target point; and projecting the relative velocity vector to a body coordinate system, calculating a drift angle and a double-rotation-sequence Euler angle, and correcting a real-time execution attitude by rotating a matrix around a Z axis. According to the method, high-precision image motion compensation can be realized, the imaging quality of a high-resolution optical satellite in a complex task scene is improved, and a technical foundation is laid for wide application of sub-meter remote sensing images.

The invention discloses a satellite autonomous task planning method and system based on a ground digital mirror image, and belongs to the technical field of spacecraft measurement and control. Through the states of a satellite-ground link synchronous satellite and a ground digital mirror image, a task request is received in the ground digital mirror image, a planning algorithm is operated, a candidate scheme is generated for high-fidelity simulation deduction and verification, a feasible scheme is evaluated and selected and converted into an instruction sequence to be uploaded to a satellite, the satellite executes an instruction and feeds back the state, and a closed loop is formed. According to the method, a complex planning and verification process is placed in a ground digital mirror image, the efficiency and reliability of task planning are remarkably improved by using the strong computing power of the ground, meanwhile, the autonomous operation capability of a satellite is improved, and the problems that traditional satellite task planning is slow in response and poor in reliability, and the autonomous planning capability of a pure satellite is weak can be solved.

The invention discloses a giant remote sensing satellite group task planning method based on atomic task decomposition, and relates to the technical field of task planning of remote sensing satellites, and the method comprises the steps: obtaining an observation task submitted by a user, and determining an observation mode of the observation task; performing atomic task decomposition to obtain a corresponding subtask set; determining a constraint condition and a target function in combination with the subtask set; minimizing the objective function to obtain an optimal task observation scheme; and allocating each sub-task to the corresponding remote sensing satellite based on the optimal task observation scheme so as to execute the observation task. According to the method, a complex multi-satellite collaborative observation task is refined into an operable atomic task, so that the complexity of the problem can be reduced, the allocation process of satellite resources is optimized, efficient integration and execution of a dynamic user target and an autonomous observation target are realized, the adaptability and response speed of a remote sensing satellite when facing a complex observation demand are improved, and the remote sensing system is suitable for being popularized and applied. Therefore, the rapid change of user requirements can be adapted.

Systems, apparatus, articles of manufacture, and methods are disclosed for space-based monitoring of aircraft. An example space-based apparatus to monitor aircraft includes communication interface circuitry to access aircraft operation data; machine-readable instructions; and programmable circuitry to be programmed by the machine-readable instructions to: identify an anomaly based on the aircraft operation data; prepare an updated mission plan for an aircraft based on the anomaly; and communicate the updated mission plan to the aircraft.

A system includes multi-module satellites designed to autonomously or semi-autonomously perform experiments in a space environment for investigators. Each multi-module satellite includes multiple testing modules each associated with a different investigator. Each testing module includes a testing computer that runs experiments. A main flight computer on the multi-module satellite communicates with the testing computers for each module, passing information between the testing computers and a main control system. To perform an experiment in a testing module, the multi-module satellite system collects and stores observation data from sensors of the testing module, controls a controller of the testing module to perform the experiment, transmits the observation data from the testing module to an application associated with the testing module via a main control system, and executes instructions receives from the application via the main control system by controlling the controller of the testing module based on the received instructions.

The invention relates to the technical field of aerospace engineering and remote sensing, in particular to an extremely-low-orbit median-inclination-angle remote sensing satellite and a design method.The satellite runs on a median-inclination-angle orbit, and a spherical crown type windward side and smooth side plate transition structure is adopted to reduce aerodynamic resistance and homogenize heat flow; the detector and the refrigerator are connected with the main heat dissipation surface through a heat conduction structure to form a cold end heat dissipation channel; when a satellite is in a launching state, the solar panel is folded on the Zs side of a satellite body, the satellite is unfolded in the + Ys direction after entering an orbit, an included angle is formed between the normal of the solar panel and the-Zs direction, the SAR antenna is folded on the Ys side of the satellite body, the satellite is unfolded in the-Ys direction after entering the orbit, and the normal of the array plane of the SAR antenna points to the + Zs direction, so that a sunshade and counterweight structure is formed; the positive and negative Xs surfaces are provided with double Hall electric thrusters, the orbit solar angle is calculated through the GNSS and the sun vector, yaw switching is executed, and forward and reverse flight switching is achieved.

The invention relates to the technical field of aerospace equipment, in particular to a fairing closing method and device and a fairing. The method comprises the steps that when the target distance of a steering wheel does not reach a preset distance, the target distance of the steering wheel is adjusted to be the preset distance so that the circle center of a first rectification half cover and / or a second rectification half cover can coincide with the circle center of the bottom of a satellite; when the gear shaper on the rectification half cover is not aligned with the gear shaper hole in the satellite, the rectification half cover is controlled to rotate in situ, so that the gear shaper on the rectification half cover is aligned with the gear shaper hole in the satellite; when the first rectification half cover and the second rectification half cover are not aligned, the height of a steering wheel in the first rectification half cover and / or the second rectification half cover is adjusted, so that the first rectification half cover and the second rectification half cover are aligned; and when the first rectification half cover and the second rectification half cover meet preset conditions, the first rectification half cover and the second rectification half cover are closed. According to the method, the operation complexity is reduced, and the fairing closing convenience is improved.

An MLI panel and method of attaching the same to a space body wherein the MLI panels are attached by way of a plurality of electrically conductive interference fit or snap-fit locking elements.

According to the present invention, the search operation of a target area using multiple mobile bodies can be performed more appropriately or more efficiently. [Solution] The present invention relates to a control system for searching for an object by controlling the operation of a plurality of mobile bodies equipped with measuring sensors capable of detecting an object to be searched, the control system comprising: a search target value acquisition unit that acquires search target information relating to the target of the search; a search related information acquisition unit that acquires search related information including at least one of search condition information relating to the conditions of the search and search state information relating to actual values ​​or predicted values ​​of the search; and a search method determination unit that determines or changes at least one of the number of mobile bodies, arrangement distribution, formation, search movement speed, search turning speed, search movement path, search movement direction, and search straight-line movement time of the plurality of mobile bodies used for the search, based on the search target information and the search related information.

Disclosed is a satellite which includes an Attitude Determination and Control System (ADCS) and at least one radiative surface, where the orientation of the satellite with respect to the sun is change

A propulsion system 1 with three propulsion points of a space module 10 for orbital deployment for at least one satellite, said propulsion system 1 comprising a chassis 2 having exactly three first housings 6 each formed to receive a propulsion unit 3 and at least one second housing 7 formed to receive a tank 4, at least one liquid fuel tank 4 arranged in the second housing 7, and exactly three propulsion units 3, each propulsion unit 3 arranged in one of the three first housings 6 and having at least one thruster 30.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

A system and method for decelerating space objects are provided. The system comprises at least one enclosure that is inflated by a medium, such as a gas, liquid, solid particles, or any mixture thereof. The enclosure can be transported to a predetermined orbital location in the path of the space object. Upon penetrating the enclosure membrane, the object passes through the medium and is thereby decelerated leading to decay of its orbit and / or disintegration thereof.

In a space traffic management system (500), space traffic management devices (100) individually mounted in a plurality of mega-constellation business devices and in a debris removal business device (45) are connected to each other via a communication line (200). The debris removal device (45) performs Active Debris Removal (ADR) against debris formed by a satellite managed by a first mega-constellation business operator. The debris removal device (45) acquires real-time high-accuracy orbital information of a satellite group of a second mega-constellation business operator in a timeframe in which a debris removal satellite, during orbital descent, passes through an orbital altitude region where the satellite group of the second mega-constellation flies, the debris removal satellite passing through the satellite group while ensuring flight safety.

Methods, systems, and devices for satellite operations are described. Positioning data for the non-geostationary satellites may be received from the non-geostationary satellites via one or more satellites in an orbit higher than the orbit of the non-geostationary satellite. Based on the positioning data, trajectories of the non-geostationary satellites may be calculated. Based on the trajectories calculated for the set of non-geostationary satellites, a distance between the set of non-geostationary satellites may be predicted to fall within a threshold distance, and an alert may be communicated to one or more operators of the set of non-geostationary satellites that a distance between the set of non-geostationary satellites is predicted to fall within the threshold distance.

To solve a problem that it costs a great deal of money to move a novel technology to space and make a chance for demonstration and experiment for establishing peripheral technologies required for employment, and although there is a method of obtaining funds by making the demonstration and experiment itself be entertainment that people view as a method of coping with the cost thereof, photographing a small satellite from the ground is difficult.SOLUTION: On the basis of a position of a satellite as a subject calculated by a GPS receiver mounted in the satellite, a position calculated by a GPS receiver mounted in a satellite for photographing mounted with a camera, and a posture direction of the satellite for photographing obtained by a star tracker mounted in the satellite for photographing, a video of a designated satellite is made available at any time by photographing the satellite as the subject while controlling an orientation of the mounted camera or an orientation of the satellite for photographing, and transmitting the video to the ground.SELECTED DRAWING: Figure 1