64 results about "Orbital plane" patented technology

A system includes a plurality of satellites orbiting around a celestial body in a plurality of orbital planes. Each satellite includes an imaging device having a field of view (FOV) to capture an image of a sky that includes celestial image features of resident space object (RSO), stars, and / or planets. The satellite processor is configured to input the image data for each image into a pre-processing software pipeline to generate for each image a contrast-enhanced image data replica of each image, and an enhanced image data replica of each image, to receive from an output of a known-unknown RSO split data processing pipeline, a determination that a candidate RSO is a known RSO stored in an RSO catalog, or an unknown RSO, and to assign to the candidate RSO based on the determination, an RSO ID of the known RSO listed in the RSO catalog, or a new RSO ID.

The invention discloses an internet constellation dynamic routing optimization method and system, and the method comprises the steps: setting orbit plane parameters and satellite parameters, and carrying out the modeling of an orbit and a constellation; performing link visualization and time-varying graph generation according to ephemeris data and the constructed constellation model; defining a related space, a network and a function based on a deep reinforcement learning algorithm; and based on an MADDPG framework, introducing a multi-agent collaborative decision-making mechanism, and outputting action selection and execution results after collaborative optimization. The system comprises a model construction module, an embedding module, a reinforcement learning design module and a collaborative optimization module. By using the method, the dynamic change of the low-orbit satellite network can be effectively dealt with, satellite resources are fully utilized, and efficient and reliable communication is realized. The method can be widely applied to the field of wireless communication.

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

The invention discloses an LEO track single event effect inversion method and device, a medium and a product, and relates to the field of space environment monitoring and evaluation, and the method comprises the steps: carrying out the training and optimization of a first LSTM neural network and a second LSTM neural network at the same time through a training data set, and obtaining a classification model and a regression model; the current position and the current time are pushed forwards for a set number of hours, the average value of the space weather parameters of each hour is input into a classification model and a regression model at the same time, and a numerical value output by the classification model and a particle flux predicted value at the current position output by the regression model are multiplied; obtaining a final particle flux prediction value at the current position; obtaining the particle flux of the LEO orbital plane according to the final predicted value of the particle flux at each position to be predicted on the LEO orbital plane; and obtaining the single event upset rate of the LEO orbital plane according to the particle flux of the LEO orbital plane. The method can accurately predict the particle flux of the LEO orbital plane and the single event upset rate of the LEO orbital plane.

The invention discloses a radar imaging detection satellite constellation design method, which comprises the following steps of: decomposing a constellation into a plurality of orbit surfaces, uniformly describing each orbit surface through four parameters, constructing an asymmetric constellation structure, and establishing a multi-target optimization model for solving by taking the total cost of the constellation, the maximum revisit time and the detection performance as targets. Non-dominated quantizer sorting is introduced in the sorting process, and satellite constellation design is completed. According to the method, single-satellite-based optimization of the salted fish is converted into orbital plane-based optimization, the optimization variable dimension is remarkably reduced, the method is particularly suitable for large-scale constellation design comprising hundreds of satellites, meanwhile, the Pareto frontier is obtained through non-dominated sorting, various constellation configuration schemes emphasizing different performances are provided, and the method is suitable for large-scale constellation design. Compared with an existing constellation design method, the problems that the to-be-optimized parameters are increased along with the number of satellites, and efficiency is low are solved, the scheme output by the algorithm is more reasonable through the improved sorting mode, and the method can be used for constellation design.

A method for node sensitivity analysis affecting the degradation of the communication performance of a giant constellation. The present invention relates to the field of spacecraft system simulation. It solves the problem that in the prior art, there is a lack of a method for measuring the sensitivity of the constellation to changes in the number of satellites, which brings changes in performance such as communication delay. The method includes: reading the parameter values under different two-line element data segments according to the format of the two-line element data record and saving them to the SGP4 orbit model structure variable; dividing different sub-constellations within the giant constellation according to the orbital inclination or the semi-major axis of the orbit; setting that the satellites within each orbital plane form a routing domain, setting the gateway on / off condition, setting that the satellite nodes can communicate within the domain, and restricting the connectivity of the nodes in different domains; updating the connectivity between the nodes, outputting the on / off situation and delay between the nodes; and performing routing and outputting the routing result to complete the node sensitivity analysis. It is also applicable to the field of measuring the communication delay of giant constellations.

The invention discloses a satellite deployment method, system and device and a storage medium, and relates to the technical field of satellite deployment. The method comprises the following steps: for a backup satellite adjusted on the same orbital plane, obtaining a relative phase angle between the backup satellite and a failure satellite after the backup satellite and the failure satellite run after the transfer time based on the transfer time required by the backup satellite to climb to a target orbit from a drift orbit, and obtaining the drift time based on the relative phase angle; for the backup satellite adjusted across the orbital plane, the running time of the backup satellite is obtained, the right ascension of the relative ascending node between the backup satellite and the failed satellite after the backup satellite and the failed satellite run after the running time is obtained based on the running time, and the adjusting time is obtained based on the right ascension of the relative ascending node; and setting the ignition time of the backup satellite based on the drift time and the adjustment time. The method has the beneficial effects of low energy consumption and short consumed time.

The invention discloses a satellite network laser inter-satellite link establishment method and system, electronic equipment and a storage medium, and mainly solves the problem of low availability of a satellite network established in the prior art due to sun-sun influence. According to the scheme, the method comprises the following steps: inputting constellation configuration parameters of a satellite network, and calculating satellites which are uniformly distributed on each orbit surface; according to the input constellation configuration parameters, establishing a same-orbit inter-satellite laser link between all satellites in the constellation and adjacent satellites in the same orbit plane where the satellites are located; according to the constellation configuration parameters, calculating boundary phase factors used for judging different-orbit link connection modes; and according to the phase factor of the constellation and the boundary phase factor, establishing a different-orbit laser inter-satellite link between satellites on all adjacent orbits, and completing the construction of the whole-network inter-satellite laser link. According to the method, the optimal link connection configuration can be screened out in the satellite network construction stage, the availability of the satellite network is improved, and the method can be used for inter-satellite link sun-sun avoidance.

The invention provides a laser link pointing calculation method and system in GEO satellite attitude adjustment. The method comprises the following steps: S1, establishing a laser link two-dimensional turntable coordinate system and a satellite body coordinate system; s2, calculating in a satellite body coordinate system to obtain a first laser link pointing vector and a conversion matrix; s3, calculating in the reference coordinate system of the laser two-dimensional turntable to obtain a second laser link pointing vector; and S4, calculating according to the first laser link pointing vector, the second laser link pointing vector and the conversion matrix to obtain an azimuth angle and a pitch angle of the laser link two-dimensional turntable. The method can adapt to the stable tracking and pointing of the laser link to the link building target satellite under the attitude yaw U-turn scene of the link building satellite, and during the attitude yaw U-turn maneuvering period, the azimuth axis rapid driving is carried out when the laser link two-dimensional turntable is selected to be vertical to the orbital plane, so that the link building interruption time is shortest, and the optimal efficiency of link building between the orbital satellites is ensured.

The invention discloses a visibility prediction method for a low-orbit communication satellite, a terminal and a storage medium. The method comprises the following steps: initializing terminal parameters and target constellation parameters; acquiring ephemeris data of a corresponding satellite based on the target constellation parameter; calculating precession parameters of a satellite orbit in combination with the ephemeris data and the terminal parameters; distributing reference satellites for each orbit of the constellation according to the precession parameters; and predicting the visibility of all satellites of the constellation to the terminal based on the reference satellite and the terminal parameters. According to the method, the reference satellite represents the motion characteristics of the whole orbital plane, so that the calculation complexity of visibility prediction is remarkably reduced, and efficient and rapid prediction of the visibility of the low-orbit communication satellite is realized.

The invention provides a method for autonomously reducing power consumption in a shadow period of a low-orbit satellite, and the method comprises the steps: setting a first threshold configuration and a second threshold configuration which comprise different threshold ranges for each heater, and enabling the upper and lower limits of the second threshold configuration to be not lower than those of the first threshold configuration; establishing a mapping relation between a satellite orbital plane solar incident angle and an orbital period illumination duration, and setting a critical solar incident angle as a shadow period control starting condition; the method comprises the following steps: monitoring accumulation time of non-discharge of a storage battery pack in real time, triggering threshold switching condition judgment when a solar incident angle is within a certain range in combination with current illumination period duration and preset heater advanced temperature rise time, and switching from first threshold configuration to second threshold configuration for preheating if a specified condition is met. And after the storage battery is discharged, the first configuration is automatically recovered. Therefore, according to the invention, the discharge depth and the peak discharge current of the storage battery can be autonomously reduced, and the on-orbit energy safety management capability is improved.

The application discloses an initialization method of a conic surface under a projection linearization track surface, and belongs to the technical field of calculation, reckoning or counting. The method first obtains a preliminary estimation of a target track through three groups of observation data; determines a conic surface containing an initial track surface according to initial track information and geometric knowledge; estimates the missing height information in combination with the conic surface, and obtains more position information of the target in combination with the existing angle information; projects the position information of the target into a two-dimensional section, re-estimates the track parameters by using a least square algorithm, and more accurate track parameters can be obtained; then the iteration process is repeated by using the newly obtained track information until a certain initial orbit determination accuracy requirement is met. The algorithm firstly fits the height information of the target by using a conic surface coordinate system, and realizes linear fitting of the position vector by using a track surface projection method, so that the calculation amount is small and the accuracy is high.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapidly transferring an elliptic orbit different-plane target to the same orbit, and particularly relates to a fuel-saving control on-satellite autonomous solving method for rapidly transferring the elliptic orbit different-plane target to the same orbit from a different orbit plane. Comprising the following steps: judging orbital plane normal correction or orbital shape correction according to orbital elements of an initial satellite and a target; solving a fuel-saving orbital plane normal correction control strategy according to orbital plane normal orbital element difference of the satellite and the target; solving a fuel-saving orbit shape correction control strategy according to the in-orbit element difference of the orbit planes of the satellite and the target; and a high-precision orbit extrapolation model is adopted to carry out iterative correction on the fuel-saving control strategy. The method is suitable for orbital transfer of any elliptical orbit target by any near-earth space elliptical orbit satellite, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The application discloses an antenna switching method for inter-satellite networking. When the included angle between the direction vector of any antenna in a closed state and the direction vector of the opposite satellite is less than or equal to a preset value within a continuous specified time length, the antenna is opened, and the currently opened antenna is closed.

The invention provides a navigation constellation autonomous orbit determination method based on space-based observation, which comprises the following steps: establishing inter-satellite routing, and permanently establishing a link based on phases between satellites in the same orbit plane; a laser inter-satellite link observation value is obtained, and a vector between every two satellites in the same orbital plane is obtained based on laser inter-satellite link distance measurement and angle measurement observation, that is, the laser inter-satellite link observation value represents vectors from the satellite i to the satellite j; coordinate system conversion: converting the laser inter-satellite link observation value from a satellite body coordinate system to an inertial system; establishing an observation equation; establishing a motion equation; and generating a normal equation, combining the observation equation and the motion equation to generate the normal equation, and performing inversion to obtain an orbit determination result. According to the method, the feasibility of realizing autonomous orbit determination of the navigation constellation through laser inter-satellite link and star-sensitive observation is explained. The direction precision of the laser inter-satellite link is equivalent to or even higher than that of a fixed star sensor, and a high-precision and long-term stable orientation reference can be provided for a normal equation, so that the autonomous orbit determination precision and reliability are improved.

The invention provides a satellite switching method, a satellite switching device and electronic equipment, which can be applied to the technical field of satellite communication. Based on the method, when the user equipment needs to be switched from a source satellite to a target satellite, a first angle difference between the user equipment and the source satellite based on a first orbital plane and a second angle difference between the user equipment and the target satellite based on a second orbital plane can be firstly obtained; wherein the first orbital plane is an orbital plane where a source satellite is located, and the second orbital plane is an orbital plane where a target satellite is located; determining target switching time according to the first angle difference and the second angle difference; correspondingly, the user equipment switches from the source satellite to the target satellite based on the target switching time. By calculating and utilizing the angle difference based on the orbital plane, the related data processing process can be effectively simplified, and the corresponding switching time can be quickly and accurately determined, so that the satellite switching of the user equipment can be accurately and timely realized.

The application provides a stable hierarchical routing method, device and electronic equipment of a star-ground fusion Internet, and relates to the technical field of communication. The method comprises the following steps: dividing the star-ground fusion Internet into three routing domains, the first routing domain comprising a plurality of ground nodes, the second routing domain comprising a plurality of orbital planes, and the third routing domain comprising a plurality of orbits in the same orbital plane; determining the first routing domain according to a ground node address, the ground node address being determined based on the coordinates of the geographic block where the ground node is located in the routing coordinate system; in the case where the first routing domain and the second routing domain are interconnected, using a relay ground station or an opportunistic link to route data from a source end to a destination end; and in the case where the first routing domain and the third routing domain are interconnected, routing data from the source end to the destination end based on an inter-satellite topology state and a routing coordinate system. The scheme of the application can construct a stable and efficient routing hierarchy in a dynamically running star-ground fusion Internet.

Various approaches for the deployment and coordination of network operation processing, compute processing, and inter-satellite communication coordination, within one or multiple satellite non-terrestrial networks, are discussed. Among other examples, a data center located at one or more satellites operating in a middle Earth orbit (MEO) plane, geosynchronous orbit (GEO) plane, or high-Earth elliptical orbit (HEO) plane, may be used to provide network and data processing operations for a low-Earth orbit (LEO) constellation.

The present invention relates to the field of automatic control technology, and more specifically to a satellite-onboard coupling control method and system for long-term staring between satellites in different orbits. The method comprises: determining a reference time based on the relative position relationship between a host satellite and a target satellite; obtaining an optimal vector matrix based on the position and velocity relationship between the host satellite and the target satellite at the reference time; obtaining an optimal vector direction based on the relationship between a diagonal in a diagonal plane of the field of view of a space turntable and an optimal orientation vector of the diagonal plane of the field of view of the space turntable; obtaining the attitude of the host satellite after rotation adjustment based on an initial attitude matrix and an optimal vector matrix; and performing a long-term staring of the host satellite turntable at the target satellite based on the optimal vector direction, the attitude of the host satellite after rotation adjustment, an installation relationship matrix, a conversion relationship matrix, and the relative position relationship between the host satellite and the target satellite. The present invention optimizes the staring duration between satellites in different orbits and improves the performance of information data transmission.

As the ion beam is accelerated, the radii of the closed orbits gradually increase, and the centers thereof move in a direction approaching the peripheral edge portion along a predetermined radial direction of the cavity, and upon reversing the direction of movement, move further toward the center of the cavity. The intensity distribution in the orbital plane of the main magnetic field is designed to realize the foregoing feature. Thus, an accelerator is provided that is compact and that enables the energy of an extracted beam to be changed, that enhances the efficiency of beam injection into the accelerator from an external ion source, and that improves a dose rate of the resulting extracted ion beam.

The invention provides a space-based service load migration method and system based on thermodynamic diagram dynamic region division. The system comprises a thermodynamic diagram generation module, a dynamic region division module, a migration decision module and a fault-tolerant execution module. The dynamic region division module generates a service load thermodynamic diagram according to the thermodynamic diagram generation module, divides a network region, identifies a service hotspot region and quantifies the level; the migration decision module plans a service load migration path according to the service hotspot area; and the fault-tolerant execution module searches for replacement nodes to generate a plurality of paths, and executes load migration according to the optimal migration path. According to the method, global calculation is decomposed into parallel calculation of multiple orbit planes, global balance of space-based satellite networking platform loads is achieved, and the defect that under low-orbit rasterized giant constellation networking, resource waste and network congestion of partial regions are caused by uneven service loads is overcome.

The present invention discloses a method for analyzing the optimal pursuit start azimuth based on the condition of pulsed maneuver in orbital pursuit and evasion, including: setting the pursuit and evasion to be carried out on the same orbital plane, establishing an orbital coordinate system with the initial position of the evading spacecraft as the origin based on the orbital altitude and the initial relative distance between the evading spacecraft and the pursuing spacecraft, and determining the initial states of the evading spacecraft and the pursuing spacecraft; based on the initial states of the evading spacecraft and the pursuing spacecraft, using a pulsed pursuit and evasion game strategy solution method based on game tree search to obtain the relative distance under the optimal strategy; the pursuing spacecraft starts to pursue from different start azimuths respectively, compares the relative distances at all azimuths under the optimal strategy, and obtains the optimal pursuit start azimuth under this initial condition. The present invention takes into account the limited action set of the pursuit and evasion spacecraft and the influence of the natural evolution of the spacecraft, obtains the optimal pursuit start azimuth under the same orbital plane, and is helpful for the subsequent research on the selection of the initial position in orbital pursuit and evasion.

The present application discloses a satellite rescue method, apparatus, device, and storage medium. The present application relates to the field of satellite rescue technology. The satellite rescue method includes: obtaining solar position information, satellite positioning data, and imaging parameters of an optical payload carried by a rescue satellite, wherein the satellite positioning data includes orbital parameters of the faulty satellite and the orbital parameters of the rescue satellite; adjusting the orbital parameters of the rescue satellite based on the orbital parameters of the faulty satellite so that the rescue satellite and the faulty satellite share the same orbital plane; controlling the rescue satellite to orbit around the faulty satellite based on the solar position information and / or imaging parameters to obtain fault information about the faulty satellite; adjusting the orbital parameters of the rescue satellite to control the rescue satellite to approach the faulty satellite until the relative positions of the faulty satellite and the rescue satellite reach a preset distance range, and repairing the faulty satellite using the rescue satellite based on the fault information. The present application can solve the technical problems of low efficiency and success rate of satellite rescue.

A system includes a plurality of satellites in orbit around a celestial body in a plurality of orbital planes. Each satellite includes an imaging device having a field of view (FOV) to capture an image of a sky that includes celestial image features of resident space object (RSO), stars, and / or planets. The satellite processor is configured to receive image data from sensors in the imaging device, to increase a positional detection accuracy of the celestial image features by defocusing imaging optics of the imaging device, to capture image data in a volume of the sky as the FOV of the imaging device on each satellite moves in an orbital plane, and generate by a centralized computer at least 1,000 celestial image features using the image data transmitted from each of the satellites.

The invention discloses a double-satellite remote sensing system and method based on inter-satellite laser communication and tracking and pointing integration, and relates to the technical field of space remote sensing. The system is composed of a first satellite and a second satellite which fly back and forth in the same orbital plane, and the first satellite and the second satellite are respectively provided with a multispectral camera, a hyperspectral camera and a transmission system integrating tracking and pointing, inter-satellite laser communication and multi-dimensional calibration functions. The transmission system comprises an inter-satellite laser communication load and a Beidou short message backup load, a stable laser link can be established autonomously, and tracking and pointing precision is guaranteed through three calibration modes of fixed star, satellite-ground and inter-satellite. The system transmits trigger information through an inter-satellite laser link, and realizes collaborative observation of multispectral general survey-hyperspectral detailed survey. The method comprises the steps of calibration preposition, laser link establishment, collaborative observation, calibration maintenance and the like, and supports data collaboration and emergency communication. The timeliness and observation precision of double-satellite remote sensing are improved, and the high-reliability and high-timeliness requirements of scenes such as earth surface anomaly monitoring and the like can be met.

An embodiment of the present application discloses an ephemeris prediction method and device, which relate to the field of satellite navigation technology and are used to reduce the amount of ephemeris parameter broadcast data. The method may include: a server obtains EOP data and historical ephemeris data for a preset time period; performs satellite orbit prediction and satellite clock error prediction based on the EOP data and historical ephemeris data to obtain a predicted orbit and predicted clock error for a future preset time period; after fitting the predicted orbit into orbital parameters, encodes the orbital parameters of satellites belonging to the same orbital plane to obtain an orbital parameter code for each orbital plane; after fitting the predicted clock error into clock error parameters, encodes the clock error parameters to obtain a clock error parameter code; and sends the ephemeris parameter code to a terminal device, where the ephemeris parameter code includes a clock error parameter code and an orbital parameter code for each orbital plane.

To provide a linear motion guide device capable of securing smooth ball movement while mitigating impact force regardless of the behavior of a ball entering a track surface, and a manufacturing method thereof.SOLUTION: In a linear motion guide device, when a plane passing through a widthwise center of a slider body and a longitudinal axial line of a guide rail is defined as a vertical center plane CP, a plane perpendicular to the vertical center plane CP and passing through a center line of a straight part of a rolling passage is defined as a horizontal reference plane HP, and a plane parallel to the vertical center plane CP and intersecting a second inclined surface and a third inclined surface is defined as a vertical reference plane VP, a point P2 on a first inclined plane that is farthest from the vertical reference plane VP is farther away from the vertical reference plane VP than a point P1 on a first raceway plane that is farthest from the vertical reference plane VP, and a point P4 on the second inclined surface and third inclined surface that is farthest from the horizontal reference plane HP is farther away from the horizontal reference plane HP than a point P3 on the first raceway plane that is farthest from the horizontal reference plane HP.SELECTED DRAWING: Figure 5

This invention discloses a method for initial orbit determination of a non-cooperative target under space-based angle-only extremely short arc observation, belonging to the technical field of initial orbit determination. This invention addresses the failure of existing initial orbit determination methods under extremely short arc conditions. The method includes: using the unit normal vector of the target orbital plane as the variable to be optimized; obtaining the relationship between the slant distance and the unit normal vector based on the geometric relationship between the target position vector, the slant distance between the space-based platform and the target position, the target azimuth unit vector, and the space-based platform position vector; transforming the solution of the unit normal vector into the solution of the X-axis and Y-axis components of the unit normal vector based on the normalization characteristics of the unit normal vector, and determining the constraints of the solution; and employing a heuristic algorithm to solve for the X-axis and Y-axis components of the unit normal vector, thereby determining the unit normal vector and achieving initial orbit determination. This invention improves the accuracy and robustness of initial orbit determination.

The invention provides a sun-synchronous orbit inclination angle adjusting method and device, and the method comprises the steps: determining a position relation between an initial orbit and a target orbit, enabling two intersection points of the initial orbit and the target orbit to be located above an equator, and taking one of the intersection points as a satellite orbit transfer point; determining an orbital transfer angle required for adjusting the satellite from the initial orbital plane to the target orbital plane; according to the system state of the satellite and the task target, the single ignition duration of the satellite and the ignition moment and shutdown moment of the satellite are comprehensively determined; determining a satellite ignition direction; the speed increment of satellite monopulse ignition is analyzed and calculated, and the total impulse required for changing the inclination angle of the satellite orbit is determined according to the speed increment; the total satellite ignition orbit number is analyzed and calculated according to the total impulse and the actual impulse generated by each time of ignition; and based on the satellite orbital transfer point, the satellite ignition direction, the satellite ignition time and the shutdown time, the satellite is controlled to carry out secondary ignition orbital transfer, and then inclination angle adjustment is completed.

The invention provides a spacecraft cluster density deployment method and device based on decoupling control, and belongs to the technical field of universe navigation. According to the method, the state of each spacecraft in the spacecraft cluster can be directly and accurately adjusted from the inside and outside of the orbital plane through the density control speed field of the spacecraft cluster, and the density control effect of the spacecraft cluster is improved while the calculation complexity of density control of the spacecraft cluster is reduced.