99 results about "Orbit plane" patented technology

A satellite constellation system includes a plurality of satellites, wherein each satellite is configured to orbit in one of a plurality of orbital planes of the satellite constellation system at an altitude range corresponding to low earth orbit (LEO), wherein each of the plurality of orbital planes includes a corresponding one of a plurality of satellite subsets of a plurality of satellites, wherein a first orbital plane of the plurality of orbital planes corresponds to a first coverage area facilitated by a plurality of navigation signals transmitted by ones of the plurality of satellites in the first orbital plane, wherein the plurality of navigation signals transmitted by the ones of the plurality of satellites in the first orbital plane each have a first primary pseudo-random noise (PRN) code; and wherein a second orbital plane of the plurality of orbital planes corresponds to a second coverage area facilitated by a plurality of navigation signals transmitted by ones of the plurality of satellites in the second orbital plane, wherein the plurality of navigation signals transmitted by the ones of the plurality of satellites in the second orbital plane each have a second primary PRN code.

The invention discloses an autonomous low orbit constellation Internet of Things traffic prediction method and device based on satellite spatial-temporal feature mapping, electronic equipment and a storage medium. The method comprises the steps of traffic prediction model autonomous training, hop-by-hop information transmission, model parameter updating, autonomous traffic prediction and cyclic traffic prediction. The method fully explores the space-time characteristics of the satellite traffic, achieves the nonlinear mapping between the space-time characteristics of the satellite and the traffic deviation of the adjacent satellites in the prediction process in the same orbit plane, provides a mechanism for intelligently and autonomously transmitting the traffic information between the low orbit constellation Internet of Things satellites and updating the prediction model, and improves the prediction accuracy of the low orbit constellation Internet of Things satellites. According to the method, efficient and accurate autonomous low-orbit constellation Internet of Things traffic prediction in a traffic sudden change scene is realized, the method is suitable for low-orbit constellation Internet of Things traffic prediction which is large in scale and high in network topology change speed, the calculation complexity is reduced, sudden change traffic can be autonomously processed, and waste of a large number of calculation resources is avoided.

A satellite constellation forming system forms a satellite constellation (20) including two orbital planes each having a different normal direction and in each of which the same number of satellites fly. A satellite constellation forming unit gradually changes an orbital altitude of at least one orbital plane of the two orbital planes from a state in which satellite passage timings of satellites flying in the two orbital planes do not coincide with each other at an intersection neighborhood point (Pc) between the two orbital planes in each of the two orbital planes, while maintaining the state in which the satellite passage timings do not coincide with each other.

A satellite constellation (100) includes a plurality of artificial satellites (111 to 116) that number a multiple of 6. Each of the plurality of artificial satellites orbits in an inclined circular orbit a plurality of times a day. Normals to a plurality of orbital planes corresponding to the plurality of artificial satellites are shifted by an equal angle from each other in an azimuth direction. The plurality of orbital planes make up one or more orbital plane sets each consisting of six orbital planes. Timings at which the six artificial satellites orbit on the six orbital planes of each orbital plane set are synchronized with each other.

To continuously provide service while avoiding collision with a space object, by control of orbits of satellite groups and control of passage timing.SOLUTION: A megaconstellation business device is a business device for managing 100 or more satellite constellations. The satellite constellation provides service by coordination between satellite groups. The megaconstellation business device includes a satellite constellation formation part 11 for forming a satellite constellation having a plurality of orbital planes in which a plurality of satellites fly at the same nominal orbital altitude on the respective orbital planes. The megaconstellation business device includes a satellite constellation formation part 11 for continuously providing service while avoiding collision, by control of orbits of satellite groups and control of passage timing, after having predicted possibility of collision with a space object falling down in an orbit, a launching rocket, a satellite transiting in the orbit, or debris.SELECTED DRAWING: Figure 28

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 relates to a space-based networking radar multi-coverage fusion target tracking method and system, and the method comprises the steps: carrying out the preprocessing of target echo data obtained through the observation of a same target through a plurality of radar nodes, which are disposed on each track plane of a low-orbit space, of a space-based networking radar, and obtaining radar measurement data, the target measurement data are three-dimensional measurement data under a Cartesian coordinate system with a target as an original point, unscented Kalman filtering is carried out on the three-dimensional measurement data to obtain a plurality of corresponding filtering data, sequential fusion is carried out on all the filtering data according to a covariance cross fusion algorithm to obtain fusion state estimation of the target, and the fusion state estimation of the target is obtained. And target tracking is realized. By adopting the method, the precise tracking of the space target can be realized, the low-orbit space target orbit coverage rate can be improved, the cooperative networking operation of the space-based surveillance radar can be realized, and the precision and efficiency of space target monitoring can be effectively improved.

The invention provides a decentralized federated learning method and device for a low earth orbit satellite and a storage medium, and relates to the technical field of satellite communication, and the method comprises the steps: employing a dual-stage adaptive model aggregation mechanism, employing an in-orbit aggregation algorithm in the first stage; in the second stage, cross-orbit plane propagation is adopted, and the propagation turns are dynamically adjusted through a multi-turn Gossip protocol diffusion model to balance the convergence speed and the communication cost; a model self-compensation mechanism is adopted, and for the problem of cross-orbit inter-satellite link packet loss, a receiver compensates lost data packets by using local parameters, so that retransmission requirements are reduced, and robustness and communication efficiency are improved. The low earth orbit satellite-oriented decentration federated learning method provided by the invention is superior to the existing decentration federated learning method in the aspects of convergence performance, communication efficiency and link robustness.

The present invention discloses an aerodynamic layout of a large-lift-surface ultra-low-orbit aircraft, which includes an aircraft body, a group of aerodynamic sails, and a group of solar cell wings. The body of the aircraft provides a space for equipment installation, and its cross-section is an equilateral triangle structure with a side length of W and a length of L. The large-lift-surface aerodynamic sails are located directly below the aircraft body, providing the main mechanical load and the area for solar cell placement. The solar cell wings are attached to the upper surface of the aerodynamic sails. The designed large-lift-surface aerodynamic layout can effectively increase the lift-to-drag ratio of the aircraft under the design requirements of a fixed antenna angle, and change the orbital plane of the aircraft through an aerodynamic assist method, so as to achieve a large-range orbital inclination maneuver for the aircraft in the ultra-low orbit with the minimum fuel consumption.

Disclosed is a method for ejecting at least two satellites (2; 2a, 2b) from a launcher (3), said at least two satellites (2; 2a, 2b) being oriented along a central axis coinciding with a longitudinal axis (A) of the launcher and secured together side by side, the method comprising: a. an initial orientation of the direction of the longitudinal axis (A) of the launcher, so as to form a predetermined angle (θ) with an orbital speed axis (X), the predetermined angle (θ) being between 20° and 60°, b. a first separation comprising the simultaneous separation from the launcher (3) of said satellites (2; 2a, 2b) along the longitudinal axis (A) with a first predetermined differential speed (dV1) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining secured together, then c. a second separation comprising the separation of said satellites (2; 2a, 2b) from each other along an axis perpendicular to the central axis and included in the orbital plane, with a second predetermined differential speed (dV2).

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 invention relates to the field of satellite orbit calculation, in particular to a low-orbit satellite ephemeris analysis method based on a satellite orbit plane coordinate system, which comprises the following steps: acquiring TLE two-line orbit data of a target satellite, geodetic coordinates of a terminal, UTC time and earth rotation average angular velocity; calculating the state of the satellite in the orbital plane; calculating a transformation matrix from a geocentric inertial coordinate system ECI to a satellite orbit plane coordinate system; the LBH coordinates of the terminal are converted into ECEF rectangular coordinates; a rotation matrix from the ECEF to the ECI is constructed, and the ECI system speed of the terminal is calculated; according to the transformation matrix, transforming the ECI coordinate of the terminal into a final satellite orbit plane coordinate system; performing simplified calculation based on a unified calculation framework of a satellite orbit plane coordinate system; and outputting the distance, the relative speed, the Doppler frequency offset and the propagation time delay result which are calculated under the satellite orbit plane coordinate system. And the calculation complexity and the realization difficulty of high-precision orbit forecasting are obviously reduced.

The invention discloses an elliptical reference orbit plane circle satellite formation configuration design method. The method comprises the steps of S1, selecting the orbit number of plane circle formation reference satellites; s2, selecting an elliptical orbit reference satellite RIC coordinate system; s3, establishing a relative kinetic equation of the elliptical reference orbit satellite formation flight; s4, establishing a periodic solution of the linear relative kinetic equation of the elliptic reference orbit; s5, giving out constraint conditions of the elliptical orbit plane circle formation configuration; s6, designing plane circle formation parameters; s7, calculating a relative position speed parameter of the accompanying satellite at the initial moment in the RIC coordinate system of the reference satellite; s8, calculating position and speed parameters of the accompanying satellite in the geocentric inertial coordinate system at the initial moment; and S9, calculating the orbital elements accompanying the initial moment of the satellite, and completing the design of the plane circle formation satellite configuration.

The application aims to provide a remote sensing satellite descending node local time control method based on orbit height adjustment, which has simple control and high control precision. The method comprises the following steps: S1, acquiring orbit elements of a remote sensing satellite, wherein the orbit elements are instantaneous elements or two-line elements; S2, calculating a current descending node local time of the satellite based on the orbit elements; S3, calculating a drift deviation of the current descending node local time from a nominal descending node local time △T ; S4, calculating a current orbit plane rotation rate based on a J2 perturbation parameter of the earth; S5, calculating a difference between the current orbit plane rotation rate and an earth revolution rate; S6, designing an expected orbit plane rotation rate based on the difference calculated in the step S5; S7, calculating an expected orbit height according to the expected orbit plane rotation rate; and S8, adjusting the orbit height of the satellite to the expected orbit height. The application is applied to the technical field of remote sensing satellite measurement and control.

Positioning computer for satellite having an inter-satellite optical communication system on an anti-Earth face. Positioning computer for a satellite (10) intended to move in an orbital plane (P), the satellite being intended to comprise: - a casing (32) having an Earth face (34) defining a Z axis of the satellite intended to point towards the Earth (12), - at least one optical system for communicating with another satellite located in the orbital plane, the first optical system being intended to be located on an anti-Earth face (46), and to have an azimuth field of view around the Z axis, the field of view being limited by an occlusion angle due to the casing.The calculator is adapted to obtain a parameter representing a solar angle (β) defined by the Sun (16) with the orbital plane, and to perform the calculation of a parameter representing the yaw angle (ψ) of the satellite using at least the parameter representing the solar angle and a parameter representing the shutter angle. Figure for the abbreviation: Figure 1.

The application relates to a rotation error correction method of a distributed autonomous navigation constellation, and comprises the following steps: S1, each satellite calculates the orbit plane orientation error of the satellite in orbit and distributes the result to all visible satellites in the constellation in real time; S2, each satellite fits a local constellation rotation quantity according to the orbit plane orientation error of the satellite and the collected orbit plane orientation error; S3, each satellite determines a virtual satellite orbit observation quantity according to the local constellation rotation quantity of the satellite; S4, each satellite constructs a conditional constraint equation of virtual observation according to the virtual satellite orbit observation quantity and the second type of non-singularity orbit root number; and S5, each satellite solves the satellite orbit through filtering according to the conditional constraint equation of virtual observation, so that the overall rotation effect of the constellation is inhibited and corrected. The application can inhibit the rotation effect of the constellation, improve the distributed autonomous orbit determination precision of the satellite, does not increase the inter-satellite data interaction communication pressure, and can stably maintain the constellation configuration for a long time.

As the ion beam is accelerated, the radius of the spiral orbit gradually increases, and the center of the spiral orbit moves in the predetermined radial direction of the gap toward the direction close to the peripheral portion, and then the moving direction of the spiral orbit is reversed and further moves toward the center of the gap. To achieve this, the intensity distribution in the orbit plane of the main magnetic field is designed. Thus, as a small accelerator capable of changing the energy of the extracted beam, the efficiency of beam injection from the external ion source into the accelerator is improved, and as a result, the dose rate of the extracted ion beam is improved.

1. When a rotating body is rotated and fixed to its rotational axis, a new N*BALY lifting force is created. Fn*Baly is a parabolic line characteristic: Fn*Baly = av 2 . Quite different from the normal lifting force with Cotag line: Fn = mv 2 .Cot(ê) / r. See Fig. A, C. 2. N*LEVY Parallel Space Dominion Satellite System: Satellites orbit the Earth in an orbital plane parallel to the equator. Unlike today's satellites, they all orbit the Earth's center. The new system has fewer satellites, evenly distributed, no overlap, no collisions, and stable system connectivity. Fig. D, E, F. 3. New Nguyen Levy formula for the precise calculation of spherical surface: Sn*Levy = n 2 .R 2 & spherical volume: Vn*Levy = 1 / 3.π 2 .R 3, see Fig. J, K, eliminating the 21.5% error of the old formula. It ensures precision in design and manufacturing, increases product quality, reduces costs, and, above all, builds self-confidence among young people and the technical elite. A new, intuitive method accurately measures everything: curves, areas, volumes... that current measuring tapes cannot.

The application discloses a heterogeneous satellite constellation configuration correction method based on stable earth coverage performance. The method comprises the following steps: constructing a first query table and a second query table; for each orbit plane, using a working orbit parameter optimization method, combining the first query table and the second query table, analyzing and calculating dynamic visibility conditions and a visibility list, and then correcting the working semi-major axis and the latitude argument AoL of each satellite on the orbit plane; for each orbit plane, according to the working semi-major axis correction value of the satellite on the orbit plane, combining the first query table, calculating the right ascension of the ascending node RAAN correction value of the satellite; and the working semi-major axis correction value, the latitude argument AoL correction value and the right ascension of the ascending node RAAN correction value of each satellite on all orbit planes constitute the configuration parameters of the heterogeneous satellite constellation. The application restrains the drift of the earth coverage index of the heterogeneous satellite constellation, reduces the constellation control frequency, and avoids the additional increase of the number of satellites due to the drift of the coverage index and the multiple artificial iterative design.

The invention relates to the technical field of satellite positioning, in particular to a Walker constellation-oriented hierarchical interval satellite selection method, which comprises the following steps of: firstly, establishing a Walker constellation of a low-orbit satellite, setting an orbit plane number, a satellite number and a stage number, carrying out satellite screening on the Walker constellation by utilizing the stage number, outputting all selected satellites, and finally, forming a constellation by the screened satellites. The constellation reconstructed by the satellite selection method still keeps the characteristics of a Walker constellation, reduces the redundancy of a constellation system and the repeated coverage between satellites, avoids huge calculation amount caused by excessive satellites, and effectively saves the constellation cost.

an LEO satellite includes a light projecting element that emits emission light to another LEO satellite, an optical telescope, an optical phased array, a light receiving element that receives incident light from the other LEO satellite, a distance measurement unit that measures a distance to the other LEO satellite based on at least one of the emission light and the incident light, and a control unit. The control unit captures the other LEO satellites by scanning emission light using the optical telescope and receives incident light from the other LEO satellites for the other LEO satellites on the same orbital plane, and captures the other LEO satellites by scanning emission light using the optical phased array and receives incident light from the other LEO satellites for the other LEO satellites on different orbital planes.

The invention belongs to the technical field of inter-satellite laser link establishment pointing, and provides an inter-satellite laser link establishment pointing method and system suitable for a low-orbit mobile communication constellation in order to solve the problem that an existing azimuth angle and pitch angle pointing technology based on an east-north-sky coordinate system has defects. According to an inter-satellite laser communication ephemeris mutual transmission mechanism, a mathematical model of inter-satellite laser link establishment angle pointing requirements is established, and azimuth angle and pitch angle calculation formulas of laser link establishment of two adjacent satellites in a constellation orbital plane, satellites in the orbital plane and satellites in two adjacent orbital planes are deduced. On one hand, by introducing an orbital coordinate system, an inter-satellite azimuth angle and a pitch angle are defined; on the other hand, a mathematical model of an inter-satellite laser link establishment pointing angle is obtained based on a positioning orbit determination ephemeris data source of an on-satellite GPS. And the inter-satellite laser link establishment pointing technology can be realized through a two-dimensional rotating mechanism of the laser. The method can be widely applied to an inter-satellite laser link establishment pointing scene of a low-orbit mobile communication constellation.

The invention discloses a two-way rocket launching window searching method, system and device and a medium, and the method comprises the steps: rapidly calculating two nearest launching windows of an ascending orbit and a descending orbit of a rocket according to the attributes of a central body, target orbit plane parameters, the geographic coordinates of launching points and the flight duration of the rocket; according to the method, the problem is converted into numerical calculation based on space geometry, the crossing moment is directly solved by using the geometrical relationship between the normal vector of the orbital plane and the inertial position vector of the launching point location and adopting efficient numerical methods such as dichotomy and the like, and complex dynamic iteration is avoided. According to the method, the calculated amount is greatly reduced, rapid searching of the launching window is achieved, the method is suitable for an interstellar detection task scene needing rapid response or multiple times of calculation, and the real-time performance of task planning is remarkably improved.

The application discloses a space tether formation state intelligent sensing method for accurate observation, which comprises the following steps: step 1, establishing a dynamics model of the space tether formation; step 2, selecting a sensor and establishing an observation model; and step 3, using a learning-based particle filtering algorithm to intelligently sense the state of the space tether formation. The application takes a space three-body tether formation system as a research object and studies the state sensing problem of the formation system out of the orbital plane.

When a deorbiting satellite, which is one of artificial satellites belonging to a first satellite constellation, deorbits and drops from an orbit of a first orbital altitude zone, a second satellite constellation widens a relative angle of any adjacent two orbital planes of a plurality of orbital planes and thereby allocates a free orbit area in a second orbital altitude zone. The deorbiting satellite passes through the free orbit area of the second orbital altitude zone.

A satellite constellation forming system (100) forms a satellite constellation which is composed of a satellite group and in which the satellite group cooperatively provides a service. The satellite constellation has a plurality of orbital planes in each of which a plurality of satellites fly at the same nominal orbital altitude. A satellite constellation forming unit (110) continues providing the service while avoiding a collision between satellites by both or one of control of an orbital altitude and control of a passage timing of a satellite group flying in a region where the plurality of orbital planes intersect.

The invention belongs to the technical field of artificial intelligence and block chains, and relates to a low-orbit satellite constellation-oriented model collaborative credible training system and method. The system comprises a main chain and a plurality of fragments, the main chain is used for task release and model mutual learning, and the fragments are used for model distributed training; the method comprises the following steps: fragmenting the constellation according to the orbit plane distribution characteristics of the low-orbit satellite constellation and completing configuration; the nodes in the fragments are trained, evaluated and aggregated according to identities allocated by the intelligent contracts, aggregation model parameters are generated, node reputation updating is carried out, after the main chain obtains the aggregation model parameters uploaded by the fragments, a plurality of nodes with sufficient energy are selected to complete a model mutual learning process, and then optimized model mutual learning results are fed back to the fragments. And a final training result is obtained through multi-round iteration. According to the method, the interference of the malicious edge satellite nodes in the low-orbit satellite constellation on the model training process can be effectively inhibited.

The invention discloses a low-thrust spacecraft asymmetric halo orbit design method, and belongs to the field of spacecraft orbits. The problem that the periodicity and stability of the orbit of a low-thrust spacecraft cannot be guaranteed under the condition that the plane of the halo orbit is asymmetric is solved. The method comprises the following steps: calculating a halo orbit satisfying a resonance relation with a second gravitational body orbit period; the eccentricity rate of the second gravitational body orbit is continued from zero to the actual eccentricity rate through numerical continuation in the first step, and a zero-thrust resonance halo orbit is obtained; through numerical continuation in the second step, the thrust acceleration amplitude is continued from zero to a preset upper limit while the thrust direction of the spacecraft is kept in the sun-spacecraft connecting line direction; through numerical continuation in the third step, the thrust taper angle is continued from zero to the preset upper limit under the gravity plane coordinate system; and through numerical continuation in the fourth step, the thrust azimuth angle is continued from zero to a preset value under the gravity plane coordinate system, and finally the asymmetric halo orbit is obtained. The method is suitable for circular and elliptical restrictive three-body problems.

A spaceborne heterogeneous sensing federated learning method based on a spiking neural network comprises the following steps: firstly, initializing a global model according to a satellite-ground cooperative computing network architecture, and distributing the global model to satellite nodes of each orbit plane; thirdly, introducing a pulse neural network into each satellite node to serve as a bottom layer calculation engine, and performing low-power-consumption model training of differentiated step lengths on local remote sensing data; then, in each orbit plane, satellite nodes carry out local communication through inter-satellite links, based on the actual training step length of each satellite, a single-step normalized model update quantity and an orbit effective step length are calculated, and in-orbit model weighted aggregation for eliminating computing power bias is executed; and after a plurality of rounds of in-orbit aggregation, the ground station server collects each orbit model and executes cross-orbit global aggregation. According to the method, the problem of inconsistent federated learning targets caused by heterogeneous satellite computing power can be effectively eliminated while the satellite-borne edge computing energy consumption can be remarkably reduced, and the convergence speed and prediction precision of the model in a network environment are improved.

A method for planning in-orbit refueling tasks for coplanar GEO orbits. The present invention relates to a method for planning in-orbit refueling tasks for coplanar GEO orbits. The purpose of the present invention is to solve the problem that the existing in-orbit round-trip refueling for coplanar GEO orbits often adopts a single-layer or two-layer optimization strategy, which will lead to falling into local optimal solutions during the optimization process, and the convergence speed is not fast enough, making it difficult to find the optimal solution, resulting in the fuel consumption of the service satellite when traveling back and forth between the service station and the target satellite for refueling tasks not being able to reach the optimal value. The process is as follows: Step 1, establish a model for planning in-orbit refueling tasks for coplanar GEO orbits; Step 2, solve the model for planning in-orbit refueling tasks for coplanar GEO orbits according to the CGAPB three-layer optimization algorithm to obtain the optimal service sequence of the service satellite, the optimal service time sequence, and the optimal time node sequence for the service satellite to return to the service station. The present invention is used in the field of aerospace technology.