34 results about "Orbit plane" patented technology

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 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 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.

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.

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.

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.

The application discloses a kind of unloading method and device of asymmetric configuration satellite angular momentum, the method includes: by solar sensor, the solar vector under satellite body system is acquired, and the solar vector is converted into the solar vector under orbit system;The included angle β of the solar vector under the orbit system and orbit plane is calculated, and the rotation angle of solar wing rotation axis and the preset angle of biasing axis are set according to the included angle β of the solar vector under the orbit system and orbit plane;The angular momentum of flywheel system is calculated and is converted into the angular momentum under satellite body system;According to the modulus of satellite to be unloaded angular momentum and preset unloading threshold value are compared, according to the comparison result, it is judged whether magnetic moment device unloading needs to be started, and according to the comparison result, the control instruction of magnetic moment device is output.The application effectively reduces angular momentum accumulation and space moment accumulation;With the advantages of saving fuel, low power consumption requirement and high attitude stability, it provides a theoretical basis for angular momentum unloading of weak magnetic field planet explorer.

The application discloses a method for calibrating a roll angle of an emergency anti-solar time of a satellite on the earth, which is beneficial to reestablishing a three-axis attitude of the satellite on the earth. Firstly, an angle between a projection of a sun vector on an orbit plane and an orbit system-X axis is calculated; then, whether a current orbit position meets a requirement is determined according to the angle between the projection of the sun vector on the orbit plane and the orbit system-X axis, and whether an absolute attitude reference is available at present is determined, if the current orbit position meets the requirement and the absolute attitude reference is available at present, the roll angle is calibrated at a current control beat; otherwise, the roll angle is not calibrated at the current control beat; if the roll angle is calibrated at the current control beat, the roll angle is calculated based on the absolute attitude reference; if the roll angle is not calibrated at the current control beat, the roll angle is calculated by using gyro inertia angular velocity integration. The application can provide high-precision and reliable attitude information for reestablishing the attitude of the satellite on the earth.

The invention discloses an automatic division method and system for a satellite constellation orbital plane, and relates to the technical field of spaceflight measurement and control. The method comprises the following steps: firstly, calculating an orbital plane unit normal vector based on a satellite orbit inclination angle and an ascending node right ascension, and performing first-level spatial pointing clustering by adopting a density clustering algorithm to form an orbital plane group; secondly, in the same orbital plane group, by taking a semi-major axis and an eccentricity rate as characteristics, automatically determining an optimal shell layer number by utilizing a Gaussian mixture model and a Bayesian information criterion, and realizing second-stage orbital energy shell layer division; and finally, in the same shell layer, sorting and interval statistical analysis are performed on the right ascension of the ascending node, recursive subdivision is performed through dynamic threshold segmentation in combination with a range-based dual-condition verification mechanism, and fine division of a third-level orbital plane is completed. The method realizes full-process automation from orbit data to division results, has the advantages of high division precision, strong adaptability, good engineering practicability and the like, and is suitable for operation and maintenance management of large-scale heterogeneous satellite constellations.

The application provides a kind of super low orbit satellite intelligent orbit control method and system, comprising the following steps: calculating the orbit plane semi-major axis deviation Δa and the descending node position deviation ΔL;According to the orbit plane semi-major axis deviation Δa and the descending node position deviation ΔL, control coefficient K is calculated;According to the control coefficient K, deviation threshold ha, hL is calculated;It is judged whether "deviation Δa exceeds deviation threshold ha" or "deviation ΔL exceeds deviation threshold hL", if it exceeds, the control instruction of thruster is calculated according to the orbit plane semi-major axis deviation Δa, the descending node position deviation ΔL and the control coefficient K, otherwise, the control instruction of thruster is zero.The application can make the satellite realize intelligent autonomous orbit height maintenance control when in orbit, and the precision of orbit maintenance can be significantly improved, so as to save the fuel of satellite carried thruster and prolong the in-orbit life of satellite.

The application discloses a kind of Walker constellation's transit time window fast calculation method to ground area.This method is for the fast calculation of ground area transit time, first, given time range information, constellation orbit information and ground area information are acquired, second, the intersection of orbit plane and ground area is analyzed, the time window of satellite possible transit event is screened, then, the characteristics of Walker constellation satellite uniform distribution on orbit plane are combined, the satellite orbit position in potential transit time window is quickly calculated, finally, the transit time window is obtained by judging whether the subpoint is inside or outside the area.The method can quickly calculate the time window of satellite transit ground area by analyzing the intersection of orbit plane and ground area, and combining the characteristics of Walker constellation satellite uniform distribution on orbit plane.The advantages of the method are: (1) significantly improve the calculation efficiency; (2) suitable for large-scale Walker satellite constellation.

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

The application discloses a high-precision space target orbit prediction method based on numerical fitting, and relates to the technical field of space target orbit prediction.The frequency domain decomposition of the application is combined with a harmonic compensation model to completely eliminate seasonal system drift caused by the gravitation of the sun and the moon; high-frequency components are stripped of attitude noise through light pressure parameterization and smooth difference to avoid the transmission of perturbation coupling error; a harmonic order adaptive mechanism captures the main disturbance term with the least amount of calculation; least square batch estimation realizes one-time solution of coefficients; template library preloading technology effectively compresses the time consumption of new target initialization; in addition, the application introduces a geomagnetic modulation factor to dynamically suppress compensation distortion during magnetic storms, and a sliding window mechanism guarantees real-time updating of coefficients, and stable prediction is still maintained under extreme events; photometric curve inversion is adopted instead of attitude measurement to make non-cooperative target prediction possible; tangential component directional compensation ensures the stability of the orbit plane and reduces satellite position fuel consumption.

An electronic device, according to one embodiment of the present disclosure, comprises: a communication circuit; a display; a memory storing instructions; and at least one processor, wherein the instructions, when executed individually or collectively by the at least one processor, may instruct the electronic device to: identify an orbital plane of a satellite that has entered a communicable area; determine, on the basis of the orbital plane, a service-available time for communication using the satellite; store an actual service time using the satellite; identify an inter-satellite gap time on the basis of the actual service time; and provide, via the display, schedule information of a non-terrestrial network service related to at least one of the determined service time, the actual service time, or the inter-satellite gap time.

The application relates to a minimum relative image movement imaging method of a low-orbit remote sensing satellite on a high-orbit target, and solves the problem of poor imaging effect on the high-orbit target. The application is convenient for the low-orbit remote sensing satellite to perform long-exposure imaging on the high-orbit target under the condition that the low-orbit remote sensing satellite is in an inertial space stable attitude. The method is as follows: the low-orbit remote sensing satellite observes a high-orbit target to be observed at an angle alpha, the relative speed of the high-orbit target to be observed relative to the low-orbit remote sensing satellite is zero in the projection of the relative direction normal; the high-orbit target to be observed is located on the orbit plane of the low-orbit remote sensing satellite, and the normal is located on the orbit plane; the angle alpha satisfies:

A satellite constellation (110) includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

The application provides a collision risk satellite combination processing sequencing optimization method, steps are as follows: according to TLE orbit root number data, orbit pre-division is carried out, satellites located on the same orbit plane are divided together; a perigee-apogee pre-screening method is used to carry out pre-screening of satellite collision combination, and satellite and debris combinations generating collision risks are obtained; TLE orbit root number data is combined with an orbit prediction model and a fast screening trajectory space intersection algorithm to determine the position of trajectory intersection and overlap in space; high collision risk satellite combinations and high collision risk regions are determined; a collision combination priority processing sequencing model is constructed in combination with satellite states and satellite characteristics; and a gradient descent method is used to optimize parameters of the multi-factor collision combination priority processing sequencing model. The application improves the screening speed of high collision risk combinations and improves the overall safety of a satellite constellation by using the geometric screening method of the intersection angle between two orbits and the gradient descent method to optimize multi-factor parameters.

The present application relates to a low-orbit satellite network switching optimization method and system based on a named data network, the system comprising: obtaining the number of orbital planes, the number of satellites in a single orbit and the topological connection relationship between satellites; obtaining a consumer access satellite set and a producer access satellite set that can provide services for consumers within the total service time; determining the average routing hop count between each consumer access satellite and the producer access satellite within its own service time slot; constructing a low-orbit satellite network switching optimization model with the minimum average routing hop count as the optimization target; solving the low-orbit satellite network switching optimization model based on the preset minimum hop count graph switching strategy to obtain the shortest path switching scheme. This makes the consumer side switch different access satellites while meeting the minimum average routing hop count constraint and selecting the shortest path, ensuring the minimum delay of the consumer during network communication.

A positioning computer for a satellite intended to move in an orbital plane, the satellite being intended to include: a body having an Earth face defining an axis Z of the satellite intended to point towards the Earth, 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 and have a field of view in azimuth around the axis Z, the field of view being limited by an obturation angle due to the body. The computer is adapted to obtain a parameter representing a solar angle defined by the Sun with the orbital plane, and to perform the calculation of a parameter representing the yaw angle of the satellite using at least the parameter representing the solar angle and a parameter representing the obturation angle.

The invention discloses a double-layer LEO satellite network cross-layer link establishment method and system based on pushing of a talker, and relates to the technical field of wireless communication networks, the method comprises the following steps: discretizing system time according to a preset scheduling period, and completing a cross-layer link establishment decision facing a next scheduling period in a current scheduling period; dividing lower-layer access satellites into regions according to orbit planes, calculating a state evaluation value based on operation state information in each region, and pushing satellites of talker persons; the talker satellites report states to a ground control center, and the ground control center performs visibility judgment based on ephemeris data and distributes target upper-layer backbone satellites to the talker satellites; and establishing a cross-layer link and updating a cross-layer topology according to a distribution result when the next scheduling period starts. According to the technical scheme, on the premise that the constellation structure is not changed, the cross-layer link establishment scheduling scale and the signaling overhead are reduced, and the cross-layer link availability and the network service bearing capacity are improved.

The invention relates to an orbit control method for different orbit plane low-thrust satellite formation rapid networking. Relates to the field of near-earth orbit satellite formation, in particular to an orbit control method for different-orbit-plane low-thrust satellite formation rapid networking. The control quantity which must be completed through thrust is converted into the drift distance which can be naturally completed through the orbit dynamics process, correction is carried out with the assistance of a small amount of direct control, and therefore the total control frequency and fuel consumption are remarkably reduced. The method comprises the following steps: acquiring a formation flying target; relevant parameters required by orbit control are calculated; constructing an ascending node right ascension control change model, and establishing an ascending node right ascension change equation; solving an ascending node right ascension change equation to obtain a minimum control circle number and a standard control vector; calculating a phase drift rate according to the adjustment time required by the right ascension of the ascending node; and combining the control vectors to form a final control plan.

The application discloses a method for constructing a cross-orbit chain topology of a mega-constellation network, and the method comprises the following steps: S1, acquiring the number of uniformly distributed orbit planes in the mega-constellation network P and the number of uniformly distributed satellites in each orbit plane S ; S2, each satellite on each orbit plane establishes an inter-orbit satellite interlink with the satellites adjacent to the orbit plane, meanwhile, each satellite on each orbit plane establishes an inter-orbit satellite interlink with the corresponding satellite on the corresponding orbit plane by selecting at least two cross numbers to form a cross-orbit chain topology of the mega-constellation network, and the greatest common divisor of all cross numbers is 1. The cross-orbit chain topology of the application makes the inter-orbit satellite interlink span the satellites of several orbit planes, under the premise of ensuring global connectivity, significantly improves the long-distance transmission efficiency, breaks through the scale bottleneck of a single-layer constellation, reduces the dependence on the ground gateway station, and makes the mega-constellation network become a high-performance infrastructure complementary to the ground network.

This invention discloses a topology design method, system, electronic device, and storage medium to ensure the capacity robustness of satellite networks. It primarily addresses the shortcomings of existing technologies that do not fully consider the impact of network topology on capacity robustness and the difficulty in efficiently selecting suitable topologies as the network scales up. The implementation scheme involves: obtaining the parameters of the satellite network to be designed and the failure probability of inter-satellite links; setting a connection factor based on the inter-satellite link failure probability and determining the offset of each orbital plane that satisfies the connection factor constraint; establishing connection rules for intra-orbit links within the orbital plane based on the inter-satellite link distance; and establishing connection rules for inter-orbit links between orbital planes based on the inter-satellite link distance and offset, thereby obtaining a satellite network topology that meets the capacity robustness requirements. By introducing link failure probability constraints during the topology design stage, this invention effectively enhances the capacity of bottleneck links and ensures the network capacity robustness of the satellite network in the event of link failure, and can be used for satellite network topology design.

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