46 results about "Orbital plane" patented technology

The invention discloses a calculation method of subsatellite points and photographic point trajectory self-intersection points of a near-earth regression orbit satellite. Based on an orbital perturbation model of the earth oblateness (J2), the mean influence of the earth oblateness on the long-term perturbation of the right ascension of the ascending node, the perigee argument and the aplanatic angle is utilized, so that the mean drift period and the orbital intersection period of the satellite orbital plane relative to the earth resonate, and revisit regression of the orbital relative to thesubsatellite point and the photographic point trajectory of the earth surface is formed. After the orbit inclination angle, the eccentricity ratio, the regression days and the period number are given,the orbit semi-major axis can be solved, then the longitude of the sub-satellite point trajectory self-intersection point is analyzed and solved, the latitude of the self-intersection point under thecorresponding precision is screened out through numerical screening, and the distribution situation of the sub-satellite point trajectory self-intersection point in one regression period can be obtained. The invention provides a practical near-earth regression orbit sub-satellite point trajectory self-intersection point calculation method, and the method has important application value in near-earth surveying and mapping, remote sensing, reconnaissance, InSAR and other regression orbit satellite measurement tasks.

The invention relates to a satellite positioning technology, and discloses a measurement, operation and control system and method for a low-orbit satellite constellation. The measurement, operation and control system comprises N common-orbit inter-satellite links corresponding to N orbital planes, each common-orbit inter-satellite link is used for data transmission between node satellites of the corresponding orbital plane, and N is larger than or equal to 2; and the system further comprises a satellite-ground link which is used for specifying the node satellites running to the intersection ofthe N orbital planes in each common-orbit inter-satellite link as hub satellites, and each common-orbit inter-satellite link performs data transmission with at least one ground station arranged nearthe latitude of the intersection of the N orbital planes through the hub satellites. According to the embodiment of the invention, while the low-orbit satellite constellation measurement, operation and control requirements are met, the technical complexity of the satellite is reduced, the cost of the satellite and the ground station is reduced, and the engineering feasibility of real-time measurement, operation and control of the low-orbit satellite constellation is improved.

The invention discloses a satellite steady sun-facing orientation method taking earth pointing deviation as constraint, and belongs to the technical field of spacecraft attitude control. The stable sun-facing orientation method comprises the following two steps: establishing an expected attitude of a satellite by two steps: firstly, establishing an intermediate attitude of the satellite on the basis of a satellite-geocentric connecting line direction, a satellite orbital plane normal direction and a satellite advancing direction; and then rotating the middle attitude around an Euler axis for acertain angle to reduce the sun pointing deviation of the satellite and ensure that the included angle between the expected earth axis and the earth center direction is not greater than the constraint angle. The expected attitude of the satellite obtained according to the strategy can keep stable change, the singular phenomenon that the expected attitude of the satellite violently changes in a short time under the sun-earth-satellite collinear condition caused by a traditional strategy is avoided, and therefore peak energy consumption in the sun-oriented process of the satellite is greatly reduced.

Provided is a method and system of evaluating a constellation spare strategy based on a stochastic time Petri net, which is applied in the technical field of constellation operation management. The method comprises: constructing a single satellite STPN model and an orbital plane STPN model, and establishing a navigation constellation STPN model that includes multiple spare strategies according to the single satellite STPN model and the orbital plane STPN model; establishing an availability model according to the number of malfunctioning satellites and the constellation value (CV) in the navigation constellation STPN model, and establishing a cost model according to operating costs of the navigation constellation STPN model; and evaluating the navigation constellation STPN model using the availability model and the cost model, and determining a target spare strategy from the multiple spare strategies according to an evaluation result.

The invention provides a continuous low-thrust high-orbit target orbital transfer approaching method and system. The continuous low-thrust high-orbit target orbital transfer approaching method comprises the following steps that S1, calculating an inclination angle vector and adjusting the speed increment of the inclination angle vector to judge the approaching type; s2, solving a continuous low-thrust orbital transfer strategy by adopting an in-plane phase modulation mode for coplanar approaching; s3, for different-plane small-dip-angle approaching, calculating a low-thrust orbital transfer strategy in the mode that the orbital plane is adjusted firstly and then in-plane phase modulation is used; and S4, and calculating a low-thrust orbital transfer strategy by approaching a target at an ascending/descending intersection point for different-plane large-dip-angle approaching. According to the method, the problems of coplanar approaching, different-plane small-dip-angle approaching and different-plane large-dip-angle continuous low-thrust approaching of the high-orbit target are effectively solved.

The invention discloses a satellite determination method and device based on a Doppler effect. The method comprises the steps of selecting a plurality of satellites which are most similar to orbital parameters of an orbital plane and have satellite topping time difference within set time as alternative satellites; sorting the alternative satellites according to the satellite topping time, and taking the satellite of which the topping time is ranked in the middle as a reference satellite; adjusting the tracking angle of a ground station antenna to always align the antenna with the reference satellite; calculating the Doppler frequency shift corresponding to the reference satellite, and carrying out compensation correction on the received satellite downlink signal by taking the Doppler frequency shift as a compensation amount; and determining a signal source satellite corresponding to the signal according to the frequency deviation condition of the compensated and corrected satellite downlink signal. According to the satellite determination method and device based on the Doppler effect, the signal source satellite corresponding to the signal can be determined when the received satellite signal cannot be analyzed on the ground or the satellite signal does not contain the real-time position information of the satellite, the implementation process is simple, and the cost is low.

The invention provides a routing table distribution method and device of a low-orbit constellation system. The method comprises the steps of when a ground management and control center needs to updatea satellite routing table, obtaining routing tables of all satellites on an orbital plane where a transit satellite is located; splitting the routing table into a plurality of first routing table data packets through the ground management and control center; sequentially sending the plurality of first routing table data packets to the transit satellite; after the transit satellite receives all the first routing table data packets, integrating all the first routing table data packets into a routing table; splitting the routing table into a plurality of second routing table data packets throughthe transit satellite; and respectively distributing other second routing table data packets, except the second routing table data packet of the transit satellite, in the plurality of second routingtable data packets to all satellites on the orbital plane where the transit satellite is located through the transit satellite. According to the invention, the routing table uploading efficiency can be improved, the routing table is distributed by fully utilizing the fixed same-orbit inter-satellite link, the link is stable, and the operation is simple.

The invention provides a large-scale low-orbit satellite constellation deployment method, device and equipment and a storage medium. The method comprises the steps of controlling a first group of satellites to enter a nominal orbit to form a first orbit surface after a rocket is separated from a plurality of satellites, controlling the second group of satellites to the Nth group of satellites to stay in a parking orbit, wherein the multiple satellites are divided into N groups, and the height of the nominal orbit is larger than that of the parking orbit, determining the adjustment time of each satellite in the ith group of satellites based on the right ascension difference of the ascending node between the first orbital plane and the ith orbital plane to be formed by the ith group of satellites, wherein the adjustment time comprises the parking time and the climbing time, and based on the parking time, controlling each satellite in the ith group of satellites to stay in a parking orbit, and based on the climbing time, controlling each satellite in the ith group of satellites to climb to a nominal orbit from the parking orbit to form the ith orbit surface so as to form a satellite constellation.

The embodiment of the invention discloses a low-orbit satellite two-degree-of-freedom solar wing control method, which comprises the following steps: S1, acquiring the angular velocity of a satellite, judging whether the absolute value of the angular velocity is greater than a velocity threshold value, if yes, setting a control instruction of a rotating shaft and a swinging shaft of a solar wing as a stalling keeping instruction, and then performing S5, otherwise, performing S2; S2, acquiring a sun vector in a satellite body system, and converting the sun vector into an orbital coordinate system; S3, calculating an included angle between the orbital plane and the sun vector, judging whether the absolute value of the included angle is smaller than an angle threshold value or not, if so, setting the solar wing as continuous sun tracking control, and otherwise, setting the solar wing as fixed-angle sun tracking control; S4, judging whether the rotating shaft and the swinging shaft of the solar wing are controlled at the same time, if yes, executing S5, and otherwise, adjusting the instruction of the rotating shaft; and S5, outputting control instructions of the rotating shaft and the swinging shaft to control the solar wing to move.

The invention belongs to the technical field of atmospheric density sensing methods and space physics, and particularly relates to a hot layer atmospheric density prediction method based on distributed orbit atmospheric sensing units, which comprises the following steps: obtaining P distributed orbit atmospheric sensing units; in combination with an energy dissipation rate or semi-major axis attenuation, calculating an atmospheric density correction ratio of a track surface thermal layer where each distributed track atmospheric sensing unit is located; periodically and dynamically selecting the multiple space targets are uniformly distributed in the P distributed orbit atmosphere sensing units at different places; in all P distributed track atmosphere sensing units, calculating the atmospheric density of the space point of the space target on the corresponding orbital plane every 2-30 seconds; periodically acquiring a full-space atmospheric density data set {rhoO} within 3-24 hours; calculating a spherical harmonic coefficient, calculating the corrected inflection point temperature and escape layer temperature to form a thermal layer atmospheric density correction model, and obtaining the dynamically corrected thermal layer atmospheric density by using the correction model to predict the atmospheric density of any position in the future space.

The invention discloses a method for realizing narrow view field load pointing calibration by satellite attitude maneuver assistance. The method comprises the following steps that a target staring attitude is established according to orbit parameters of a local satellite and a target satellite and a theoretical value of a load installation matrix; in a vertical load visual axis plane in a satellite staring coordinate system, an Archimedes spiral is calculated from a central point, and the scanning direction of a load visual axis is planned in real time; a scanning attitude of a satellite relative to a staring coordinate system in real time is calculated according to a load visual axis direction planned by a spiral line by taking an orbital plane normal as a reference vector, a target attitude of the satellite relative to an orbital coordinate system is calculated in combination with the staring attitude, and an expected angular velocity of the satellite is obtained; the deviation of the attitude and the angular velocity is calculated according to the current attitude and the angular velocity of the satellite so as to realize spiral line scanning in the load signal direction; after the two parties capture the signals, the current attitude is recorded, the actual measurement value of the installation matrix corresponding to the load is calculated in combination with the theoretical attitude, and normal communication of the narrow-view-field load can be realized without spiral scanning again.

The embodiment of the present invention discloses a geomagnetic energy storage low orbit space debris off-orbit delivery orbit attitude coupling adjustment method, which includes the following steps: Step 100, decomposing the entire geomagnetic energy storage low orbit period into energy gathering arcs and energy releasing arcs section, and configure the momentum wheel energy-absorbing device on the three axes of the spacecraft; step 200, respectively determine the energy-gathering arc section and the energy-discharging arc section when the spacecraft performs energy storage and accumulation around the y-axis of the orbital plane and the z-axis of the orbital plane; Step 300, after determining the energy-gathering arc segment, for the energy storage targets of different orbital coordinate systems Y-axis or Z-axis, complete the de-orbiting operation of space debris based on the geomagnetic energy storage low-orbit space debris de-orbiting control method; Step 400, After the energy release arc is determined, the three-axis momentum wheel energy absorbing device of the spacecraft is unloaded through the generated magnetic moment.

The invention relates to the technical field of spatial information networks, and provides a laser link distribution method for optimizing network throughput, which comprises the following steps: carrying out visibility analysis; selecting orbital planes of the IGSO satellite and the MEO satellite to establish a link; selecting an MEO satellite to establish a link with a ground base station and calculating communication overhead; recording a link establishment routing table; and circulating the steps and recording the link establishment routing table with the least output hops. According to the method, the balance between the on-satellite data and the landing data is completed through distribution of the laser links, the problem that the landing amount of the on-satellite data is small is solved, the inter-satellite network has good throughput, and reference is provided for networking of the next generation of Beidou navigation system.

The invention relates to a method for designing a medium-orbit elliptical orbit remote sensing satellite constellation, which comprises the following steps of: S1, designing an elliptical frozen orbit of a medium-orbit elliptical orbit remote sensing satellite by utilizing the characteristic that the elliptical orbit of the medium-orbit elliptical orbit remote sensing satellite stays at an apogee for a long time; the method specifically comprises the following steps: S11, selecting an orbit inclination angle i; s12, selecting perigee height hp and apogee height ha; s13, selecting a perigee argument omega; s14, selecting the right ascension omega of the ascending node; s2, aiming at the characteristic that the elliptical orbit of the medium-orbit elliptical orbit remote sensing satellite is a non-uniform orbit with high perigee movement speed and low apogee movement speed, designing a special Walker constellation of the medium-orbit elliptical orbit remote sensing satellite; the method specifically comprises the following steps: S21, calculating a mean anomaly of a satellite in an orbital plane; s22, calculating the number s of satellites in each orbital plane; and S23, calculating the plane anomaly difference of the adjacent orbital planes. The method has the advantages that the method is oriented to optical remote sensing satellites, targets in key areas are highly revisited, and hotspot targets are monitored with high time resolution and high spatial resolution.