51 results about "Orbit+Face" patented technology

The invention provides a space-ground integrated hybrid routing method, including calculating the link distance between satellites in the orbital plane and based on the time-varying link distance between orbital satellites, establishing a time-varying two-way Manhattan satellite network model; The queue delay between the satellite and the neighbor satellite node and the queue delay between the satellite-ground link with the ground node; calculate the time interval that does not affect the network routing performance, and predict the link transmission time between the neighbor node and the second nearest neighbor node within the time interval Delay and queue delay; according to the satellite routing algorithm based on predicting the next nearest neighbor, calculate the link delay of satellite network routing; when the source terminal communicates with the destination terminal, based on the minimum mean square error prediction mechanism, predict each link on the ground route The queue delay of the route; under the 5G-satellite integrated architecture, the goal is to minimize the total link delay of the space-ground integrated hybrid route, and select the optimal path for the on-satellite 5G-satellite node and the off-satellite 5G-satellite node , to get the set of integrated hybrid routes of heaven and earth.

The invention discloses a method for rapidly estimating a long-term orbital rendezvous optimal speed increment under J2 perturbation. The method includes the steps of calculating the orbital element difference of a non-maneuvering spacecraft and a target at the rendezvous moment, calculating an optimal control amount when only the ascending node ascensional difference is eliminated, substituting the optimal control amount as the initial value into a control amount optimizing model for eliminating the semi-major axis, dip angle and ascending node ascensional difference at the same time to obtain a new optimal control amount, modifying the control amount according to the phase difference in an orbital plane, modifying the control amount according to the eccentricity ratio vector difference,and finally obtaining the estimated optimal rendezvous speed increment. The method realizes the purpose of rapidly estimating the optimal rendezvous speed increment when precise solution is not neededand is high in estimating precision and suitable for the situations with large initial ascending node ascensional difference from a target orbit and small element differences from other orbits.

One embodiment of the present invention discloses a low-orbit satellite dual-degree-of-freedom solar wing control method, including: S1: Acquiring the angular velocity of the satellite, and judging whether the absolute value of the angular velocity is greater than a speed threshold, and if greater, the solar wing's S5 is performed after the control commands of the rotation axis and the swing axis are set to stop and maintain the command, otherwise S2 is performed; S2: obtain the sun vector under the satellite body system, and convert it to the orbital coordinate system; S3: calculate the orbital plane and the sun vector The angle between the vectors, and judge whether the absolute value of the angle is less than the angle threshold, if less, set the solar wing to continuous sun tracking control, otherwise set it to a fixed angle sun tracking control; S4: judge the sun Whether the wing rotation axis and the swing axis are controlled at the same time, if they are controlled at the same time, go to S5, otherwise adjust the command of the rotation axis; S5: output the control commands of the rotation axis and swing axis to control the movement of the solar wing.

The invention discloses a method for designing a spectrum coexistence constellation between low-orbit constellation systems, comprising the following steps: obtaining the number of orbital planes of the first satellite constellation, the initial ascending node right ascension and ascending node of the first orbital plane and the second orbital plane Right ascension spread and ascending node right ascension drift rate, as well as the orbital height and orbital eccentricity of the second satellite constellation; determine the orbital plane number of the second satellite constellation according to the orbital plane number of the first satellite constellation; The initial ascending node right ascension of one orbital plane and the second orbital plane, the distribution of ascending node right ascension, and the number of orbital planes of the second satellite constellation Obtain the initial ascending node right ascension of each orbital plane of the second satellite constellation; according to The orbital inclination of the second satellite constellation is obtained from the ascending node right ascension drift rate of the first satellite constellation, the orbital altitude and the orbital eccentricity of the second satellite constellation. The second satellite constellation obtained according to the technical solution of the present application can easily adopt a space isolation method to realize spectrum coexistence with the first satellite constellation.

The invention provides an engineering design method for a navigation satellite constellation. The method comprises the following steps: determining the height of a satellite orbit; determining a tilt angle of the satellite orbit; determining the number of satellites; determining the number of orbit planes and a phase factor; determining initial right ascension of an ascending node; judging whether a user has a design demand of performing regional enhancement by adopting a GEO satellite or not, and if so, optimizing a station position of the GEO satellite; judging whether the user has a design demand of performing regional enhancement by adopting an IGSO satellite or not, and if so, optimizing a phase angle of the IGSO satellite, and determining an orbit plane of the IGSO satellite; establishing a constellation phased deployment policy; and establishing constellation backup policies including a constellation networking phase backup policy and a constellation operation phase backup policy. With the design method, the navigation constellation which meets user demands, is excellent in performance and small in satellite number, and has relatively good compatible interoperability with an existing GNSS can be designed.

The invention discloses a method for suppressing internal frequency interference of a low-orbit constellation system. The low-orbit constellation system includes satellites running on at least three orbital planes, wherein each orbital plane includes at least one satellite, including: calculating the first satellite The first right ascension of the intersection point of the beam edge and the earth's surface, where the first satellite is any satellite in any one of the at least three orbital planes; calculate the second right ascension of the intersection point of the beam edge of the second satellite and the earth's surface , where the second satellite is a satellite separated by two orbital plane distances from the first satellite, and the second satellite has the same latitude argument as the first satellite; calculate the same latitude argument u of the first right ascension and the second right ascension c , according to the calculated latitude argument u c Determine the satellite beam shutdown range. Compared with the prior art, the technical solution proposed by the present invention can effectively suppress the internal frequency interference of the low-orbit constellation system on the premise of realizing the full coverage of the ground beam by the constellation system in the high-latitude area.