7 results about "Eclipse" patented technology

There are provided methods of processing satellite state data, comprising receiving satellite state data in the form of multiple separate files via one or more ground stations and compiling the received satellite state data into a single dataset accessible via an application programming interface and searchable by time range. There are further provided methods of processing satellite state data comprising receiving raw satellite state data; receiving manoeuvre data relating to one or more scheduled manoeuvres of the satellite; and filtering the received raw satellite state data in an orbit determination process to provide filtered satellite state data, wherein the manoeuvre data is used in the filtering of the received raw satellite state data. There are further provided methods of scheduling a satellite manoeuvre comprising: receiving parameters for one or more planned manoeuvres to move the satellite from a current orbit to a new orbit, wherein the parameters include a time and duration of each of the one or more planned manoeuvres; receiving times of eclipses of the Sun by the Earth during future orbits of the satellite; and scheduling the manoeuvre to take place according to the determined parameters and the times of eclipses.

The invention discloses a mechanical arm precise shaft hole assembling method based on lunar eclipse heuristic scene observation and simulation. A visual sensor installed at the tail end of a mechanical arm is adopted for collecting image data of a to-be-assembled hole piece; performing target detection and positioning on a to-be-assembled hole piece in the image data through a deep learning-based YOLOv8 model, performing image acquisition on the to-be-assembled hole piece again by the visual sensor, and inputting a lunar eclipse assembly scene into a YOLOv8-seg network model for detection to obtain instance segmentation output information of the lunar eclipse assembly scene, the instance segmentation output information is input into a simulation learning strategy network model, and corresponding mechanical arm execution actions are output according to different lunar eclipse assembly scenes; and operating the mechanical arm to execute the action, returning to the process of obtaining the lunar eclipse assembly scene, inputting the obtained lunar eclipse assembly scene into the YOLOv8-seg network model to be continuously executed, finally obtaining the mechanical arm execution action, and repeatedly executing until the lunar eclipse assembly scene cannot be obtained, so that the alignment process is completed.

The application provides a method and system for evaluating the stability margin of a long-distance retrograde orbit spacecraft. First, the current orbit state information of the target spacecraft is obtained. Then, based on the evaluation allowed time range, evaluation time interval, speed error size range and speed error size interval, an error parameter combination set is constructed. Next, based on the current orbit state information and the error parameter combination set, the stability evaluation parameters and the eclipse evaluation parameters corresponding to the error parameter combination are determined through a high-precision ephemeris model. Second, based on the stability evaluation parameters and the eclipse evaluation parameters, the target error parameter combination is determined. Finally, the stability margin is determined based on the target error parameter combination. This method combines the characteristics of long-distance retrograde orbits, engineering application requirements and the influence of eclipses, and selects parameters through the error parameter combination set to determine the stability margin. In this way, the efficiency and accuracy of determining the stability margin of the spacecraft can be effectively improved.

The application discloses a spacecraft-oriented radiator design method and system, and belongs to the technical field of spacecraft thermal control. In order to solve the problems of non-uniform area and aperture, inaccurate extreme condition identification and lack of cross-validation in the prior art, the method comprises the following steps: obtaining an orbit and a radiator parameter; calculating an orbit position, a solar vector and an eclipse; independently calculating five types of external heat flow sub-items of a solar direct radiation, an earth albedo, an earth infrared, a deep space radiation and an earth radiation on a front surface and a back surface of the radiator, and summing up to obtain a total net heat flow density; according to a solving mode, a first mode of inversely solving a radiator area with a fixed temperature constraint is executed, or a second mode of inversely solving a balance temperature with a fixed area is executed; and finally, energy closure diagnosis and sub-item conservation checking are performed. Through double-sided sub-item modeling, reciprocal solving and three-stage extreme condition searching, the application realizes the improvement of design accuracy, the reduction of calculation cost and the quantitative evaluation of model self-consistency, and is suitable for the thermal control design of various low earth orbit spacecrafts.

The invention discloses a data processing and identification method based on space-based optical measurement, comprising the following steps: S1, reading space-based optical data; S2, judging whether the data is reasonable, eliminating the unreasonable data, and entering step S3 if it is reasonable; S3, performing data conversion on the reasonable data; S4, determining the initial orbit; S5, if the initial orbit is successfully determined, using the orbit determination result to compare and identify with the TLE in the target set, judging the threshold conditions according to the semi-major axis, eccentricity, and ascending node right ascension, and obtaining suspected targets that meet the threshold conditions; if the initial orbit is determined unsuccessfully, extrapolating and inverting the TLE in the target set, comparing the pointing data with the measured data, giving the right ascension and declination judgment threshold conditions, and selecting suspected targets that meet the threshold conditions; S6, further eliminating the suspected targets through identification rules, finding the closest target, giving a confidence evaluation, and outputting the identification results. The invention solves the problem of large amount of space-based observation data and difficulty in identification.

The application provides a sunlight direct irradiation avoiding method and system for intersatellite laser communication. The method comprises the following steps: judging the included angle in real time according to the satellite broadcast sun vector and the current optical axis pointing vector, and autonomously entering the sun-eclipse avoiding mode if the included angle is less than 3 degrees; converting the sun vector to the azimuth-elevation two-dimensional coordinate system of the optical antenna, and determining the azimuth, elevation value and azimuth-elevation change direction of the sun vector; and autonomously ending the sun-eclipse avoiding and re-establishing the chain after the included angle between the pointing vector and the sun vector is greater than 3 degrees. The technical scheme provided by the application does not need to be predicted, the included angle is calculated in real time according to the current laser terminal optical axis pointing and the satellite broadcast sun vector, the sun-eclipse avoiding mode is autonomously entered if the included angle is less than 3 degrees, and no manual instruction is needed, so that the sunlight direct irradiation can be effectively avoided in the avoiding process.

The application relates to a sun-eclipse-avoiding method and system for a spaceborne laser terminal based on optical path switching and pointing maintenance, which comprises the following steps: based on satellite attitude, orbit data and laser terminal pointing data, a real-time included angle theta between a sun vector and a laser communication link vector is calculated in real time; a first protection threshold theta1 is preset, and it is judged whether physical optical path switching is started; when the real-time included angle theta is less than or equal to theta1, it is determined that it is in a sun-eclipse danger stage, and the incident optical path is switched from a communication receiver to a light shield plate area; during the optical path switching, the pointing of the laser terminal to the communication target is maintained unchanged; the real-time included angle theta is continuously monitored, and when theta is greater than theta1, the optical path is restored from the light shield plate area to the communication receiver; and after the optical path is restored, the communication link is recovered in stages based on the unchanged terminal pointing, so that the communication link is re-established. Through the cooperative mechanism of hardware optical path switching and pointing maintenance, the communication receiver is quickly protected, and the link is quickly recovered.