3 results about "Eclipse" patented technology

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 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.