In this paper, an autonomous orbit determination method for satellite using a large field of view star sensor is presented. The simulation of orbit under atmospheric drag perturbation are given with expanded Kalman fi...In this paper, an autonomous orbit determination method for satellite using a large field of view star sensor is presented. The simulation of orbit under atmospheric drag perturbation are given with expanded Kalman filtering. The large field of view star sensor has the same precision as star sensor and a sufficient filed of view. Therefore ,the refraction stars can be observed more accurately in real time. The geometric relation between the refracted starlight and the earth can be determined by tangent altitude of the refraction starlight. And then the earth center can be determined in satellite body frame. The simulation shows that the precision of the mean square deviation of satellite’s position and velocity is 5m and 0.01m/s respectively. The calculated decrement of the semi-major axis in one day is close to the theoretical result, and the absolute error is in the range of decimeter when the altitude of orbit is 750 km. The simu- lateion of orbit of different initial semi-major axis shows that the higher the altitude of orbit is, the smaller the dec- rement of the semi-major axis is, and when the altitude of orbit is 1700 km the decimeter of the semi-major axis is 10-7 km.展开更多
基金Project CXJJ-84 supported by Science and Technology Innovation Foundation of Chinese Academy of Science
文摘In this paper, an autonomous orbit determination method for satellite using a large field of view star sensor is presented. The simulation of orbit under atmospheric drag perturbation are given with expanded Kalman filtering. The large field of view star sensor has the same precision as star sensor and a sufficient filed of view. Therefore ,the refraction stars can be observed more accurately in real time. The geometric relation between the refracted starlight and the earth can be determined by tangent altitude of the refraction starlight. And then the earth center can be determined in satellite body frame. The simulation shows that the precision of the mean square deviation of satellite’s position and velocity is 5m and 0.01m/s respectively. The calculated decrement of the semi-major axis in one day is close to the theoretical result, and the absolute error is in the range of decimeter when the altitude of orbit is 750 km. The simu- lateion of orbit of different initial semi-major axis shows that the higher the altitude of orbit is, the smaller the dec- rement of the semi-major axis is, and when the altitude of orbit is 1700 km the decimeter of the semi-major axis is 10-7 km.
