Filter Algorithm On Spacecraft Autonomous Navigation In Large Elliptical Orbit

With the extensive application of space technology, higher demands on theautonomous navigation capability of the spacecraft. The purpose of this paper is toinvestigate the autonomous navigation of spacecraft in the large elliptical orbit.Strap-down Inertial Navigation System (SINS) is a completely autonomousnavigation, whose navigation accuracy is high in a short time. But the navigationerrors will accumulate with time, resulting the decreasing of the navigation accuracy.Working for long hours under the conditions, other measuring devices are requiredto correct the error. For low earth orbit spacecraft, the availability of GPS system isgood, with the high navigation accuracy, so the integrate navigation mode of SINSand GPS is presented to correct the navigation parameter. For high earth orbitspacecraft, considered with the unavailable of GPS, put forward the method of starsensor to correct gyro drift, and achieve the purpose to improve system navigationaccuracy. The main contents of this thesis are as follows.Firstly, the navigation equations of SINS are established in the geocentricinertial coordinate. The calculation formulas of the perturbation of the earth’snon-spherical, the resistance of the air, the solar radiation pressure and the gravitationof the sun and moon are given, and simulation of the perturbations divided inlow-orbit and high-orbit. Apply Cowell perturbation method to get the spacecraftnavigation information, including position, velocity, angular velocity and specificforce information. Then the simulation models of the gyro and the accelerometer areestablished. The simulation results verify the characteristics of the SINS system thatcan’t work alone for long.Operation principles of the GPS system are analyzed, combining with the SINSsystem to establish the status error equation and the measurement equation of theSINS/GPS integrated navigation system. Then the kalman filtering techniques in theSINS-GPS system are studied. The simulation results show that the SINS/GPSsystem increased the system navigation accuracy, and verify the feasibility of theSINS-GPS integrated navigation system.For high orbit spacecraft, put forward the integrated navigation method usingthe SINS system and star sensor assisted. Extended Kalman Filter technique isdesigned to verify the performance of the system. Simulation results show that theaccuracy of this integrated navigation is higher than SINS system.