运载火箭上面级惯性与天文组合导航系统设计

针对运载火箭上面级惯性导航随时间累积而误差增大以至不能满足长时间工作要求的问题,对采用星敏感器和地球敏感器修正惯性导航误差的方案进行了研究。首先,导出了上面级常用坐标系定义和姿态转换矩阵。然后,根据惯性导航的误差传播特性、星敏感器测量方程和地球敏感器的模拟测量方程,给出了组合导航的状态方程和观测方程。最后,设计了基于Matlab/dSpace仿真平台的星敏感器在导航回路中的半物理仿真实验。实验结果表明,组合导航使惯性导航位置误差矢量和从1.1719×104m减小到1.0367×103m,速度误差矢量和从11.2827m/s减小到3.6626m/s,姿态误差从0.1°减小到5′,说明了该组合导航方案能够有效修正惯性导航时间累积误差,半实物仿真实验验证了惯性/天文组合导航方案的可行性与正确性。

Design of INS/CNS integrated navigation system for launch vehicle upper stage

As the errors of the Inertial Navigation System (INS) in a launch vehicle upper stage increase significantly with time and cannot meet the requirements of long working hours,an INS/CNS (Celestial Navigation System) integrated navigation system is studied. Firstly, the coordinate systems and the attitude transformation matrix are defined. Then, the state equation and measurement equation of the integrated navigation system are raised from the INS error propagation equations and measurements of the star tracker and earth sensor. Finally, a semi-physical simulation experiment system with the star tracker in the loop is designed based on Matlab/dSpace simulation environment. Experimental results indicate that the integrated navigation system decreases the inertial navigation position errors from 1.171 9×10⁴ m to 1.036 7×10³ m, the velocity errors from 11.282 7 m/s to 3.662 6 m/s, and the attitude errors from 0.1° to 5'. Furthermore,the experimental results show that the integrated navigation system can correct the inertial navigation errors effectively, and also confirm that the INS/CNS integrated navigation system is feasible and appropriate.