一种改进的捷联惯性/天文组合导航方法  

A Improved SINS/CNS Integrated Navigation Method

作　　者：郎希开 何小飞 姬占礼 何竹 吉云飞 Lang Xikai;He Xiaofei;Ji Zhanli;He Zhu;Ji Yunfei(Beijing Aerospace Times Optical-electronic Technology Co.Ltd,Beijing 1000941,China;College of Information Systems and Management,National University of Defense Technology,Changsha 410073,China)

机构地区：[1]北京航天时代光电科技有限公司,北京100094 [2]国防科学技术大学信息系统与管理学院,长沙410073

出　　处：《战术导弹技术》2018年第6期95-100,共6页Tactical Missile Technology

基　　金：国家自然科学基金(61573113)

摘　　要：针对传统的捷联惯性/天文组合导航系统(SINS/CNS)不能精确估计加速度计偏置而导致导航误差发散的问题,提出了一种改进的组合导航方法。该方法将与位置相关的高度角误差、方位角误差以及由气压高度表得到的高度误差引入到系统的量测方程中,从而抑制了位置误差的发散,解决了传统方法中加速度计估计位置误差时存在的问题。根据上述原理以弹道导弹为载体推导了组合系统的线性模型,最后通过卡尔曼滤波实现了状态估计。仿真结果表明:该方法的导航精度优于传统方法,有效抑制了由加速度计偏置造成的位置误差,验证了该方法的有效性。Considering that traditional strap-down inertial/celestial integrated navigation method( SINS/CNS) cannot accurately estimate the accelerometer bias,which can cause the divergence of navigation errors,a modified integrated navigation algorithm is proposed. The proposed method introduces the height angle error and azimuth angle error related to position and the height error which is obtained by the pressure altimeter into the measurement equation of the system. The method inhibits the divergence of position error and solves the problem existing in the accelerometer estimation of the position error in the traditional method. Based on the above principle,the linear model of the integrated navigation system which is used in the ballistic missile is constructed. Finally,the estimation of the state of system is realized through the Kalman filter. The simulation results indicate that the precision of navigation based on the proposed method is better than that of the traditional method and corrects the navigation error caused by the accelerometer bias effectively,which shows the validation of the proposed method.

关 键 词：组合导航 高度角 方位角 导航误差 卡尔曼滤波 

分 类 号：V294.3[航空宇航科学技术]