%0 Journal Article
%T 捷联惯性/天文/雷达高度表组合导航<br>Strapdown inertial/celestial/radar altimeter integrated navigation
%A 高自谦
%A 王海涌
%A 王永海
%A 赵振平
%A 陈垦
%A 秦绪国
%J 北京航空航天大学学报
%D 2017
%R 10.13700/j.bh.1001-5965.2016.0859
%X 摘要 提出了一种捷联惯性/天文/雷达高度表的弹道导弹组合导航方法。针对传统SINS/星敏感器组合无法从根本上解决惯导速度位置误差发散的问题，引入RA测量数据，以海拔计算高度与海拔观测高度的差值作为新的量测量，并推导了全微分方程，结合姿态误差角建立4维观测模型，针对弹道中段导航，以SINS误差方程作为系统状态模型，通过扩展卡尔曼滤波（EKF）进行组合导航解算。仿真结果表明，当SINS精度为惯导级、星敏感器测量精度10″、RA测量精度50 m时，经过1 810 s的飞行，再入点时刻速度误差小于1 m/s、圆概率误差（CEP）为1.2 km，比传统SINS/CNS方法速度和位置误差分别减小了76.1%和65.0%。<br>Abstract：Aimed at ballistic missile, a strapdown inertial navigation system/celestial navigation system/radar altimeter (SINS/CNS/RA) integrated method was proposed. Since the velocity and position errors' divergent problem of SINS can not be fundamentally solved by conventional SINS/star tracker integrated method, altitude intercept between calculated sea level elevation and observed sea level elevation which was measured by RA is introduced and total differential equation can be deduced. The four-dimensional observation model combining altitude intercept with attitude angle errors and the state model of SINS error equation are established by using extend Kalman filter (EKF) based on midcourse phase navigation. The simulation results manifest that when SINS has an inertial precision grade, star tracker has measurement precision of 10″, and RA has measurement precision of 50 m, after 1 810 s' flight, the velocity error of reentry point is less than 1 m/s and the circular error probability (CEP) is 1.2 km, with a 76.1% decrease of velocity error and 65.0% decrease of position error compared with conventional SINS/CNS method.
%K 捷联惯性/天文
%K 雷达高度表
%K 组合导航
%K 弹道导弹
%K 扩展卡尔曼滤波(EKF)&lt
%K br&gt
%K strapdown inertial/celestial
%K radar altimeter
%K integrated navigation
%K ballistic missile
%K extend Kalman filter (EKF)
%U http://bhxb.buaa.edu.cn/CN/abstract/abstract14246.shtml