北京航空航天大学学报 ›› 2017, Vol. 43 ›› Issue (12): 2505-2512.doi: 10.13700/j.bh.1001-5965.2016.0859

• 论文 • 上一篇    下一篇

捷联惯性/天文/雷达高度表组合导航

高自谦1, 王海涌1, 王永海2, 赵振平2, 陈垦2, 秦绪国2   

  1. 1. 北京航空航天大学 宇航学院, 北京 100083;
    2. 北京航天长征飞行器研究所, 北京 100076
  • 收稿日期:2016-11-09 修回日期:2017-03-06 出版日期:2017-12-20 发布日期:2017-04-07
  • 通讯作者: 王海涌 E-mail:why@buaa.edu.cn
  • 作者简介:高自谦,男,硕士研究生。主要研究方向:天文导航及组合导航技术;王海涌,男,博士,讲师,硕士生导师。主要研究方向:天体敏感器、天文导航及组合导航技术。
  • 基金资助:
    航空科学基金(20130151004)

Strapdown inertial/celestial/radar altimeter integrated navigation

GAO Ziqian1, WANG Haiyong1, WANG Yonghai2, ZHAO Zhenping2, CHEN Ken2, QIN Xuguo2   

  1. 1. School of Astronautics, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;
    2. Beijing Institute of Space Long March Vehicle, Beijing 100076, China
  • Received:2016-11-09 Revised:2017-03-06 Online:2017-12-20 Published:2017-04-07
  • Supported by:
    Aeronautical Science Foundation of China (20130151004)

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%。

关键词: 捷联惯性/天文, 雷达高度表, 组合导航, 弹道导弹, 扩展卡尔曼滤波(EKF)

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.

Key words: strapdown inertial/celestial, radar altimeter, integrated navigation, ballistic missile, extend Kalman filter (EKF)

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