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

doi:10.3788/OPE.20101811.2473

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

Article | 更新时间：2020-08-12

- Design of INS/CNS integrated navigation system for launch vehicle upper stage
- Optics and Precision Engineering Vol. 18, Issue 11, Pages: 2473-2481(2010)
- 作者机构：
  
  1. 中国科学院长春光学精密机械与物理研究所,吉林长春,中国,130033
  2. 哈尔滨工业大学航天学院,黑龙江哈尔滨,150001
- 作者简介：
- 基金信息：
- DOI：10.3788/OPE.20101811.2473
  CLC： V249.32
- Received：16 March 2010，
  
  Revised：13 April 2010，
  
  Published Online：25 November 2010，
  
  Published：25 November 2010
- 稿件说明：
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张利宾, 崔乃刚, 吕世良, 浦甲伦. 运载火箭上面级惯性与天文组合导航系统设计[J]. 光学精密工程, 2010,18(11): 2473-2481

ZHANG Li-bin, CUI Nai-gang, Lü Shi-liang, PU Jia-lun. Design of INS/CNS integrated navigation system for launch vehicle upper stage[J]. Editorial Office of Optics and Precision Engineering, 2010,18(11): 2473-2481
张利宾, 崔乃刚, 吕世良, 浦甲伦. 运载火箭上面级惯性与天文组合导航系统设计[J]. 光学精密工程, 2010,18(11): 2473-2481 DOI： 10.3788/OPE.20101811.2473.

ZHANG Li-bin, CUI Nai-gang, Lü Shi-liang, PU Jia-lun. Design of INS/CNS integrated navigation system for launch vehicle upper stage[J]. Editorial Office of Optics and Precision Engineering, 2010,18(11): 2473-2481 DOI： 10.3788/OPE.20101811.2473.

摘要

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1.171 910

m减小到1.036 710

速度误差矢量和从11.282 7 m/s减小到3.662 6 m/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。

Abstract

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 1.171 910

m to 1.036 710

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.

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references

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