Abstract: In order to gain high accuracy navigation parameters, the attitude transformation matrix between attitude angular error and mathematic platform angular error is derived for a ballistic missile in this paper and is deduced in forms of both quaternion and Euler angles. The method of “State of Space” is often used to estimate navigation parameters, in which the mathematic platform angular error is the observation value involved in the observation equation of “SINS/CNS” integrated navigation system for ballistic missile and the attitude angular error rather than the mathematic platform angular error is obtained. So, the transform matrix with both quaternion form and Euler form used to transform attitude angular error into the mathematic platform angular error is deduced in this paper. The mathematical model in launch inertial coordinate system is established, EKF and UKF are adopted to verify the transformation matrix in the simulation. The results show correctness of the transformation matrix.

Key words: Attitude error, Platform anGles error, EKF, UKF, “SINS/CNS” inteGrated Guidance system