Attitude determination system and method
AssigneeHughes Electronics Corporation
InventorsGarry Didinsky, Yeong-Wei Andy Wu
Life of this patent
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Abstract
An attitude sensing system utilizing simplified techniques and apparatus includes a spacecraft control processor which receives signals from an inertial measurement unit and two attitude sensors. The spacecraft control processor calculates a time-varying gain matrix for estimating attitude errors and gyroscope drifts corresponding to the axes of the inertial measurement unit. In special cases where the separation angle between the attitude sensor vectors is less than approximately 10 degrees, the time-varying gain matrix is computed as a fixed gain matrix and a corresponding desensitizing factor.
Claims7
A method of controlling the attitude of a spacecraft comprising the steps of:
- generating a first attitude measurement vector s.sub.m from a first attitude sensor;
- generating a second attitude measurement vector e.sub.m from a second attitude sensor;
- computing a noise covariance matrix, R, wherein ##EQU7## where .phi. is the separation angle between said first and second attitude measurement vectors and .sigma..sub.e.sup.2 is the first attitude sensor 1 .sigma. noise covariance and .sigma..sub.s.sup.2 is the second attitude sensor I .sigma. noise covariance;
- decomposing said noise covariance matrix, R, such that
- R=UDU.sup.T and
- U U.sup.T =I.sub.3.times.3 ;
- defining a timc-varying gain matrix Kc wherein ##EQU8## wherein q.sub.1 represents gyro angle random walk noise and q.sub.2 represents gyro rate random walk noise;
- and
- updating the attitude of said satellite in a manner directly related to the corresponding value of said Kc.
An attitude sensing system for a spacecraft comprising:
- first and second attitude sensors each of said respective attitude sensors providing an output vector e.sub.m, s.sub.m, to a detected source;
- an inertial measurement unit for providing a direction cosine matrix representing the attitude of said spacecraft in inertial space with respect to a yaw axis, a pitch axis, and a roll axis defining a frame of reference of said system;
- and,
- a spacecraft control processor, responsive to said output vectors and said direction cosine matrix, for providing a time-varying gain matrix for correcting the attitude of said spacecraft wherein said time-varying gain matrix Kc is defined as ##EQU10## wherein q.sub.1 represents gyro angle random walk noise and q.sub.2 represents gyro rate random walk noise.
Dependencies are inferred from the claim text.