Smooth and Real Time Spacecraft Attitude Estimation by Conciliation of the Kalman Filter Propagation and Updating Steps
- Paper number
IAC-08.C1.7.7
- Author
Dr. Roberto Lopes, Instituto Nacional de Pesquisas Espaciais (INPE), Brazil
- Coauthor
Mr. Adenilson R. da Silva, Brazil
- Year
2008
- Abstract
This paper presents a technique to mitigate disturbances on a spacecraft attitude dynamics due to the use of the Kalman filter. Kalman filter is a well-known unbiased, minimum variance Bayesian estimator of Gauss-Markov stochastic processes from sequential observations. It is suitable for real time applications and particularly useful for spacecraft attitude estimation in a feedback closed loop process for attitude control. Nevertheless, discrepancies between the dynamic and the observation error models cause discontinuities in the filter state estimate at every observation sampling time. Besides being unnatural, piecewise continuous attitude and angular velocity estimates yield a piecewise continuous feedback control. This may not be suitable especially for missions with stringent pointing stability requirements despite the high quality of their attitude control actuators. One proposes a procedure to conciliate the propagation and the updating steps of the Kalman filter at the expenses of a marginal loose in the estimate accuracy. The procedure is based on an optimization approach. A pseudo-control is added to the dynamic model that smoothly drives the state from the propagated estimate at one sampling time to the next. The pseudo-control minimizes a quadratic cost function that takes into account weighted deviations from the Bayesian estimate, the weighted pseudo-control strength, the noise covariance matrices and the control pointing requirements. The resulting state estimate is smooth. Real time feature is preserved. The technique itself may be easily adapted to any application where piecewise continuity in the state estimate represents a drawback. The potential advantage of the procedure to the pointing stability performance is discussed to the particular case of INPE’s Multi Mission Platform in an Earth observation mission using Synthetic Aperture Radar. A proportional-derivative control law is considered to the satellite three-axis attitude control in the nominal mode, based on star sensors and reaction wheels only. In order to emphasize the effect of discontinuities in the state estimate, gyro measurements are intentionally not taken into account and the control actuation is updated at a higher rate than the star sensor sampling rate. Numerical results of the control performance from digital simulation under a typical environment disturbance level are presented and compared with those obtained with the same control law when the conciliation technique is not applied. Based on the results, one concludes that the procedure is a promising original contribution to mitigate attitude control disturbances and may represent an improvement to the spacecraft pointing stability level even in absence of gyros.
- Abstract document
- Manuscript document
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