Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
1998-12-16
2001-10-02
Cuchlinski, Jr., William A. (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C244S164000, C244S165000, C244S171000, C250S330000, C701S010000, C340S973000, C340S974000
Reexamination Certificate
active
06298288
ABSTRACT:
BACKGROUND OF THE INVENTION
(a) Field of the Invention
The present invention relates generally to spacecraft or satellite attitude determination systems and, more particularly, to a method and apparatus for correcting gyro scale factor and misalignment errors to improve attitude determination performance.
(b) Description of Related Art
The term attitude is used to describe the orientation of an object with respect to a reference orientation. Attitude is of particular interest in satellite or spacecraft operations. For example, if a satellite is used in a communications application, it is necessary that the satellite be oriented in the proper direction to receive and/or transmit relevant information for the communication link.
The attitude of a satellite is determined by computations based on the output of sensors located on the satellite. Gyros and object trackers (such as star trackers, sun sensors, and earth sensors) are two types of sensors that may be used in attitude determination systems. In general, gyros are used to measure the rate at which the spacecraft is moving. By integrating gyro output, spacecraft attitude may be determined. The use of gyros in conjunction with star trackers is commonly known in the art as a stellar-inertial attitude determination system.
Object trackers such as star trackers, earth sensors, or sun sensors are used to determine the orientation or attitude of the satellite with respect to the objects being tracked. Object trackers commonly use a CCD array to measure heat or light emitted from the tracked object. Typically, satellites use an ephemeris system to determine the location of the objects being tracked with respect to the earth. An ephemeris system uses a table containing the coordinates of a celestial body at a number of specific times during a given period. Using ephemeris techniques and object trackers, spacecraft attitude with respect to the earth can be determined.
The use of gyros in an attitude determination system results in attitude errors induced by gyro scale factor and misalignment errors (GSFME). These errors appear as gyro bias errors for non-dynamic spacecraft missions such as an orbit-normal steered spacecraft orbiting in a circular orbit. In non-dynamic missions, the angular accelerations of the spacecraft about its axes are relatively small. Thus, bias errors in non-dynamic missions can easily be estimated and compensated for in the spacecraft's attitude determination system by using a standard six state stellar/inertial Kalman filter, which is well known in the art. Therefore, GSFME has not typically been of much concern because most spacecraft missions were historically not dynamic.
However, spacecraft use has expanded to dynamic missions such as sunnadir steering, perigee passing for highly elliptical orbit (HEO), and dynamic target tracking. In dynamic missions, the angular accelerations of the spacecraft about its axes are relatively large. Thus, GSFME in dynamic spacecraft missions can contribute significantly to spacecraft attitude error. Moreover, in dynamic missions, GSFMEs do not manifest themselves as bias and, therefore, cannot be easily eliminated using a six-state Kalman filter.
Accordingly, there is a need for a method and apparatus that eliminates gyro misalignment and bias errors in dynamic spacecraft missions.
SUMMARY OF THE INVENTION
The present invention provides a method and apparatus for eliminating gyro misalignment and scale factor errors in dynamic spacecraft missions. The invention may be embodied in an attitude control/measurement system for estimating the attitude of a spacecraft. The system has three inertial sensors and a stellar position sensor all coupled to a digital filter. The digital filter receives estimated attitude information from the stellar position sensor and compares it with attitude information derived from the inertial sensors. The filter produces corrective data signals representing misalignment errors between the inertial sensors and/or the scale factor errors of each inertial sensor. The system corrects the inertially derived attitude estimate using the corrective data signals.
In accordance with another aspect of the present invention, a method of estimating the attitude of a spacecraft includes the steps of receiving inertial information from three inertial sensors, correcting the inertial information for scale factor and/or misalignment errors using a first time varying gain matrix, propagating a three axis attitude using the corrected inertial information, and correcting the propagated attitude estimate for attitude errors using a second time varying gain matrix. The first and second time varying matrices are supplied by a digital filter.
The invention itself, together with further objects and attendant advantages, will best be understood by reference to the following detailed description, taken in conjunction with the accompanying drawings.
REFERENCES:
patent: 5109346 (1992-04-01), Wertz
patent: 5259577 (1993-11-01), Achkar et al.
patent: 5348255 (1994-09-01), Abreu
patent: 5452869 (1995-09-01), Basuthakur et al.
patent: 6047226 (2000-04-01), Wu et al.
Didinsky Garry
Li Rongsheng
Wu Yeong-Wei A.
Cuchlinski Jr. William A.
Gudmestad Terje
Hughes Electronics Corp.
To Tuan C
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