Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
1997-10-14
2001-08-28
Zanelli, Michael J. (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C244S164000
Reexamination Certificate
active
06282467
ABSTRACT:
TECHNICAL FIELD
The present invention relates in general to attitude control of a spacecraft and in particular to using spacecraft momentum to acquire or update the attitude or a spinning updraft.
BACKGROUND ART
Most spacecraft require attitude information relative to a reference frame in order to perform motion control or for mission operations. For earth satellites, an Earth Centered Inertial (ECI) frame is often preferred because it simplifies references for operations personnel.
Typically a spinning spacecraft, such as a satellite, may make a series of control actions to modify attitude and orbit. Preceding each control action is an attitude determination phase. Accurate attitude determination is critical to mininizing the number and extent of each control action. Preceding each attitude determination phase may be a nutation dampening phase to increase the accuracy of the attitude determination.
Many systems and methods are available for determining the attitude of spinning spacecraft. Such a method is described in U.S. Pat. No. 5,020,744 issued to Schwarzschild. The initial spin axis attitude is established by a ground station. An additional sensor, such as a sun sensor, is used to determine the attitude within the plane of rotation. Angular rate information, obtained from gyroscopes, can than be used to propagate positional information. Because gyroscopes drift, frequent recalculation of the attitude by the ground station is required.
Several difficulties exist with current methods for determining the attitude of a spinning spacecraft. A first difficulty is hat current methods allow continuous on-board attitude updates from a sun sensor only once the attitude is determined with respect to a sun normal reference frame. If the attitude is determined with respect to an ECI reference frame, the attitude must be computed on the ground and transmitted to the spacecraft with time critical commanding. Further, once the attitude in the ECI frame is uploaded to the spacecraft, the attitude can only be propagated by the data from gyroscopes. Output from other sensors, such as sun sensors, and the spin axis data cannot be used on-board by tho spacecraft to update the attitude.
SUMMARY OF THE INVENTION
As such, one object of the present invention is to provide a system and method for determining the attitude of a spinning spacecraft in any inertial frame.
Another object of the present invention is to provide a system and method for reducing or eliminating the amount of ground processing necessary to determine the attitude of a spinning spacecraft.
In carrying out the above objects and other objects and feature of the present invention, a method is provided for obtaining information about the spacecraft momentum in an inertial frame, obtaining information about the spacecraft momentum in the body frame, updating the spacecraft attitude using the inertial frame momentum data, body frame momentum data, input from at least one additional sensor, and reference information on the at least one additional sensor, and propagating the attitude using angular rate sensors when the inertial frame momentum is not observable.
In the preferred embodiment, the rotational momentum vector direction in an inertial frame is determined on the ground using measurements taken on the spacecraft and is transmitted to the spacecraft. The rotational momentum vector direction in the body frame is also obtained on the ground using known or estimated mass properties and is transmitted to the spacecraft. The momentum vector directions, together with data from an at least single axis sensor, such as a sun slit sensor and a reference, such as a sun ephemeris model in the ECI frame, are used to update the spacecraft attitude. An at least single-axis sensor, such as a sun slit sensor, together with rate information from gyroscopes are used to propagate the spacecraft attitude.
In another embodiment, the body frame momentum vector is determined on board the spacecraft by summing the rotational moments from each distinct rotating component. Each individual rotational moment is obtained by multiplying the rotational velocity, with the moment of inertia. Both the rotational velocity and the rotational momentum may be determined from sensors or can be estimated from commanded movements.
A further embodiment calculates inertial frame rotational momentum vector direction on the spacecraft. This is accomplished using additional sensors, such as a pair of horizon crossing indicators, to locate a second positional reference. An ephemeris model is then used to compute the spin axis relative to the inertial frame.
A system is also provided in accordance with the present invention for determining the attitude of a spinning spacecraft. The system includes an inertial measurement device, one or more sensors for determining position about the spin axis, ephemeris model references if some or all inertial frame momentum computation will be done on-board, a communication system if some or all momentum computation will be done on Earth, and attitude determination logic for updating and propagating the spacecraft attitude.
The above objects and other objects, features, and advantages of the present invention are readily apparent from the following detailed description of the best mode for carrying out the invention when taken in connection with the accompanying drawings.
REFERENCES:
patent: 3591108 (1971-07-01), Perkel et al.
patent: 3637170 (1972-01-01), Paine et al.
patent: 4617634 (1986-10-01), Izumida et al.
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patent: 5452869 (1995-09-01), Basuthakur et al.
patent: 5702067 (1997-12-01), Bruederle
Didinsky Garry
Shah Piyush R.
Uetrecht David S.
Gibson Eric M
Gudmestad Terje
The Boeing Company
Zanelli Michael J.
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