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Attitude Control of a Precise-Oriented Flexible Satellite by Using Constant Amplitude Thrusters

Attitude maneuver of a flexible satellite is studied in this paper. The satellite consists of a rigid main body and two flexible solar panels. To derive the satellite's equations of motion Lagrange's formulation is employed. Bryant angles of rotation is used to describe the satellite orientation relative to earth. Finite Element Method is then used to discret the motion of the solar panels. For this purpose the solar panels are divided into a number of rectangular plate elements, and only out-of-plane deformations of solar panels are considered.
When the satellite maneuvers are equipped with on-off thrusters, the resulted torques cannot vary. The torques are in constant amplitudes. Under such torques the flexible satellite experiences large vibrations of solar panels after maneuver processes. This vibration then induces attitude angle oscillations of the main rigid body of the satellite.
To achieve a precise orientation of the satellite after attitude maneuver, the torque inputs resulted by thruster must be shaped. Input shaping process must consider several constraints: desired angle displacements, angle velocities, limitations of structural vibration and attitude angle oscillation, fuel consumption, and time limitation.
Numerical simulation then be performed for attitude maneuver of the satellite under suitable shaped inputs. Its computer program is written in FORTRAN language. To solve the differential equations of satellite dynamics, Newmark method is selected.
As the result, a very precise orientation of the flexible satellite can still be achieved after maneuver if the inputs considering the system flexibility are employed for the maneuver.
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