초록 ▼

An attitude controller for a 3-axis stabilized, Earth oriented bias momentum spacecraft is described, where only 2-axis attitude measurements from the Earth sensor are available. In contrast to the classical spacecrafts that are controlled w.r.t. a fixed orbital Earth pointing reference frame, (poss

An attitude controller for a 3-axis stabilized, Earth oriented bias momentum spacecraft is described, where only 2-axis attitude measurements from the Earth sensor are available. In contrast to the classical spacecrafts that are controlled w.r.t. a fixed orbital Earth pointing reference frame, (possibly large angle) time varying reference signals are considered here, i.e. the control task consists of a tracking problem. The controller design consists of a decoupling controller and axis-related PID controllers based on yaw observer estimates.

대표청구항 ▼

[ We claim:] [1.] Method for controlling attitude of a three-axis-stabilized, earth oriented bias momentum spacecraft being suppled on its orbit with time varying attitude reference signals with respect to an orbital earth pointing reference coordinate system, and having measurement means for two-ax

[ We claim:] [1.] Method for controlling attitude of a three-axis-stabilized, earth oriented bias momentum spacecraft being suppled on its orbit with time varying attitude reference signals with respect to an orbital earth pointing reference coordinate system, and having measurement means for two-axis attitude measurement and a set of wheels for establishing an angular momentum in any desired direction and for producing torques about three axes of a spacecraft-related orthogonal coordinate system, the method comprising:feeding to a signal processing and control device said time varying attitude reference signals, measurement signals from said measurement means and signals representing an angular momentum vector of said set of wheels;said signal processing and control device generating first, second and third control signals;using the sum of said first, second and third control signals as a torque command signal for said set of wheels;for the generation of said first control signal;starting from an Euler equation describing dynamics of the spacecraft, transforming the Euler equation by introducing measurable parts .omega..sub.M and .omega..sub.M of a spacecraft angular velocity vector .omega. and its derivative .omega., respectively as derived from .PHI..sub.M .tbd.[.phi..THETA.0].sup.T, a measurable part of vector .PHI..tbd.[.phi..THETA..psi.].sup.T representing deviation of the spacecraft from its reference attitude; anddetermining said first control signal as one part of the torque vector figuring in the Euler equation, being related to said set of wheels in such a way that said one part compensates all relevant terms in the Euler equation that are measurable by means of said measurement means except for the term I.PHI., wherein I is the spacecraft inertia matrix;determining a sum of said second and third control signals as a second part of said torque vector;said second control signal being generated by one of a PID and a PD control law, using measurable components .phi., .THETA. of vector .PHI. as control input; andsaid third control signal being generated by one of a D control law using the computed derivative .psi. as controller input and a PD control law using the computed .psi. as input for a time varying observer; and the estimated observer state variables as controller input.