Attitude control system

Claim 2

An attitude control system for a spacecraft having a momentum bias comprising: a momentum wheel spinning about an axis substantially parallel to a first axis of said spacecraft to produce angular momentum for providing said momentum bias; means for rotating the angular momentum of said momentum wheel with respect to said spacecraft; means for providing an attitude error signal responsive to angular deviation of the spacecraft about one or more axes perpendicular to the first axis; magnetic torquing means for providing, when energized, magnetic dipoles for developing torque about one or more selected axes; first control means responsive to said attitude error signal exceeding a first threshold for energizing said magnetic torquing means for developing correcting torque for reducing said deviation; and second control means coupled to said means for rotating and responsive to said attitude error signal exceeding a second threshold greater than said first threshold, said second control means for causing the angular momentum of said momentum wheel to rotate in a direction to reduce said deviation.

Claim 3

The combination of claim 1 wherein said means for rotating comprises means for pivotally mounting said momentum wheel on said spacecraft.

Claim 4

The combination of claim 2 wherein said means for pivotally mounting includes control logic means for generating pivot signals, and means responsive to said pivot signals for pivoting said momentum wheel.

Claim 5

The combination of claim 3 wherein said means responsive to said pivot signals is a stepper motor.

Claim 6

The combination of claim 3 wherein said pivot signals are generated in pairs separated in time by about one-half the spacecraft nutation period.

Claim 7

The combination of claim 5 wherein said pairs of pivot signals comprise two signals of substantially equal value.

Claim 8

The combination of claim 1 wherein said first axis is the pitch axis and said attitude error signal represents deviation about the roll axis.

Claim 9

The combination of claim 1 wherein said attitude error signal includes a nutation component, and wherein said second control means generates pivot signals for rotating the angular momentum substantially independently of said nutation component.

Claim 10

The combination of claim 8 wherein said second control means generates said pivot signals in pairs of substantially equal value.

Claim 11

The combination of claim 1 wherein said attitude error signal includes a nutation component, and wherein said second control means generates pivot signals for rotating the angular momentum responsive to said attitude error signal including said nutation component.

Claim 12

The combination of claim 10 wherein said second control means generates said pivot signals in pairs of unequal value, said unequal value being responsive to the nutation component of said attitude error signal.

Claim 13

The combination of claim 1 wherein said first threshold is zero.

Claim 14

An attitude control system for a spacecraft comprising: a spinning body pivotally mounted to a platform body wherein the axis of the pivotal mounting is orthogonal to the nominal angular momentum axis of the spacecraft; means responsive to motion of the platform body in inertial space about an axis transverse to the nominal angular momentum axis, for providing a first signal representing a first error angle φ 1 about said transverse axis at a first time, and for providing a second signal representing a second error angle φ 2 about said same transverse axis at a second time one-half spacecraft nutation period later than said first time; and drive means coupled between said spinning body and said platform body and responsive to said first and second signals for pivoting the spinning body over a first pivot angle proportional to (φ 1 +φ 2 ) and one-half nutation period later pivoting the spinning body over a second pivot angle also proportional to (φ 1 +φ 2 ).

Claim 15

An attitude control system for a spacecraft comprising: a spinning body pivotally mounted to a platform body wherein the axis of the pivotal mounting is orthogonal to the nominal angular momentum axis of the spacecraft; means responsive to motion of the platform body in inertial space about an axis transverse to the nominal angular momentum axis, for providing a first signal representing a first error angle φ 1 about said transverse axis at a first time, and for providing a second signal representing a second error angle φ 2 about said same transverse axis at a second time one-half spacecraft nutation period later than said first time; and drive means coupled between said spinning body and said platform body and responsive to said first and second signals for pivoting the spinning body over a first pivot angle proportional to (φ 1 +φ 2 ) and one-half nutation period later pivoting the spinning body over a second pivot angle also proportional to (φ 1 +φ 2 ).

Claim 16

The system of claim 13 wherein said first and second pivot angles are substantially equal.

Claim 17

The system of claim 13 wherein said drive means includes a stepper motor and means for driving said stepper motor, wherein said drive means drives said stepper motor a number of steps proportional to (φ 1 +φ 2 ).

Claim 18

The system of claim 13 wherein said means responsive to motion provides said first signal responsive to said first error angle φ 1 exceeding a threshold value.

Claim 19

An attitude control and nutation damping system for a spacecraft comprising: a spinning body pivotally mounted to a platform body wherein the axis of the pivotal mounting is orthogonal to the nominal angular momentum axis of the spacecraft; means responsive to motion of the platform body in inertial space about an axis transverse to the nominal angular momentum axis, for providing a first signal representing a first error angle φ 1 about said transverse axis at a first time, and for providing a second signal representing a second error angle φ 2 about said same transverse axis at a second time one-half spacecraft nutation period later than said first time; and drive means coupled between said spinning body and said platform body and responsive to said first and second signals for generating first and second control signals proportional to (φ 1 +φ 2 ) and (φ 1 -φ 2 ), respectively, for pivoting the spinning body over a first pivot angle proportional to the sum of said first and second control signals, and one-half nutation period later pivoting the spinning body over a second pivot angle proportional to the difference of said first and second control signals.

Claim 20

An attitude control and nutation damping system for a spacecraft comprising: a spinning body pivotally mounted to a platform body wherein the axis of the pivotal mounting is orthogonal to the nominal angular momentum axis of the spacecraft; means responsive to motion of the platform body in inertial space about an axis transverse to the nominal angular momentum axis, for providing a first signal representing a first error angle φ 1 about said transverse axis at a first time, and for providing a second signal representing a second error angle φ 2 about said same transverse axis at a second time one-half spacecraft nutation period later than said first time; and drive means coupled between said spinning body and said platform body and responsive to said first and second signals for generating first and second control signals proportional to (φ 1 +φ 2 ) and (φ 1 -φ 2 ), respectively, for pivoting the spinning body over a first pivot angle proportional to the sum of said first and second control signals, and one-half nutation period later pivoting the spinning body over a second pivot angle proportional to the difference of said first and second control signals.

Claim 21

The system of claim 17 wherein said drive means includes a stepper motor and means for driving said stepper motor by numbers of steps related to said sum and to said difference of said first and second control signals.

Claim 22

The system of claim 17 wherein said means responsive to motion provides said first signal responsive to said first error angle φ 1 exceeding a threshold value.

Claim 23

The system of claim 17 wherein said means responsive to motion provides said first signal responsive to a peak value of said first error angle φ 1 .

Claim 24

The system of claim 20 wherein said first signal is provided within one-quarter of a spacecraft nutation period after said first error angle φ 1 exceeds a threshold value.