Space Systems/loral, LLC

Thomas J Holmes

John Scott Higham

Donald W Gamble

Sun Hur-Diaz

Scott W Tilley

Keith J Reckdahl

Ahmed A Kamel

John A Lehner

David Cielaszyk

David J Wirthman

Peter Timothy Gauthier

David L Cielaszyk

Christopher R Purvis

A spacecraft attitude control system that enables friction and stiction compensation for attitude control of a spacecraft using at least one reaction wheels that can spin through zero rpm without degradation in pointing. The system removes nuisance disturbances relating to friction or stiction that detract from the ability of the system to maintain or change spacecraft attitude. It can remove these effects before they are sensed by the systems attitude control sensors. A friction or stiction compensation buffer is arranged between the spacecraft attitude control subsystem and the torque actuators for the at least one reaction wheel. Since the compensation is localized to each wheel, it is possible to optionally: control wheel operating speeds to safe levels, provide wheel failure flags, provide accurate wheel tracking at slow or zero wheel speeds and provide full access to the wheel torques.

Friction and stiction compensation for spacecraft attitude control

Apparatus and a method is presented for controlling yaw and roll excursions in a spacecraft having on-board components such as magnetic torquers, a roll thruster, momentum wheels, a wheel controller, an earth sensor, and a Digital Integrating Rate Assembly (DIRA). An observer module on the spacecraft receives inputs containing information comprising the unbiased roll error from the earth sensor, yaw momentum measured from the wheel speeds, and commanded yaw and pitch momentum output from the wheel controller, and produces therefrom output signals indicative of yaw estimate, yaw momentum estimate and torque disturbances estimate. A controller module on the spacecraft receives the output signals and combines them with inputs containing information comprising minimum yaw error and roll thrust yaw controller gain and minimum yaw error and roll thrust yaw momentum controller gain, and produces therefrom a signal for commanding the operation of the magnetic torquers to change roll momentum when within the dead band of the signal and, alternatively, for commanding the operation of both the magnetic torquers and the roll thruster to change roll momentum when the threshold of the dead band is exceeded. By so doing, the yaw attitude and yaw excursions are controlled.

Magnetic torquer control with thruster augmentation

An onboard attitude control system is constructed to utilize a four reaction wheel system having a reference axis, wherein at least three of the reaction wheel spin axes are oriented obliquely to the reference axis. Current attitude is estimated based on uploaded orbital data, onboard sensed earth and sun position data, and attitude data sensed by a three axes gyroscope system. Current attitude is compared to mission attitude to calculate an error which is transformed to a trihedral axes adjustment command to actuate the reaction wheel system.

Onboard attitude control using reaction wheels

An automatic apparatus and method that is used to switch actuators, such as momentum wheels, used on a spacecraft without impacting the pointing performance of the spacecraft. The apparatus and method allows an actuator to be automatically deactivated without affecting spacecraft performance such as is caused by a failed actuator. A plurality of local actuator controllers are respectively coupled to a plurality of actuators. A system controller is coupled to the plurality of local controllers and implements a control algorithm corresponding to the present method that determines forces/torques for each of the actuators and outputs force/torque commands to each of the actuators to control the pointing direction of the spacecraft. This is achieved by estimating the drag exerted by or integral control signal produced by each actuator to determine if it is approaching failure. In the event of failure, a backup actuator is activated, the failing actuator is shut down, and the forces/torques exerted by the remaining actuators are adjusted in an amount corresponding to the force/torque exerted by the failing actuator.

Transient free actuator switching

A system and process is described for maintaining the attitude of a space craft to compensate for external disturbances. The on board computer generates a thruster on time signal derived from a comparison of estimated adjusted attitude based substantially on rigid body assumptions which to do not take into consideration the flexing reactions of on board flexible components.

Thruster on time signaling with flexure compensation avoidance

A method for controlling the direction of a spacecraft of the type having a controller for calculating thrust forces comprises a plurality of thrusters disposed on said spacecraft in a spaced relationship relative to one another and aligning and locating each of said plurality of thrusters at a predetermined angle on said spacecraft relative to first, second and third orthogonally oriented axes such that the firing of any given pair of thrusters produces a torque about one of said spacecraft axes; and firing of all of said plurality of said thrusters in equal amounts at one time to cause the spacecraft to move in a linear manner along one of said three orthognally disposed axes.

Thruster optimized pair selection

An onboard star tracking system is used to provide continuous attitude data based on a fixed star catalog. An orbital position reference for the sensed star data is estimated from uploaded data. The attitude is calculated based solely on the star sensor data referenced to the estimated orbital position and compared to mission attitude to generate an attitude adjustment.

Simplified onboard attitude control based on star sensing

An onboard star tracking system is used to provide continuous attitude and orbital position data independent of the image generating instruments. Such data is instantaneously and continually supplied to the onboard computer which translates such information into mirror position coordinates which is added to the mirror position set based on estimated data from the ground control. An error signal is generated which is used to further adjust the position of the mirror.

Satellite imaging control system for non-repeatable error

For a switching of a magnitude of system gain in the driving of a load by a control system, a method of inhibiting formation of switching transients has a step of distinguishing between an estimate of a plant state, such as the attitude of a spacecraft, and an integral control state which allows the control system to provide actuation in response to long term disturbances, such as solar pressure, in operation of the system. In the case of a feedback configuration to the system with controller in a forward branch and an estimator in a feedback branch, a loop error signal serves to drive an actuator of the load via the controller. The method includes a further step of evaluating a portion of a controller input signal which is exclusive of an integral control state and which comprises a difference between a desired state and an estimate of the plant state. The evaluation step enables observation of an occurrence of a magnitude of the difference which falls within a range of values satisfying a criteria of minimal effect on the actuator. There is a further step of switching a value of the gain during the occurrence. Also, concurrently with the switching step, there is a scaling of an amplitude of the integral state by a factor inverse to a ratio of switched gain values.

Transient-free gain switching within error threshold

An active attitude control system for a spacecraft having first, second, and third mutually perpendicular axes utilizes four flywheels, at a minimum, which can selectively provide options of fully redundant momentum bias, 3/4 redundant momentum bias, or fully redundant zero momentum bias. A plurality of reaction wheels are mounted on the spacecraft and rotatable on spin axes in a fixed configuration for together maintaining full three-axis control of the spacecraft to a predetermined attitude. A momentum wheel is also rotatable on the spacecraft about the first axis for maintaining gyroscopic stiffness. In the event of a failure of the momentum wheel, the reaction wheels may have a combined angular momentum sufficient to maintain the gyroscopic stiffness lost by the failure of the momentum wheel while maintaining full three-axis control of the spacecraft to a predetermined attitude. In one embodiment, the reaction wheels are mounted in a fixed trihedral configuration, a first reaction wheel being rotatable about a spin axis which lies in a plane of the first and second axes, and second and third reaction wheels being rotatable about spin axes which lie in a plane of the first axis and are oblique to the second and third axes.

Spacecraft control system with a trihedral momentum bias wheel

An automatic, on-board system for orbiting spacecraft that controls yaw excursions caused by solar torques and thruster firings, which system combines inputs indicative of roll and yaw momentum increases and inputs containing information comprising the unbiased roll error from the earth sensor, yaw momentum measured from the wheel speeds, and commanded yaw momentum output from the wheel controller, and produces therefrom output signals indicative of the yaw estimate and the yaw momentum estimate. These output signals are combined and processed in a controller with a mimimum yaw error and roll thrust yaw controller gain and a miminmum yaw error and roll thrust yaw momentum controller gain, and a signal is produced therefrom for commanding roll thruster firings to change roll momentum, and, in turn, control yaw attitude and yaw excursions.

Thruster control of yaw without yaw measurements

A satellite stabilization system compensates for the destabilizing torque introduced by solar radiation pressure by the emplacement of tabs of material angled relative to the back surfaces of the antenna reflectors to intercept solar rays. The material of the tabs interacts with the solar radiation in substantially the same manner as the interaction of the material on the back surface of each antenna reflector to produce a radiation pressure and consequential increment in a torque about an axis of the satellite. Each of the tabs is constructed in the manner of a plate or disk, and is attached by means of a spring mount to the back surface of one of the reflectors. The reflectors are angled relative to each other, as well as to a body of the satellite which holds feeds for illuminating the reflectors, for directing beams of radiation from the feeds towards the earth. Due to the angulation of the reflectors to each other, the total solar energy incident upon a pair of of the reflectors is generally unequal, this resulting in a net torque from the solar radiation pressure which tends to rotate the satellite. By virtue of the angulation of the tabs relative to the reflectors, an increased amount of solar radiation is received by the reflector which is angled away from the sun's rays, thereby to equalize the torque from both of the reflectors.

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Portola Valley, CA, United States of America

Thomas J Holmes