Abstract: A method and apparatus for generating navigation solutions. A global positioning system based navigation solution unit; an inertial navigation solution unit; a correction unit, a limiter, and an adding unit. The global positioning system based navigation solution unit is capable of generating a first navigation solution. The inertial navigation solution unit is capable of generating a second navigation solution. The correction unit is capable of generating a raw correction. The limiter is capable of selectively modifying the raw correction to fall within a selected range of corrections to form a correction. The adding unit is capable of adding the correction to the second navigation solution to form a navigation solution.

Abstract: A method and apparatus for generating navigation solutions. A global positioning system based navigation solution unit; an inertial navigation solution unit; a correction unit, a limiter, and an adding unit. The global positioning system based navigation solution unit is capable of generating a first navigation solution. The inertial navigation solution unit is capable of generating a second navigation solution. The correction unit is capable of generating a raw correction. The limiter is capable of selectively modifying the raw correction to fall within a selected range of corrections to form a correction. The adding unit is capable of adding the correction to the second navigation solution to form a navigation solution.

Abstract: A method and apparatus for generating navigation solutions. A global positioning system based navigation solution unit; an inertial navigation solution unit; a correction unit, a limiter, and an adding unit. The global positioning system based navigation solution unit is capable of generating a first navigation solution. The inertial navigation solution unit is capable of generating a second navigation solution. The correction unit is capable of generating a raw correction. The limiter is capable of selectively modifying the raw correction to fall within a selected range of corrections to form a correction. The adding unit is capable of adding the correction to the second navigation solution to form a navigation solution.

Abstract: An Ultra-Tightly Coupled GPS-inertial navigation system for use in a moving agile platform includes a range residual extractor that uses best curve fitting of a third order polynomial for estimating range residual. The curve-fitted residual is used to update an error Kalman filter. The error Kalman filter includes correction for navigation solution, and IMU and GPS parameters. The navigation solution together with GPS parameter corrections are used in a Tracking Predictor to generate high-sampling-rate carrier and code replicas. The curve-fitting error covariance indicates signal to noise ratio for the tracked GPS signal and may be used for early indication of interference or jamming.

Abstract: An Ultra-Tightly Coupled GPS-inertial navigation system for use in a moving agile platform includes a range residual extractor that uses best curve fitting of a third order polynomial for estimating range residual. The curve-fitted residual is used to update an error Kalman filter. The error Kalman filter includes correction for navigation solution, and IMU and GPS parameters. The navigation solution together with GPS parameter corrections are used in a Tracking Predictor to generate high-sampling-rate carrier and code replicas. The curve-fitting error covariance indicates signal to noise ratio for the tracked GPS signal and may be used for early indication of interference or jamming.

Abstract: A method for spacecraft power acquisition is provided using single-axis slit sun sensors for both wing-stowed and wing-deployed spacecraft configurations. The method for wing-deployed spacecraft includes initializing a solar wing of the spacecraft to search for sun; rotating the spacecraft about a search axis substantially parallel to a slit sun sensor field of view; monitoring the slit sun sensor for a time of arrival signal; and wherein, if the time of arrival signal occurs, the spacecraft is rotated along the search axis to an orientation where the time of arrival signal occurred and the spacecraft is placed in stable rotation about an axis substantially parallel to a solar wing longitudinal axis; and for a non-occurrence of the time of arrival signal, the spacecraft is slewed about a keyhole axis substantially perpendicular to the search axis to move the sun away from a keyhole.

Abstract: A system and a method wherein future navigation data is broadcast on one channel and current navigation data is broadcast on another channel. By instituting minor changes in message structure, anti-jamming capability can be enhanced. In accordance with one method, future navigation data is distributed by satellites to receivers for tracking of a conventional channel. In accordance with another method, part of the future navigation data (i.e., ephemeris and clock data) is distributed by satellites to receivers for use with a dataless channel.

Abstract: A method and apparatus for generating navigation solutions. A global positioning system based navigation solution unit; an inertial navigation solution unit; a correction unit, a limiter, and an adding unit. The global positioning system based navigation solution unit is capable of generating a first navigation solution. The inertial navigation solution unit is capable of generating a second navigation solution. The correction unit is capable of generating a raw correction. The limiter is capable of selectively modifying the raw correction to fall within a selected range of corrections to form a correction. The adding unit is capable of adding the correction to the second navigation solution to form a navigation solution.

Abstract: An Ultra-Tightly Coupled GPS-inertial navigation system for use in a moving agile platform includes a range residual extractor that uses best curve fitting of a third order polynomial for estimating range residual. The curve-fitted residual is used to update an error Kalman filter. The error Kalman filter includes correction for navigation solution, and IMU and GPS parameters. The navigation solution together with GPS parameter corrections are used in a Tracking Predictor to generate high-sampling-rate carrier and code replicas. The curve-fitting error covariance indicates signal to noise ratio for the tracked GPS signal and may be used for early indication of interference or jamming.

Abstract: A method for spacecraft power acquisition is provided using single-axis slit sun sensors for both wing-stowed and wing-deployed spacecraft configurations. The method for wing-deployed spacecraft includes initializing a solar wing of the spacecraft to search for sun; rotating the spacecraft about a search axis substantially parallel to a slit sun sensor field of view; monitoring the slit sun sensor for a time of arrival signal; and wherein, if the time of arrival signal occurs, the spacecraft is rotated along the search axis to an orientation where the time of arrival signal occurred and the spacecraft is placed in stable rotation about an axis substantially parallel to a solar wing longitudinal axis; and for a non-occurrence of the time of arrival signal, the spacecraft is slewed about a keyhole axis substantially perpendicular to the search axis to move the sun away from a keyhole.

Abstract: A method and apparatus for controlling a plurality of solar panels of a spacecraft is described. The method comprises the steps of providing a first step command to a first solar panel, and providing a second step command to a second solar panel at a time of a transient zero-crossing of a dynamic response of the spacecraft body to the first step command, wherein the second solar panel is disposed on an opposite side of the spacecraft from the first solar panel. The apparatus comprises a processor; a first solar panel driver, communicatively coupled to the processor, for providing a first step command to a first solar panel, and a second solar panel driver, communicatively coupled to the processor, for providing a second step command to a second solar panel at a time of a transient zero-crossing of a dynamic response of the spacecraft body to the first step command.

Abstract: A method and apparatus for estimating a slave payload attitude is disclosed. The method includes accepting a plurality of slave payload attitude measurements, deriving a model of the relative attitude of the slave payload and a master payload attitude at least in part from the plurality of slave attitude measurements, predicting the relative attitude between the slave payload attitude and the master payload attitude using the derived model, and estimating the relative attitude between the slave payload attitude and the master payload attitude at least in part from the predicted relative attitude between the slave payload attitude and the master payload attitude. Furthermore, the absolute attitude of the slave payload is computed using the relative attitude and the master payload attitude.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method uses solar tacking and similar techniques to produce differential solar torques that are used to control the momentum and attitude of the spacecraft during the appendage deployment.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method comprises the steps of predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage, computing a magnitude and a direction of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque, and storing the computed magnitude and direction of momentum in at least one of the momentum wheels before deploying the appendage.

Abstract: A system and method for performing in-orbit alignment calibration using on-board attitude sensors to improve reflector alignment after deployment to improve spacecraft pointing.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method comprises the steps of predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage, computing a magnitude and a direction of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque, and storing the computed magnitude and direction of momentum in at least one of the momentum wheels before deploying the appendage.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method uses solar tacking and similar techniques to produce differential solar torques that are used to control the momentum and attitude of the spacecraft during the appendage deployment.

Abstract: Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance attitude accuracy when compared to systems that realize independent payload estimates, determine payload attitudes. These systems also provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).

Abstract: Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance attitude accuracy when compared to systems that realize independent payload estimates. determine payload attitudes. These systems also provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).