Abstract: Provided herein are various enhanced systems, apparatuses, and techniques for optical communication among a constellation of satellites. One example includes establishing an initial cross-connect configuration comprising optical communication links among a constellation of satellites. In the initial cross-connect configuration, the satellites can optically communicate with both adjacently positioned and non-adjacently positioned satellites in the constellation. The satellites obtain angular states of the optical communication links resultant from orbital motion of the satellites. When angular states of a portion of the optical communication links reach an exclusion range, the satellites establish a subsequent cross-connect configuration to avoid optically communicating within the exclusion range.

Abstract: The present invention addresses performance limitation of gyros (e.g., MEMS gyros) by significantly reducing common-mode noise and bias effects. In one embodiment, an array of gyros, which may comprise four or more gyros, is configured so that common-mode error effects can be separated from the sensed rotational motion of the gyros and therefore removed. Removing the common-mode effects increases attitude estimation and spacecraft pointing accuracy, particularly during periods when the gyros must solely provide the attitude reference.

Abstract: The present disclosure provides systems and methods that improve the pointing accuracy of a spacecraft using temperature-sensitive gyros (e.g., MEMS gyros) by using a temperature bias model to compensate for temperature biases of the gyros and using attitude data (e.g., star tracker data) to automatically and continuously calibrate the temperature bias model over the life of the spacecraft. When star tracker data is unavailable (e.g., due to sun interference), the most recently updated temperature bias model is used in open-loop to provide improved estimation of the gyro biases and improved attitude estimation.

Abstract: A system for isolating vibration among a plurality of instruments on a spacecraft includes at least two platforms, each of which is configured to couple to and isolate vibration for a single one of the plurality of instruments. Each of the at least two platforms is configured to mount to the spacecraft. A device is also provided for isolating vibration for one instrument among a plurality of instruments on a spacecraft.

Abstract: A spacecraft includes a plurality of thrusters, a thruster firing logic, an actuator assembly, an attitude control system, and a torque calibration system. The plurality of thrusters is configured to apply torque to the spacecraft. The thruster firing logic is configured to control firing of the plurality of thrusters. The actuator assembly is configured to apply torque to the spacecraft. The attitude control system is configured to sense an attitude of the spacecraft and to provide an attitude control torque. The torque calibration system is configured to provide a thruster feedforward torque. A machine-readable medium includes instructions for a method for providing torque calibration to a spacecraft.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: A spacecraft system is provided, including a measurement device configured to measure information associated with position of a spacecraft, a filter configured to process the measured information and to provide estimated pre-maneuver information associated with position of the spacecraft and estimated post-maneuver information associated with position of the spacecraft, a propagator configured to predict post-maneuver information associated with position of the spacecraft based upon the estimated pre-maneuver information associated with position of the spacecraft and a model of a maneuver of the spacecraft, an error calculator configured to calculate an acceleration error based upon the estimated post-maneuver information associated with position of the spacecraft and the predicted post-maneuver information associated with position of the spacecraft, and a maneuver archive configured to store the acceleration error for updating the model of the maneuver.

Abstract: An attitude determination system is provided for determining attitude values of a yaw-steering spacecraft, and includes a rate sensor, at least one attitude sensor, and a processor operable to receive measured spacecraft body rates from the rate sensor, to receive measured spacecraft attitude values from the at least one attitude sensor, and to calculate estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error states, a plurality of gyro bias states, and a plurality of commanded rate dependent gyro error states.

Abstract: The present invention provides a system and a method for improving spacecraft antenna pointing accuracy utilizing feedforward estimation. The present invention takes advantage of the fact that spacecraft antenna pointing error has periodic behavior with a period of 24 hours. Thus, unlike the prior art feedback systems which blindly correct antenna pointing error continuously reacting only to presently sensed error, the present invention takes an intelligent approach and learns the periodic behavior of spacecraft antenna pointing error. Then, an estimate of antenna pointing error at a particular time going forward is predicted based on the learned model of the periodic behavior of the antenna pointing error. The predicted estimate is then used to correct or cancel out the antenna pointing error at a particular time in the future. The result is more accurate correction of spacecraft antenna pointing error by more than a factor of two.

Abstract: Various aspects of the disclosure provide hybrid spacecraft capable of hosting both a remote sensing payload and a communications payload. A hybrid spacecraft includes an isolation system for a remote sensing payload that isolates the remote sensing payload from spacecraft dynamics and disturbances, and prevents interaction between the remote sensing and communications payloads. The hybrid spacecraft may also include provisions for a communications payload including metering structure for large aperture antennas, and scalable resources for accommodating a range of remote and communications payloads.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: Momentum control is maintained in a geosynchronous orbiting spacecraft that uses a plurality of reaction wheel assemblies and a plurality of magnetic torquers to control the spacecraft momentum, each orbit of the spacecraft being comprised of a set of time steps, by determining a current momentum error for a current time step of a current orbit by adding a system momentum change determined for an immediately preceding orbit to an average system momentum determined for the immediately preceding orbit, and then subtracting a magnetic control torque momentum change determined for the immediately preceding orbit, determining a current duty cycle for each of the magnetic torquers based on the current momentum error and on a torque value applied by each magnetic torquer at each time step of the immediately preceding orbit, and commanding each magnetic torquer to operate at the current time step in accordance with its respective determined current duty cycle, wherein the magnetic torquers apply a magnetic momentum c

Abstract: A method for correcting spacecraft pointing errors. An orbit frequency distortion angle between a rate sensor and at least one attitude sensor is estimated. An estimated attitude of the spacecraft is adjusted to correct for the distortion angle.