Abstract: An inertial reference system for a spacecraft includes an attitude control assembly and a plurality of force sensors. The attitude control assembly is coupled to a spacecraft body at a plurality of attachment points and at least one of the plurality of force sensors is integrated at each of the attachment points. In one embodiment, the force sensors are piezoelectric transducers. In another embodiment, the control assembly is coupled to the spacecraft body at four attachment points. In yet another embodiment, each of the force sensors is preloaded by a bolt that attaches the attitude control assembly to the spacecraft body.

Abstract: A preload adjustment device and method are provided for momentum control devices. The preload adjustment device includes a piezodynamic preload spacer and a control system. The piezodynamic preload spacer is coupled to a bearing in the momentum control device. The piezodynamic preload spacer is configured such that the application of a control voltage to spacer causes a change in the spacer dimensions, with that change in spacer dimension adjusting the preload of the bearing within the momentum control device. The control system provides dynamic control of preload by selective application of control voltage to the piezodynamic preload spacer. This allows for adjustments of preload to compensate for changes in operating environment, improving the performance of the momentum control device. Additionally, adjustments of preload can by used to compensate for wear in the bearings that would otherwise negatively impact the life of the momentum control device.

Abstract: A vibration damping device and method for momentum control devices is provided. The vibration damping device includes a piezodynamic damping spacer and a tuning system. The piezodynamic damping spacer is coupled to a bearing in the momentum control device. The piezodynamic damping spacer is configured such that vibrations in the momentum control device are absorbed by piezodynamic damping spacer. The piezodynamic damping spacer converts these vibrations to electrical energy, where they can be dissipated by the tuning system. The tuning system provides the ability to tune the vibration damping device to most effectively absorb vibrations in specific frequency ranges. Thus, the vibration damping device is able to effectively reduce vibrations in the momentum control device.

Abstract: The present invention provides a control moment gyroscope that overcomes many of the limitations of the prior art. The control moment gyroscope provides a radial actuator mechanism to provide rotation to an inner gimbal assembly around a gimbal axis. The radial actuator mechanism comprises a circular ring at the midpoint of the gimbal axis, the circular ring extending around the outer periphery of the inner gimbal assembly housing. The radial actuator mechanism rotates the inner gimbal assembly around the gimbal axis through the use of a non-contact motor that provides rotational force directly to the outer perimeter of the inner gimbal assembly. Because the inner gimbal assembly is rotated from the midpoint of the gimbal axis, a torque module assembly at the end of the gimbal axis is not required. Thus, the overall length of the control moment gyroscope can be reduced.

Abstract: A plurality of sensors provide spacecraft attitude signals to reaction wheel assemblies which provide converted signals to a flight controller that controls the reaction wheel assemblies. The reaction wheel assemblies include a secondary power supply to power the sensors, and power can be regenerated from the reaction wheels. A microcomputer in the reaction wheel assemblies can control the reaction wheel assemblies in place of the flight controller.

Abstract: The present invention provides a control moment gyroscope that overcomes many of the limitations of the prior art. The control moment gyroscope provides a radial actuator mechanism to provide rotation to an inner gimbal assembly around a gimbal axis. The radial actuator mechanism comprises a circular ring at the midpoint of the gimbal axis, the circular ring extending around the outer periphery of the inner gimbal assembly housing. The radial actuator mechanism rotates the inner gimbal assembly around the gimbal axis through the use of a non-contact motor that provides rotational force directly to the outer perimeter of the inner gimbal assembly. Because the inner gimbal assembly is rotated from the midpoint of the gimbal axis, a torque module assembly at the end of the gimbal axis is not required. Thus, the overall length of the control moment gyroscope can be reduced.

Abstract: A reaction wheel system is provided that includes at least two rotors. The first rotor is the primary rotor that provides the large output torques to the vehicle. The second rotor is a vernier control rotor. The primary rotor and vernier control rotor each rotate about a common axis. The vernier control rotor has an inertial mass that is less than the inertial mass of the primary rotor, and rotates independently of the primary rotor. Because the vernier control rotor can be rotated independently from the primary rotor, it can be used to significantly improve the performance of the reaction wheel system. Specifically, the vernier control rotor is used to provide relatively small output torques. These relatively small output torques can be used to reduce the disturbances created by motor ripple, provide precise torque output control and/or reduce the disturbances created by static friction.

Abstract: A plurality of sensors provide spacecraft attitude signals to reaction wheel assemblies which provide converted signals to a flight controller that controls the reaction wheel assemblies. The reaction wheel assemblies include a secondary power supply to power the sensors, and power can be regenerated from the reaction wheels. A microcomputer in the reaction wheel assemblies can control the reaction wheel assemblies in place of the flight controller.

Abstract: An inertial reference system for a spacecraft includes an attitude control assembly and a plurality of force sensors. The attitude control assembly is coupled to a spacecraft body at a plurality of attachment points and at least one of the plurality of force sensors is integrated at each of the attachment points. In one embodiment, the force sensors are piezoelectric transducers. In another embodiment, the control assembly is coupled to the spacecraft body at four attachment points. In yet another embodiment, each of the force sensors is preloaded by a bolt that attaches the attitude control assembly to the spacecraft body.

Abstract: Dynamic behavior of structures subject to vibrations is monitored with light coupled into a single multi-mode optical fiber positioned within or on the structure. Vibration caused strain or deflections in the structure are detected by interference light signal intensity variations caused by differential phase changes in the transmission modes. These intensity variations are optically detected to provide an electrical output. Actuators may be positioned at the vibration nodes of the structure and energized by signals derived from the electrical output signals to suppress the vibrations.