A rotational negative-inertia converter (RNIC) has a housing enclosing a flywheel configured to rotate around an axis of symmetry; a motor with a stator attached to the housing and a rotor attached to the flywheel to rotate it around the axis of symmetry; a housing angular accelerometer attached to said housing; a flywheel angular accelerometer; and a controller configured to receive measured accelerometer values from the accelerometers. The controller is configured to drive the motor to maintain the angular acceleration of the flywheel at a value proportional to the housing angular acceleration, with a predetermined proportionality constant.

A method for calibrating an ADCS testbed comprising a DUT holder with three RNICs includes: using measured angular velocities of the DUT holder and RNIC flywheels, and ZGT data, to compute moments of inertia of the DUT holder with and without a satellite with ADCS, allowing compensation for those moments by the RNICs.

An attitude control apparatus for a satellite includes: at least three electric motors, wherein the at least three electric motors are arranged in such a way that a torque may be generated with any orientation of an associated torque vector, and a controller, wherein the controller is configured to drive the at least three electric motors based on a torque controller. The torque controller is adapted to operate the at least three electric motors outside a rest state only when an acceleration torque and a braking torque are required to execute an agile attitude change maneuver. There is also described an associated method.

The invention relates to a flywheel for stabilising the position of a spacecraft, comprising a hub means (1) for fastening the flywheel, a flywheel ring (4), which externally surrounds the hub means (1) circumferentially at a distance, a support means (3) for supporting the flywheel ring (4) on the hub means (1), and a vibration damping device (6, 8) having a tuned mass damper means (8) which is axially movable back and forth relative to the flywheel ring with respect to a rotation axis of the flywheel.

An orbit transition apparatus that transitions an orbit of a payload in outer space includes a rotating body, an adapter disposed on a center part of the rotating body for docking a payload, a launch module disposed outside of the rotating body for launching the payload, and a thruster for rotating the rotating body. The launch module may launch the payload to a target orbit.

A method for attitude control of a satellite in inclined low orbit in survival mode is disclosed, the satellite including at least one solar generator, at least one solar sensor, magnetic torquers capable of forming internal magnetic moments in a satellite reference frame having three orthogonal axes X, Y, and Z, and inertial actuators capable of forming internal angular momentums in the satellite reference frame. The at least one solar sensor has a field of view at least 180° wide within the XZ plane around the Z axis, the method including a step of attitude control using a first control law, a step of searching for the sun by means of the at least one solar sensor, when a first phase of visibility of the sun is detected, and a step of attitude control using a second control law.

An antenna system is configured for use in Low Earth Orbit (LEO) around Earth. The system has a plurality of antenna satellites coupled together to form a phased array. Each of the plurality of antenna satellites have an antenna body with an antenna and a solar cell. A processing device determines an orientation of the plurality of antenna satellites and position the phased array in the orientation based on an analysis of the solar cell of the antenna bodies facing the sun, the antenna of the antenna bodies facing the Earth, and maintaining a torque equilibrium of the phased array.

Apparatus and methods for controlling a spacecraft for a transfer orbit. The spacecraft includes a propulsion subsystem with electric thrusters that are installed with two-axis gimbal assemblies. The spacecraft also includes a controller that identifies a target spin axis for the spacecraft, determines an actual spin axis for the spacecraft during the transfer orbit, determines gimbal angles for the electric thruster(s) that adjust the actual spin axis toward the target spin axis, and initiates a burn of the electric thruster(s) at the gimbal angles.

A method for attitude control of a satellite in inclined low orbit in survival mode is disclosed, the satellite including at least one solar generator, at least one solar sensor, magnetic torquers capable of forming internal magnetic moments in a satellite reference frame having three orthogonal axes X, Y, and Z, and inertial actuators capable of forming internal angular momentums in the satellite reference frame. The at least one solar sensor has a field of view at least 180° wide within the XZ plane around the Z axis, the method including a step of attitude control using a first control law, a step of searching for the sun by means of the at least one solar sensor, when a first phase of visibility of the sun is detected, and a step of attitude control using a second control law.

A rotor assembly for deployment within a momentum control device that enables near-zero revolutions per minute (RPM) sensing, and method for making same, are provided. The provided rotor assembly utilizes a magnet coupled to the rotor shaft and a stationary sensor element to detect magnetic flux from the magnet and derive reliable near zero RPM therefrom.

A hybrid power source and control moment gyroscope (“HPCMG”) is disclosed. The HPCMG includes a control moment gyroscope (“CMG”), a first conductive bearing, and a second conductive bearing. The CMG includes a first transverse gimbal assembly, a central mass that produces a voltage potential, and a second gimbal assembly rotationally connected to the first transverse gimbal assembly. The first transverse gimbal assembly is rotationally connected to the central mass along a first axis of rotation and the central mass is configured to spin about the first axis of rotation and the first transverse gimbal assembly is configured to rotate about a second axis of rotation of the second gimbal assembly. The first conductive bearing rotationally connects the central mass with the first position of the first transverse gimbal assembly along the first axis of rotation.