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.

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.

A reverse directional thrust device by a bidirectional translative movement is provided. The reverse directional thrust device is a thruster, formed by a drive motor, a magnetic reversing mechanism, and a system within a hermetic frame. The system allows a thrust to forward a sense of direction and to reverse the thrust towards an opposite sense of direction, and starting from an expulsion of compressed air, from an interaction with a magnetic field allowing the bidirectional translative movements of components rotary, wherein two translating impulse discs perforated with curved holes and fixed to a threaded axis of rotation, and a toothed disc interposed between the two translating impulse discs. The drive motor and the magnetic reversing mechanism do not fall within a scope of the reverse directional thrust device.

A magnetically suspended control moment gyroscope comprising: a gimbal; a flywheel system, set in the gimbal; wherein the flywheel system comprises: a housing; a shaft, arranged in an inner cavity of the housing; a radial magnetic bearing, comprising: a first rotor portion and a first stator portion fixed to the shaft; an upper axial magnetic bearing and a lower axial magnetic bearing, wherein the upper axial magnetic bearing is fixed to an upper end of the first stator portion, the lower axial magnetic bearing is fixed to a lower end of the first stator portion; a wheel body, set in the radial magnetic bearing, fixed to the first rotor portion; an upper axial thrust plate and a lower axial thrust plate, wherein the upper axial thrust plate is fixed to an upper end of the wheel body, and is on an upper end of the upper axial magnetic bearing, the lower axial thrust plate is fixed to a lower end of the wheel body, and is under a lower end of the lower axial magnetic bearing.

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.

Cell-based systems may interlock in a reconfigurable configuration to support a mission. Space systems, for example, of a relatively large size may be assembled using an ensemble of individual “cells”, which are individual space vehicles. The cells may be held together via magnets, electromagnets, mechanical interlocks, etc. The topology or shape of the joined cells may be altered by cells hopping, rotating, or “rolling” along the joint ensemble. The cells may be multifunctional, mass producible units. Rotation of cell faces, or of components within cells, may change the functionality of the cell. The cell maybe collapsible for stowage or during launch.

A control moment gyroscope (CMG) is provided, selectively having a first spatial configuration and a second spatial configuration at least during operation of the CMG. In the first spatial configuration the CMG occupies a smaller volume than in the second spatial configuration. For example, in the first spatial configuration no part of the CMG projects beyond a predetermined geometrical boundary, while in the second spatial configuration, a portion of the CMG projects beyond the geometrical boundary.

A stackable pancake satellite that is configured so that a plurality of the satellites can be stacked within a payload fairing of a launch vehicle. Each satellite includes sections that are folded or rotated together prior to launch, and unfolded or rotated away from each other when deployed. A first section is a satellite body having a first side that acts as a thermal radiator and a second side opposite the first side that includes an antenna. A second section includes one or more solar panels attached adjacent to the first side of the satellite body. A third section includes a splash plate reflector attached adjacent to the second side of the satellite body that reflects signals between Earth and the antenna. When deployed, the solar panels are pointed towards the Sun and the splash plate reflector directs the signals between the Earth and the antenna.