The present disclosure relates to an axial flux motor for a reaction wheel, and method of using and making the same. The motor includes a stator and a rotor. The stator comprises a printed circuit board (PCB) including a first motor coil. The rotor is coupled to a first ring-shaped magnet having an alternating pole arrangement. In a further embodiment, the rotor includes permanent magnets, and the stator PCB includes a first motor coil, and a first high thermal conductivity element.

A control system configured to execute a safing mode sequence for a spacecraft is disclosed. The control system includes one or more star trackers that each include a field of view to capture light from a plurality of space objects surrounding the celestial body. The control system also includes one or more actuators, one or more processors in electronic communication with the one or more actuators, and a memory coupled to the one or more processors. The memory stores data into a database and program code that, when executed by the one or more processors, causes the control system to determine a current attitude of the spacecraft, and re-orient the spacecraft from a current attitude into a momentum neutral attitude.

A spacecraft attitude control module (1) according to the invention is compact and easy to assemble with additional modules to form an operative attitude control system. The module comprises a robust rectangular, preferably cubic support frame with an attitude control assembly fitted within the confines of the support frame, the assembly including a reference structure comprising a platform, a flywheel support structure (15, 18, 19, 26) and a flywheel (25). The flywheel support structure may be fixed to the platform (10) or it may be a gimbal structure that is rotatable relative to the platform. In the first case the module is a reaction wheel module. In the second case the module is a single gimbal control moment gyroscope module. A preferred embodiment includes a slanted position of the platform (10) relative to the ground plane (100) of the support frame. Another preferred characteristic is the implementation of a flywheel provided with a hollow portion (25′) into which the motor (28) that is driving the flywheel rotation is fitted. The invention is also related to an attitude control system comprising multiple modules assembled together on a support plate (35). The modules may be provided with decking plates (39, 39′) to improve the mechanical robustness of the assembly and to realize fast electrical connections to the modules.

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.

A launch lock system includes a first portion rigidly coupled to the second portion in a first state and the first portion movable in all directions relative to the second portion in a second state. The launch lock system includes a fastener subassembly coupled to the second portion, and the fastener subassembly is movable relative to the second portion from a first position to a second position. The launch lock system includes at least one pivot arm subassembly having a pivot arm movable between a first position and a second position. The pivot arm is coupled to the first portion in the first position. In the first state, the pivot arm is in the first position and cooperates with the fastener subassembly in the first position, and in the second state, the pivot arm is uncoupled from the first portion and the fastener subassembly.

An attitude control system includes one or more control moment gyro pairs, with gyros of individual of the pairs being counter-rotated to rotate the rotation axes of flywheels of the gyros of a gyro pair in opposite direction. The flywheels of a gyro pair may be in paddle configuration, with the rotation axes of the flywheels rotating in the counter-rotation through separate planes as the gyros are rotated. The rotation of the gyros of a gyro pair may be accomplished by coupling both of the gyros to a servo motor with suitable coupling gears, or by using independent servos for each gyro. The counter-rotation of gyros of an individual pair produces a resultant torque about a fixed global axis, such as the axis of a flight vehicle of which the attitude control system is a part. Further control may be accomplished for example by varying rotation speeds of the flywheels.

When calculating a gimbal angle trajectory that satisfies boundary conditions set by an attitude boundary condition setter 2131 of the ground station 21, a gimbal angle trajectory calculator 2132 calculates the gimbal angle trajectory that minimizes a period of an acceleration interval within a range that satisfies driving restrictions of a gimbal, based on a gimbal angle θ.sub.0i of a start time and a gimbal angle θ.sub.ci of a fixed interval of an attitude change. Also, the gimbal angle trajectory is calculated that minimizes a period of a deceleration interval within a range that satisfies the driving restrictions of the gimbal, based on the gimbal angle θ.sub.ci of the fixed interval and a gimbal angle θ.sub.fi of a completion time of the attitude change. The obtained θ.sub.0i, θ.sub.ci, θ.sub.fi and an attitude change period τ are transmitted to the artificial satellite as gimbal angle trajectory parameters, and the control moment gyros are controlled based on the gimbal angle trajectory parameters.

An attitude control apparatus (20) includes an ideal thrust direction calculator (22), an ideal attitude calculator (24), a target attitude calculator (26), and a torque calculator (28). The ideal thrust direction calculator (22) calculates an ideal thrust direction of a thruster. The target attitude calculator (26) calculates a target attitude that is the attitude of a satellite in which a deviation from an ideal attitude is minimized within a movement limitation of an attitude control actuator (14) while a panel surface faces the sun. The torque calculator (28) calculates a torque for turning the satellite from an actual attitude to the target attitude and transmits a torque instruction to the attitude control actuator (14).

This patent presents an attitude determination and control system based on a Quaternion Kalman Filter (QKF) with an extendable number of sensors and actuators. Furthermore, it is compatible with the spherical motor as its attitude actuator. The system includes a processor with a QKF, at least one direct attitude actuator, and at least two environmental sensors. Firstly, system dynamics calculates a first propagation attitude determination result. Next, update the first propagation with the attitude sensor measurements. Then, control the satellite's attitude via the attitude actuator closer to the attitude command provided by the user. The proposed system dynamic model could adjust the number of actuators and sensors freely without reprogramming the algorithms for new missions with new configurations on the actuators and sensors. Moreover, if some components fail, the algorithm can automatically remove those related sequences to avoid the overall failure of the system.

This patent presents an attitude determination and control system based on a Quaternion Kalman Filter (QKF) with an extendable number of sensors and actuators. Furthermore, it is compatible with the spherical motor as its attitude actuator. The system includes a processor with a QKF, at least one direct attitude actuator, and at least two environmental sensors. Firstly, system dynamics calculates a first propagation attitude determination result. Next, update the first propagation with the attitude sensor measurements. Then, control the satellite's attitude via the attitude actuator closer to the attitude command provided by the user. The proposed system dynamic model could adjust the number of actuators and sensors freely without reprogramming the algorithms for new missions with new configurations on the actuators and sensors. Moreover, if some components fail, the algorithm can automatically remove those related sequences to avoid the overall failure of the system.