Energy efficient satellite maneuvering is described herein. One disclosed example method includes maneuvering a satellite that is in an orbit around a space body so that a principle sensitive axis of the satellite is oriented to an orbit frame plane to reduce gravity gradient torques acting upon the satellite. The orbit frame plane is based on an orbit frame vector.

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.

Described herein are systems and methods for attitude determination using infrared Earth horizon sensors (EHSs) with Gaussian response characteristics. Attitude information is acquired by detecting Earth's infrared electromagnetic radiation and, subsequently, determining the region obscured by Earth in the sensors' fields of view to compute a nadir vector estimation in the spacecraft's body frame. The method can be applied when two sensors, each with known and distinct pointing directions, detect the horizon, which is defined as having their fields of view partially obscured by Earth. The method can be implemented compactly to provide high-accuracy attitude within small spacecraft, such as CubeSat-based satellites.

A spacecraft including a set of thrusters for changing a pose of the spacecraft. At least two thrusters mounted on a gimbaled boom assembly and are coupled together sharing the same gimbal angle. A model predictive controller (MPC) to produce a solution for controlling thrusters of the spacecraft by optimizing a cost function over a receding horizon using a model of dynamics of the spacecraft effecting a pose of the spacecraft and a model of dynamics of momentum exchange devices of the spacecraft effecting an orientation of the spacecraft. A modulator to modulate magnitudes of the thrust of the coupled thrusters determined by the MPC as pulse signals specifying ON and OFF states of each of the coupled thruster, wherein the ON states of the coupled thrusters sharing the same gimbal angle do not intersect in time. A thruster controller to operate the thrusters according to their corresponding pulse signals.

A control system for controlling an operation of a spacecraft. A model predictive controller (MPC) produces a solution for controlling thrusters of the spacecraft. The MPC optimizes a cost function over a finite receding horizon using a model of dynamics of the spacecraft effecting a pose of the spacecraft and a model of dynamics of momentum exchange devices of the spacecraft effecting an orientation of the spacecraft. The optimization is subject to hard and soft constraints on angles of thrusts generated by thrusters. Further, the hard constraints require the angles of thrusts in the solution to fall within a predetermined range defined by the hard constraints. The soft constraints penalize the solution for deviation of the angles of thrusts from nominal angles corresponding to a torque-free thrust passing through the center of the mass of the spacecraft. A thruster controller operates the thrusters according to the solution of the MPC.

An operation of a spacecraft is controlled using an inner-loop control determining first control inputs for momentum exchange devices to control an orientation of the spacecraft and an outer-loop control determining second control inputs for thrusters of the spacecraft to concurrently control a pose of the spacecraft and a momentum stored by the momentum exchange devices of the spacecraft. The outer-loop control determines the second control inputs using a model of dynamics of the spacecraft including dynamics of the inner-loop control, such that the outer-loop control accounts for effects of actuation of the momentum exchange devices according to the first control inputs determined by the inner-loop control. The thrusters and the momentum exchange devices are controlled according to at least a portion of the first and the second control inputs.

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 propulsion system for small artificial satellites comprises a plurality of engines (2) fixable to a frame (101) of a satellite (100); a control unit (3) connected functionally to the engines (2) for sending at least one activation signal (AS) for activating at least one engine (2); the system is selectively configurable at least between a first configuration in which at least one of the engines (2) is activated for correcting the orbit of the satellite (100) and a second configuration in which at least one of the engines (2) is activated for dispersing said satellite (100) relative to another adjacent satellite.

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.

The invention relates to a flywheel device for position stabilization of a spacecraft, comprising a carrier (1), a rotor (2), a magnetic drive (4) for the rotatingly driving the rotor (2) relative to the carrier (1), and a roller bearing (3) arranged between the rotor (2) and the carrier (1). A magnetic force can be generated between the rotor (2) and the carrier (1) by means of the magnetic drive (4) in order to pre-stress the rolling bearing (3). The outer diameter (A) of the rotor (2) can have, for example, only a maximum of 2.5 times of the rolling bearing diameter (W).