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.

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 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.

A spacecraft for changing an orbit or an attitude of a target in outer space by irradiating the target with a laser, the spacecraft includes: a laser apparatus configured to generate the laser; a focusing unit configured to converge the laser; a detecting unit configured to acquire detection information including a distance between the spacecraft and the target; and an irradiation control unit configured to control the focusing unit on the basis of the distance so that the laser converges on the target.

In an example embodiment, an attitude, orbit, and de-orbit control system (AODCS) for a satellite is provided. In an example embodiment, the AODCS system comprises one or more selectively retractable booms. The one or more selectively retractable booms are collectively configured to provide a selectively adjustable drag during de-orbiting of a satellite over a predefined de-orbiting time.

The invention relates to small satellites capable to fly in formation (10), in particular nano- or picosatellites with a mass of 10 kg or less, for LEO applications, comprising a housing (12) and at least one plug-in board (14) arranged in the housing (12) with a predetermined functionality and a propulsion system (16) for generating a directed pulse in the direction of the flight trajectory T.sub.k. It is proposed that the small satellite (10) comprises an independent and autonomously working collision avoidance system (18), which is capable of adapting a trajectory correction T.sub.kk of the trajectory T.sub.k by the propulsion system (16), when a collision with an object (30) is expected. In a further independent aspect, the invention relates to a formation (100) composed of several small satellites capable to fly in formation (10), wherein a relative position and flight trajectory T.sub.k of each small satellite (10) is modifiable via the independently and autonomously working collision avoidance system (18).

The present disclosure relates to an axial flux motor comprising a stator and a rotor. The stator comprises a first motor coil, a second motor coil, a first hall sensor, and a second hall sensor, and the rotor comprises a rotor platform member, an actuator magnet array arranged in an alternating axial polarity arrangement, a trigger magnet array, and a rotating magnetic return path member.

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 control system for a momentum wheel device is specified, wherein the momentum wheel device is a real momentum wheel device (1) and has a momentum wheel which is driven by a motor, and wherein a simulated momentum wheel device (2) is provided which simulates the behaviour of an ideal momentum wheel on the basis of an ideal physical model (12). The rotational speed of both the real momentum wheel device (1) and of the simulated momentum wheel device (2) can be changed by a torque command (6). A comparator device (11) is provided for comparing the real rotational angle (9) of the real momentum wheel device (1) and the simulated rotational angle (14) of the simulated momentum wheel device (2) and for generating a fault signal (15) corresponding to a deviation between the real rotational angle (9) and the simulated rotational angle (14). The fault signal (15) can be conducted to a control device (3), in order to actuate the motor on the basis of the fault signal (15), for the purpose of reducing the deviation.

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.