Apparatus and methods for controlling a spacecraft for a transfer orbit. The spacecraft includes a momentum subsystem that stores angular momentum relative to a center of mass of the spacecraft, and a propulsion subsystem that includes electric thrusters. A controller identifies a target spin axis for the spacecraft, determines gimbal angles for electric thruster(s) that so that thrust forces from the electric thrusters are parallel to the target spin axis, and initiates a burn of the electric thruster(s) at the gimbal angles. The controller controls the momentum subsystem to compensate for a thruster torque produced by the burn of the electric thrusters. The momentum subsystem is able to produce a target angular momentum about the center of mass, where a coupling between the target angular momentum and an angular velocity of the spacecraft creates an offset torque to counteract the thruster torque.

A propulsion system for the orbit control of a satellite in Earth orbit driven at a rate of displacement along an axis V tangential to the orbit comprises two propulsion modules, fixed to the satellite, and facing one another relative to the plane of the orbit, each of the propulsion modules comprising, in succession: a motorized rotation link about an axis parallel to the axis V; an offset arm; and a plate supporting two thrusters, suitable for delivering a thrust on an axis, arranged on the plate on either side of a plane P at right angles to the axis V passing through a center of mass of the satellite; each of the two thrusters being oriented in such a way that the thrust axes of the two thrusters are parallel to one another and at right angles to the axis V.

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.

Embodiments of Reaction Wheel Assembly (RWA) systems are provided, which include multi-faceted bracket structures to which RWAs are mounted. In one embodiment, the RWA system includes a bracket structure, which is assembled from multiple (e.g., two to four) interchangeable panels. Each bracket panel may define or include a mount bracket to which an RWA is mounted. In certain embodiments, the bracket panels may include integral bearing cartridge features, which contain the spin bearings of the RWAs. The interchangeable panels may have interconnect features, which align and which possibly interlock to position the panels in a precise angular relationship when the multi-faceted bracket structure is assembled. In other embodiments wherein the bracket structure is assembled from two interchangeable panels or produced as a single piece, the multi-faceted bracket structure may have a peaked form factor supportive of two RWAs, which are mounted to the bracket structure in a back-to-back relationship.

Embodiments provided herein include rotatable magnets (e.g., spherical dipole magnets) disposed within sets of coils that can be used to operate the magnets as reaction/momentum spheres and/or as control moment gyroscopes. The coils are able to exert three-dimensional torques onto the magnet in order to effect attitude control of a satellite or other system. The coils can also optionally exert translational forces onto the magnet in order to maintain the magnet in position and avoid contact with static components. Diamagnetic materials can also be included to provide stabilizing repulsive magnetic forces to maintain the magnet in position and/or to reduce the necessary performance of the coils with respect to applying stabilizing translational forces.

A system for generating magnetic fields in one or more axis, the system comprising a primary electromagnet comprising a first coil having a first axis wherein the first coil is formed of a superconductor, a cooling element configured to cool the first coil below the critical temperature of the superconductor, a power source configured to energise the primary and secondary and electromagnets, wherein the primary electromagnet comprises a frame member, and wherein the frame member is suspended from at least one bracket by a thermally insulating structural member and/or a thermally insulating spring.

Disclosed is an artificial satellite including a battery pack capable of dissipating heat, at least one radiator capable of conveying the heat dissipated by the battery pack into space, and a low-dissipation equipment item having an individual power flux density of less than 250 watts/m.sup.2. The satellite includes a thermally insulating cover delimiting, together with the radiator, an interior isothermal zone in which thermal control takes place by radiation, the battery pack and the low-dissipation equipment being arranged in thermally insulating cover. The battery pack has an operating range of between 0 C. and 50 C. and preferably of between 10 C. and 30 C.

A method of controlling the attitude of a spacecraft in spinning around itself with a non-zero total angular momentum H.sub.TOT. The spacecraft includes a set of inertia flywheels configured to form an internal angular momentum H.sub.ACT. The axis of the total angular momentum H.sub.TOT is aligned with a principal axis of inertia of the spacecraft, in the course of which the inertia flywheels are controlled to form an internal angular momentum H.sub.ACT. The following expression, in which J is the inertia matrix of the spacecraft:

H.sub.actJ.sup.1(H.sub.totJ.sup.1H.sub.tot)

is negative if the principal axis of inertia targeted is the axis of maximum inertia of the spacecraft and is positive if the principal axis inertia targeted is the axis of minimum inertia of the spacecraft.

A rotation suppressing device 1 includes: a body 10; a shaft 20 extending outward from the body 10 and configured to rotate about a first rotation axis A.sub.1; a rotation part 30 configured to rotate about a second rotation axis A.sub.2 together with the shaft 20; a capture part 40 fixed to the rotation part 30 and configured to capture space debris D; a braking part 50 configured to suppress rotation of the shaft 20; and a body rotation suppressing part 60 configured to suppress rotation of the body 10 occurring when the braking part 50 operates.

A method controls an operation of a spacecraft according to a model of the spacecraft. The method determines control inputs for controlling concurrently thrusters of the spacecraft and momentum exchange devices of the spacecraft using an optimization of a cost function over a receding horizon subject to constraints on a pose of the spacecraft and constraints on inputs to the thrusters. The cost function includes components for controlling the pose of the spacecraft and a momentum stored by the momentum exchange devices. The method generates a command to control concurrently the thrusters and the momentum exchange devices according to at least a portion of the control inputs.