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).

A de-orbiting system for a space vehicle may include one or more lithium ion (Li-ion) batteries configured to release hot gases to be used for thrusting during de-orbiting of the apparatus. The system may also include one or more heaters surrounding each of the one or more Li-ion batteries, which are configured to send each of the one or more Li-ion batteries into a thermal runaway. The thermal runaway causes the one or more Li-ion batteries to release stored electrochemical energy within each of the one or more Li-ion batteries.

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.

A method of de-spinning a rotor of a propulsion system includes providing one or more spinning rotors rotatably mounted on a frame with a bearing having a bearing outer race, bearing balls, and bearing inner race; providing a force mechanism coupled with the one or more spinning rotors for applying a load to the one or more spinning rotors; and loading an outer portion of the outer bearing race, bearing ball, and inner bearing race of the bearing, a load on the outer portion of the bearing race, bearing ball, and inner bearing race of the bearing corresponding to a force applied to the one or more spinning rotors by the drive mechanism. The one or more spinning rotors de-spin at a rate corresponding to the load on the bearing balls.

A spacecraft has a flat antenna array having an edge and a middle portion. A reconfigurable reaction-momentum wheel is coupled to the antenna array to roll and/or pitch the antenna array in small magnitudes. The reconfigurable reaction-momentum has a reaction operating state or mode (high-torque, low momentum) and a momentum operating state or mode (low-torque, high momentum). A thruster is coupled to the antenna array to move the antenna array.

Enclosures for facilitating activities in space, and associated systems and methods, are disclosed. A representative system includes a spacecraft having an enclosed interior volume (which can be formed by an inflatable membrane) and one or more unmanned aerial vehicles (UAVs) carried by the spacecraft and positioned to deploy into the enclosed interior volume. The system can include a remote-control system to control the one or more UAVs from a terrestrial location while the spacecraft is in space. A wireless charging system can provide electrical power to the one or more UAVs. A representative method includes configuring one or more controllers to launch a first spacecraft to a first orbit, launch a second spacecraft to a second orbit, move the first spacecraft to the second orbit, dock the first spacecraft with the second spacecraft, and broadcast an event within an interior volume of the first spacecraft to a terrestrial location.

A novel cooling system for a superconducting electromagnet (740) that is suitable for use in satellite (700), or at least one or more components of the electromagnet (740) is disclosed. A satellite (700) and electromagnetic control system (705) for position control of such a satellite (700) are also disclosed. In one embodiment, the superconducting magnet control system (705) comprises at least one superconducting electromagnet (740) with at least one cooling element and at least one cryocooler (735). The cryocooler (735) is thermally coupled with the cooling element thereby enabling cooling of the superconducting electromagnet (740) or at least one or more components thereof through the cooling element solely by conduction cooling.

Systems and methods are described for a satellite control system that exhibits improved stability and increased efficiency by implementing a non-linear dynamic inversion inner-loop control algorithm coupled with an eigen vector outer-loop control algorithm. Thus, the attitude determination and control system (ADACS) may operate using commands to rotate directly about an eigen vector. Additionally, the outer-loop control system includes a feed-forward control element to enhance pointing accuracy when tracking moving targets.

A Fiber-fed Pulsed Plasma Thruster (FPPT) has an anode, a coaxial insulator, and a fiber propellant feed system. At least two cathodes insulated from each other are configured about the coaxial insulator to define an interior profile shaped into a nozzle region. At least one igniter fitted through each cathode. Wherein when the igniters are triggered, the igniters expel electrons toward the anode to ignite a primary high energy discharge between the anode and the cathodes thereby creating a plasma that vaporizes the fiber propellant. The dissociated fiber propellant combines with the primary high energy discharge to create a partially or fully ionized plasma, that is electromagnetically and electrothermally accelerated to produce predominantly {right arrow over (j)}×{right arrow over (B)}{right arrow over (j)}×{right arrow over (B)} thrust.