Systems and methods are described herein for mounting a thruster onto a vehicle. A thruster mounting structure may comprise a first, second, and third rotational joint, a boom, and thruster pallet, and a thruster attached to the thruster pallet. The first rotational joint may be attached to the vehicle and configured to rotate in a first axis. The first rotational joint may be connected to the boom and configured to pivot the boom about the first axis. The boom may be connected to the second rotational joint, which is connected to the third rotational joint and configured to rotate the third rotational joint in the first axis. The third rotational joint may be connected to the thruster pallet and configured to pivot the thruster pallet in a second axis that is perpendicular to the first axis.

A circuit for igniting and sustaining an electron discharge includes an ignitor circuit. The igniter circuit includes a high voltage transformer and a switch connected in series between a primary of the transformer and a DC source return. The switch is configured to receive a driving signal. A reset circuit is connected in parallel to the primary of the high voltage transformer. A first rectifier is connected in series between a secondary of the high voltage transformer and a keeper. A terminal of the secondary of transformer is connected to a cathode. The circuit for igniting and sustaining the electron discharge also includes a sustaining circuit having a current source with a return connected to a cathode and a second rectifier connected in series between the current source and the keeper.

A satellite propulsion system and methods of operating the same include a first ionization stage and a second acceleration stage. The first ionization stage has a plasma source configured to produce an arc discharge and emit a preliminary plasma. The plasma source includes an external magnetic field configured to magnetize the arc discharge. The second acceleration stage has an accelerator positioned in series with the plasma source. The accelerator is configured to accelerate the preliminary plasma out through the accelerator, thereby creating an accelerated plasma flow. The application of an activation threshold voltage to the accelerator results in a surge in system performance parameters.

A dual mode engine for propelling a spacecraft, including a combustion chamber having a flange end, an open nozzle end, and an enclosed chamber portion extending therebetween, a propellant tank in fluidic communication with the combustion chamber, an electronic controller, a power source operationally connected to the electronic controller, and a fluid flow motivator operationally connected to the electronic controller and connected in fluidic communication with the propellant tank. The engine further includes a chemical propulsion portion further including a propellant inlet port operationally connected to the combustion chamber and disposed adjacent the flange end, an ignition trigger electrode positioned in the combustion chamber adjacent the propellant inlet port and operationally connected to the electronic controller and operationally connected to the power source, wherein the propellant inlet port is fluidically connected to the propellant tank. The engine also includes an electric propulsion portion, further including at least two spaced electrodes for ionizing the propellant positioned in the combustion chamber adjacent the nozzle end, a plurality of attitude control thrusters operationally connected to the electronic controller and in fluidic communication with the propellant tank, and a plurality of respective valves, each respective valve fluidically connected between a respective attitude control thruster and the propellant tank, wherein the propellant inlet port is fluidically connected to the propellant tank.

Techniques for deorbiting a satellite include executing an orbit transfer maneuver that transfers the satellite from an operational orbit to an interim orbit. The operational orbit is substantially geosynchronous and has (i) an inclination of greater than 70 degrees; (ii) a nominal eccentricity in the range of 0.25 to 0.5; (iii) an argument of perigee of approximately 90 or approximately 270 degrees; (iv) a right ascension of ascending node of approximately 0; and (v) an operational orbit apogee altitude. The interim orbit has an initial second apogee altitude that is at least 4500 km higher than the first apogee altitude, and the interim orbit naturally decays, subsequent to the orbit transfer maneuver, such that the satellite will reenter Earth's atmosphere no longer than 25 years after completion of the orbit transfer maneuver.

The present disclosure relates to a launch system, a launch vehicle for use with the launch system, and methods of launching a payload utilizing the launch vehicle and/or the launch system. The disclosure can provide for delivery of the payload at a terrestrial location, an Earth orbital location, or an extraorbital location. The launch vehicle can comprise a payload, a propellant tank, an electrical heater wherein propellant, such as a light gas (e.g., hydrogen) is electrically heated to significantly high temperatures, an exhaust nozzle from which the heated propellant expands to provide an exhaust velocity of, for example, 7-16 km/sec, and sliding electrical contacts in electrical connection with the electrical heater. The launch vehicle can be utilized with the launch system, which can further comprise a launch tube formed of concentric electrically conductive tubes, as well as an electrical energy source, such as a battery bank and associated inductor.

A spacecraft includes a first deployable module and a second deployable module and is reconfigurable from a launch configuration to an on-orbit configuration. In the launch configuration, the first deployable module is adjacent to the second deployable module. The first deployable module includes a first solar array, the first solar array being rotatable, in the on-orbit configuration, about a first axis of rotation, and the second deployable module includes a second solar array, the second solar array being rotatable, in the on-orbit configuration, about a second axis of rotation, the second axis of rotation being separated by a substantial distance from the second axis of rotation.

A hybrid propulsion engine for a satellite that includes a satellite housing and electronics carried by the satellite housing, may include a rechargeable electrical power source carried by the satellite housing. The hybrid propulsion engine may also include a combustible fuel supply carried by the satellite housing, and an engine housing defining a combustion chamber for fuel from the combustible fuel supply. The hybrid propulsion engine may also include an electrical arc heater associated with the combustion chamber and configured to supply additional heat from the rechargeable electrical power source, and an exhaust nozzle downstream from the combustion chamber.

A space-based solar power station, a power generating satellite module and/or a method for collecting solar radiation and transmitting power generated using electrical current produced therefrom is provided. Each solar power station includes a plurality of satellite modules. The plurality of satellite modules each include a plurality of modular power generation tiles including a photovoltaic solar radiation collector, a power transmitter and associated control electronics. The power transmitters can be coordinated as a phased array and the power generated by the phased array is transmitted to one or more power receivers to achieve remote wireless power generation and delivery. Each satellite module may be formed of a compactable structure capable of reducing the payload area required to deliver the satellite module to an orbital formation within the space-based solar power station.

A thruster for a micro-satellite is disclosed. The thruster includes a voltage source, an inductor and a resistor. A switching device is coupled to the inductor and the resistor. The thruster includes an exterior electrode composed of a first propellant, an insulator located coaxially within the exterior electrode and an interior electrode composed of a second propellant located coaxially to the insulator and the exterior electrode. An exterior housing has a proximate end and an opposite distal end with a thrust channel. The exterior housing holds the exterior electrode, the insulator and the interior electrode. The switching device is coupled to either the exterior electrode or the interior electrode. The switching device is switched to pulse voltage from the inductor to create an arc between the exterior electrode and the interior electrode. Either the exterior electrode or the interior electrode may serve as a cathode to generate a plasma jet.