Techniques for placing a satellite into a highly inclined elliptical operational orbit (HIEO) having an argument of perigee of 90 or 270 include executing an orbit transfer strategy that transfers the satellite from a launch vehicle deployment orbit to the operational orbit. The launch vehicle deployment orbit is selected to have an argument of perigee of approximately 90 greater than the argument of perigee of the operational orbit, and to be substantially lower than the operational orbit. The orbit transfer strategy includes (i) an apsidal rotation of approximately 90, at least a substantial part of the apsidal rotation being attained without expenditure of any satellite propellant; and (ii) an electric orbit raising maneuver to attain an apogee altitude and a perigee altitude required by the HIEO.

This patent discloses a powerplant for an aerial vehicle comprised of Perovskite-Silicon tandem photovoltaic solar cells covering the wings and fuselage, a lithium-sulfur battery, a high-pressure unitized regenerative proton exchange membrane (PEM) device, and hydrogen tanks. The PEM device has a fuel-cell mode and an electrolysis mode. During level flight, the PEM device operates in fuel-cell mode, converting hydrogen into electricity. The electricity is used to run a plurality of pairs of permanent magnet synchronous motors, coupled to propellers, and mounted in a coaxial rotor configuration. During level flight, the array of solar cells re-charges the LiS battery pack. During takeoff and landing, the LiS battery pack supplements the electricity generated by the PEM device in fuel-cell mode. On the ground, the solar cells provides electricity to the PEM device, which operates in electrolysis mode, converting water into hydrogen gas, which is then stored in the hydrogen tanks.

A propulsion system includes a motor having a pair of parallel longitudinal tubes. Each longitudinal tube includes a first end and a second end, and delimits an internal volume filled with a fluid. Each longitudinal tube includes a projectile, configured to move longitudinally in the internal volume, fixedly secured to a propeller. Each longitudinal tube includes a mechanism to launch the projectile in the internal volume, from the first end. The propeller being arranged so as to transform a translational movement of the projectile into a rotational movement. Each longitudinal tube includes a device to slow the rotation of the projectile in the internal volume, arranged at the second end, a device to return the projectile toward the first end, and a heat dissipation device. The propellers of the longitudinal tubes are contrarotating.

One or more embodiments relates to an air-breathing plasma thruster including a thruster wall, an anode, a cathode, and at least one ring electrode. The thruster wall defines a cylindrical channel, the cylindrical channel having a first end and an opposing second end in fluid communication with the first end, where the cylindrical channel is adapted to receive incoming airflow. The anode is at the first end of the channel and the cathode is at the second end of the channel opposite the first end. The at least one ring electrode is positioned on the thruster wall.

One or more embodiments relates to an air-breathing plasma thruster including a thruster wall, an anode, a cathode, and at least one ring electrode. The thruster wall defines a cylindrical channel, the cylindrical channel having a first end and an opposing second end in fluid communication with the first end, where the cylindrical channel is adapted to receive incoming airflow. The anode is at the first end of the channel and the cathode is at the second end of the channel opposite the first end. The at least one ring electrode is positioned on the thruster wall.

An electron propulsion engine utilizes the acceleration of electrons to propel a spacecraft through space. The acceleration of the electrons in space emits electromagnetic radiation which can be used to propel the spacecraft. The radiation emission also decelerates the electrons, which allows the decelerated electrons to be recycled for reuse by the electron propulsion engine. The electron propulsion engine includes a first engine module, a second engine module, and an engine control system. The first engine module and the second engine module correspond to two mirror structures that form the electron propulsion engine. The engine control system facilitates the automatic control of the operation of the electron propulsion engine. The first engine module and the second engine module each includes a vacuum housing. The vacuum housing of each engine module is a D-shaped structure that facilitates the acceleration of the electrons and the resulting radiation emission to propel the spacecraft.

Bi-modal propulsion systems and related methods are generally described. In some embodiments, a bi-modal propulsion system may employ a single propellant for both chemical thruster(s), operating at elevated pressures, and electrical thruster(s) (e.g., electro spray thruster), operating at reduced pressures. The propellant pressure may be reduced to a desired operational range of the electrical thruster(s) using any appropriate construction including, for example, capillaries configured to reduce the pressure of the propellant to an operational range of the electrical thruster(s). In some embodiments, the reduced pressure of the propellant may be lower than a vapor pressure of at least one volatile component of the propellant, leading to the formation of bubbles within the propellant line. The presence of alternating gas and liquid phases along a flow path between a propellant tank and the electrical thruster(s) may help to electrically insulate the electrical thruster from the rest of the system.

One or more embodiments relates to an air-breathing plasma thruster including a thruster wall, an anode, a cathode, and at least one ring electrode. The thruster wall defines a cylindrical channel, the cylindrical channel having a first end and an opposing second end in fluid communication with the first end, where the cylindrical channel is adapted to receive incoming airflow. The anode is at the first end of the channel and the cathode is at the second end of the channel opposite the first end. The at least one ring electrode is positioned on the thruster wall.

A circuit for igniting and sustaining an electron discharge includes an ignitor circuit. The ignitor 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 spacecraft for the distribution of electrical energy to client craft at points situated in free space, in orbit and/or on a celestial body includes a main structure equipped with an electric thruster, with a chemical thruster and with a solar generator, a first fuel container for fuel intended for the electric thruster, and a second fuel container for fuel intended for the chemical thruster. The spacecraft is able to be modulated such that the main structure can be coupled/decoupled alternatively to/from the first container or the second container, the first container and the second container are able to be coupled/decoupled to/from one another, and the solar generator can be deployed or retracted.