In an on-orbit configuration, a spacecraft having a center of mass and a pitch axis passing through the center of mass includes a main body, a first solar array, and a first thruster is operable in a geostationary orbit with the first solar array deployed, proximate to a first north or south surface of the main body, such that a rotational axis of the deployed first solar array is substantially parallel to the pitch axis. The first thruster is disposed proximate to a second north or south surface of the main body, the first thruster having a thrust vector that is approximately coaligned with the pitch axis, the second surface being opposite to the first north or south surface. No solar array is proximate to the second surface.

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, and an exhaust nozzle from which the heated propellant expands to provide an exhaust velocity of, for example, 7-16 km/sec. The launch vehicle can be utilized with the launch system, which can further comprise a launch tube formed of at least one tube, which can be electrically conductive and which can be combined with at least one insulator tube. An electrical energy source, such as a battery bank and associated inductor, can be provided.

Spacecraft propulsion systems and methods featuring a first storage tank containing a metallic propellant and a second storage tank containing a non-metallic propellant are provided. A selected one of the metallic propellant and the non-metallic propellant is supplied to an electric propulsion thruster, depending on an operational mode of the spacecraft. The metallic propellant is stored at a relatively high density, while the non-metallic propellant is stored at a lower density than the metallic propellant. Moreover, the non-metallic propellant is preferably utilized to produce thrust through the electric propulsion thruster during operational maneuvers, while the metallic propellant is reserved for producing thrust through the electric propulsion thruster during end-of-life, such as deorbiting, maneuvers.

A CubeSat satellite has a solar cell and an arc propulsion system. The arc propulsion system includes a power source, a first lead coupled to the power source and to a cathode, and a second lead coupled to the power source and to an anode. The solar cell is used as the cathode of the propulsion system.

A vacuum arc thruster device having a cathode rod disposed within a concentric insulator tube, and an anode electrode located at the distal edge of the insulator tube, separated from the cathode rod by the insulator tube. A controlled feeding mechanism moves the cathode towards the distal exit plane in a helical motion, the cathode rotating as it moves forward. The cathode rod is fixed in the center of a headless screw segment, which is rotated within a screw thread on the internal surface of a cylindrical wall of the device. As the erosion rate is concentrated at the exit plane, the screw action path enables uniform erosion around the cathode circumference, and cathode linear motion that can be matched to the radial erosion rate. The feeding rate and hence the thrust are proportional to the input power, which can be regulated by the pulse frequency.

A space vehicle (1), in particular a satellite or a probe, capable of executing high-thrust maneuvers or high specific impulse maneuvers. According to the invention, the space vehicle (1) includes at least one solid-propellant chemical thruster (10-14) and at least one electric thruster (20-24).

An example method for producing thrust includes injecting a neutral gas into a cavity between an outer electrode and an inner electrode of a thruster, ionizing the neutral gas within the cavity into a plasma, causing the plasma to form into a plasma arc between the end of the inner electrode and the exhaust orifice of the outer electrode, generating a magnetic field that applies pressure on the plasma arc, maintaining stability of the plasma arc, and exhausting the plasma arc out of the exhaust orifice based on the applied pressure of the magnetic field, thereby producing thrust.

A stationkeeping strategy for a satellite disposed in a TAP orbit includes controlling parameters of the orbit such that, for a constellation of two satellites disposed in the orbit, the constellation provides substantially continuous coverage of a polar region. The stationkeeping strategy includes one or more of: establishing an initial Right Ascension of Ascending Node (RAAN) of the operational orbit such that naturally caused orbital drift results in a mid-life RAAN of approximately 0 degrees (360 degrees); and controlling Argument of Perigee (ARGP), only indirectly, by performing orbit maintenance maneuvers only to directly control one or more of the operational orbit apogee altitude, the operational orbit perigee altitude, and inclination within a respective required range.

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, and/or compactible structures and deployment mechanisms used to form and deploy such satellite modules and power generation tiles associated therewith are provided. Each satellite module and/or power generation tile may be formed of a compactable structure and deployment mechanism capable of reducing the payload area required to deliver the satellite module to an orbital formation within the space-based solar power station and reliably deploy it once in orbit.

An arcjet thruster system for a spacecraft is provided. The arcjet thruster system may include a power supply that includes a radio-frequency start power supply and a continuous direct-current power supply, each selectively coupled to electrodes of an arcjet for initiation and maintenance of an arc between the electrodes. A radio-frequency/direct-current control module may be provided for selectively coupling the radio-frequency start power supply and a continuous direct-current power supply. The radio-frequency start power supply may be used to initiate an arc that is then sustained by the continuous direct-current power supply.