A propulsion apparatus includes: first and second conducting wires fixed with an interval therebetween; and a power supply configured to output an alternating current. The alternating current output from the power supply having a phase difference of 90 degrees is made to flow through the first and second conducting wires.

A propulsion system for spacecraft is based on an electric engine that expels propellant to achieve thrust. The propellant is first ionized to generate a plasma. Plasma particles are selectively accelerated via a pulsed laser that accelerates predominantly the electrons in the plasma. The electrons are expelled first, forming a space charge that acts as a virtual cathode to accelerate the positive ions. Interactions between the laser beam and plasma electrons are predominantly through the ponderomotive force.

In one embodiment, a method includes collecting and compressing air with an air-scoop that includes an air inlet that air molecules enter the air-scoop through at approximately an orbital speed when the air-scoop is moving through an atmosphere at approximately the orbital speed and a rotor configured to be rotated by a motor at approximately a rotational speed. The method also includes providing the collected and compressed air for supply to a thruster as propellant.

A system for regulating the flow rate of a propellant fluid for an electrical thruster of a space vehicle, the vehicle including a tank of propellant fluid and a flow rate regulator connected to the outlet of said tank; the flow rate regulator including a heater element controlled by a computer and adapted to heat the propellant fluid and to modify its physical properties so as to vary the flow rate of propellant fluid leaving the tank; said system being characterized in that the computer also includes a plurality of empirical calibration curves that have been determined empirically for defining the flow rate of propellant fluid as a function of the magnitude of heating and as a function of environmental parameters, such that said computer also performs a function of determining the flow rate of the propellant fluid.

Methods and apparatus to methods and apparatus for performing propulsion operations using electric propulsion system are disclosed. An apparatus includes a space vehicle including means for performing propulsion operations without using a chemical propulsion system.

A device for forming a quasi-neutral ion-electron beam, including: a chamber; a set of means for forming an ion-electron plasma in the chamber; and means for extracting and accelerating charged particles from the plasma out of the chamber. The particles are capable of forming the beam and the extraction and acceleration means that include a set of at least two grids located at one end of the chamber.

Systems, methods, and apparatus for a structural propellant for ion rockets (SPIR) are disclosed. In one or more embodiments, a method for in-space propulsion of a spacecraft involves removing, by a removal device, a portion of a structure of the spacecraft. The method further involves feeding, by the removal device, the portion into a Hall thruster system. Further, the method involves utilizing, by the Hall thruster system, the portion as propellant to produce thrust. In one or more embodiments, the structure is an upper stage of the spacecraft. In at least one embodiment, the upper stage comprises at least one structural support and/or at least one upper stage housing. In some embodiments, the structure is manufactured from magnesium, bismuth, zinc, and/or indium.

An electrospray diffusion pump with an upper vacuum chamber coupled to a lower vacuum chamber by a cylinder having an aperture in the upper chamber at the center of a conductive extractor ring. A conductive tube is positioned in the upper chamber and is axially aligned with the conductive extractor ring. The conductive tube is coupled to receive a conductive or semi-conductive spray fluid. A voltage source is coupled between the conductive tube and the extractor ring and adjusted to form a Taylor Cone that provides a jet of charged droplets at the tip of the conductive tube, the charged droplets are attracted to the extractor ring and pass into the aperture, then through the cylinder into the second chamber. The charged droplets have nearly zero vapor pressure and transfer ambient gas at a first pressure from the upper chamber to the lower chamber at a lower pressure.

An aerodynamic platform or spacecraft including a habitable 1G centripetal force rotating gravity producing interior corridor within an aerodynamic shell and an aerodynamic drone booster launch system with reentry and reuse capability.

A method of providing a suitable candidate liquid for an electrospray system is provided. At a first step an aperture radius for an aperture of the electrospray system (10) through which the liquid to be electrosprayed is drawn is obtained. Next, a corona threshold electric field curve as a function of relative permittivities of candidate liquids is calculated to determine the electric field at which undesirable corona discharge will occur. The maximum surface tension that can be electrosprayed by the system is calculated and then a candidate liquid which has a chosen relative permittivity and a surface tension that is equal to or less than the maximum surface tension is provided, to thereby provide a suitable candidate liquid with an appropriate surface tension to result in electrospray that meets the requirements of the electrospray system.