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.

Apparatus, computer-readable medium, and method for calculating a mass of propellant in a propellant storage volume. A density of the propellant is calculated based on a detected pressure and temperature of the propellant and actual thermodynamic properties of the propellant. A volume of the propellant storage volume is calculated based on a measured nominal volume and changes to the nominal volume based on pressure. The mass of the propellant is calculated by multiplying the calculated volume of the propellant storage volume by the calculated density of the propellant in the propellant storage volume.

The invention is a Hall thruster that does not have any discharge channel, and magnetic pole piece. The Hall thruster utilizes permanent magnets to produce magnetic field with strong radial component in front of an annular anode, and expands propellant directly into vacuum through the anode acting also as a gas distributor. The invention reduces mass and complexity of conventional Hall thrusters, and offers a radical solution to discharge channel and magnetic pole piece erosion problem.

A space vehicle system, a method of manufacturing a multiple space vehicle system, and a method of disposing space vehicles into Earth orbit are disclosed. The space vehicle system may include a first space vehicle including a first core structure with a first wall thickness. The space vehicle system may include a second space vehicle including a second core structure with a second wall thickness, the second wall thickness different from the first wall thickness, and the second core structure releasably attached to the first space vehicle in a stacked configuration.

An ion thruster, comprising: a discharge chamber for accelerating ions towards one direction; an inflow opening for intake of a propellant into the discharge chamber; a discharge cathode, shaped in a form of a propeller, for releasing electrons in the discharge chamber, thereby ionizing the propellant in the discharge chamber, wherein the discharge cathode is rotatable around an axis, thereby propelling inward to the discharge chamber the propellant thereof; an outflow opening for exhausting the ions from the discharge chamber; and an accelerator electrode, shaped in a form of a propeller, for accelerating the ions towards the one direction of the outflow opening, wherein the accelerator electrode is rotatable around an axis, thereby propelling outward from the discharge chamber the ions and neutral atoms thereof; wherein the ion thruster comprises electromagnetic coils for generating a magnetic field inside the discharge chamber.

Various material compositions of a heat shield for a spacecraft are described. The heat shield can be formed by multi-dimensional weaver or three-dimensional (3-D) printer. Furthermore, the heat shield can be configured with a superconducting coil.

An electric propulsion device for a satellite equipped with at least four active thrusters exerting a parallel thrust upon a transfer, the device comprises means for detecting a thruster failure and means for reorienting the thrusters, and comprises means for computing a reorientation angle of the thrusters remaining active upon a failure of a thruster, the value of the angle being computed to reorient at least two of the remaining thrusters in order to cancel the total torque about the center of mass of the satellite.

Methods and apparatus to methods and apparatus for performing propulsion operations using electric propulsion system are disclosed. An example method includes deploying a space vehicle including an electric propulsion system; and using the electric propulsion system for attitude control and orbit control, no other propulsion system provided to enable the attitude control and the orbit control.

An ion electrospray device with a porous reservoir and at least one porous emitter includes a porous compliant interface sandwiched between the porous reservoir and the porous emitter for transferring fluid from the porous reservoir to the porous emitter. The interface has a characteristic capillary pressure stronger than the characteristic capillary pressure of the porous reservoir to fill the porous emitter with fluid via a fluid injection section of the interface and before the porous reservoir is then partially filled with fluid. Emitter leakage and propellant bridging problems are addressed.

A space vehicle engine (10) comprising a chemical thruster having a nozzle (30) for ejecting combustion gas, together with a Hall effect thruster. The engine is arranged in such a manner that the nozzle serves as the ejection channel for particles ejected by the Hall effect thruster when it is in operation. The engine can deliver high thrust with low specific impulse or relatively low thrust with large specific impulse.