A reconfigurable power processing unit for a spacecraft including a plurality of power modules. Each of the power modules includes a first power source and a second power source. The first power source and the second power source are configured to be in series in a first state and in parallel in a second state. A plurality of contactors connect each power module to at least one of another power module in the plurality of power modules and a power processing output and are configured to control the state of the power modules.

A Fiber-fed Pulsed Plasma Thruster (FPPT) has an anode, a coaxial insulator, and a fiber propellant feed system. At least two cathodes insulated from each other are configured about the coaxial insulator to define an interior profile shaped into a nozzle region. At least one igniter fitted through each cathode. Wherein when the igniters are triggered, the igniters expel electrons toward the anode to ignite a primary high energy discharge between the anode and the cathodes thereby creating a plasma that vaporizes the fiber propellant. The dissociated fiber propellant combines with the primary high energy discharge to create a partially or fully ionized plasma, that is electromagnetically and electrothermally accelerated to produce predominantly {right arrow over (j)}{right arrow over (B)}{right arrow over (j)}{right arrow over (B)} thrust.

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 constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit. Each of the first orbit and the second orbit is substantially circular with a radius of approximately 42,164 km. The first orbit and the second orbit have a respective inclination with respect to the equator within a range of 5 to 20. The first orbit has a first right ascension of ascending node (RAAN1) and the second orbit has a second RAAN (RAAN2) of approximately RAAN1+90.

A constellation of Earth-orbiting spacecraft, the constellation having an orbital maneuver lifetime life (OML), includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit, each of orbit being substantially circular with a radius of approximately 42,164 km and having a respective inclination with respect to the equator specified within a range of 10 to 20. The first orbit has, at beginning of life (BOL), a first right ascension of ascending node (BOL-RAAN1) and the second orbit has, at BOL, a second RAAN (BOL-RAAN2) the BOL-RAAN1 and the BOL-RAAN2 being separated by a first angular separation -RAAN1. A first stationkeeping delta-V (V1) applied over the OML to the first spacecraft, in combination with a second delta-V (V2) applied over the OML to the second spacecraft, maintains the -RAAN1 approximately constant and an actual inclination within specification, and V1 approximately equals V2.

A space propulsion system includes an electrostatic thruster with a first electrical load; a resistojet; a propellant fluid feed circuit; and an electrical power supply circuit including a first power supply line and a first switch for selecting between connecting the first power supply line to the resistojet and connecting the first power supply line to the first electrical load of the electrostatic thruster. The propulsion system thus enables a space propulsion method to be applied that includes a switching step for selecting a first propulsion mode in which the resistojet is activated, or a second propulsion mode in which the electrostatic thruster is activated.

Disclosed is a method (70) for monitoring a phase for transferring a satellite (20) from one earth orbit, called initial orbit, to another earth orbit, called mission orbit, in particular a transfer using electric propulsion unit. The monitoring method includes a step for estimating the direction of the satellite during the transfer phase by way of an earth array antenna (30) including a plurality of elementary antennas (31), each elementary antenna having a primary radiation lobe with a width greater than or equal to 20, the elementary antennas (31) being oriented such that their respective fields of vision overlap, the direction of the satellite being estimated based on at least one useful phase difference measurement between signals corresponding to a target signal, transmitted by the satellite and received on a pair of elementary antennas (31).

A deployment algorithm that populate a group of companion satellites around a center satellite, and a control algorithm for maintaining the companion satellites in close formation around the center satellite.

A method to control a sunlight acquisition phase of a spacecraft with a nonzero angular momentum of an axis D.sub.H. The spacecraft includes a solar generator configured to rotate about an axis Y. The spacecraft actuators are controlled to place the spacecraft in an intermediate orientation in which the axis Y is substantially orthogonal to the axis D.sub.H. The solar generator is controlled to orientate the solar generator towards the sun. The spacecraft actuators are controlled to reduce the angular momentum of the spacecraft. The actuators of the spacecraft engine are controlled to place the spacecraft in an acquisition orientation in which the axis Y is substantially orthogonal to the direction of the sun with respect to the spacecraft.

Disclosed is a thrust-producing device that can generate force in any medium (in air, in space, underwater). The device, in and of itself, does not have an external intake, or exhaust. Thrust is produced when electro-mechanical and/or thermal energy is applied to a closed, pressurized container of gas so as to produce an imbalance in the internal pressure within the container. A measure of corrosion prevention can be achieved by using N.sub.2, an inert gas, or some other gas that will not chemically react with the other constituents in the container. Although this invention was principally intended as an advancement in the field of space propulsion systems, it has wide applicability across all modes of transportation; and has applications in stability and control, as well as propulsion.