A method for estimating a direction of a satellite in the transfer phase. The direction of the satellite is estimated relative to a measurement antenna by executing steps for measuring the reception power, by the measurement antenna, of a target signal emitted by the satellite, for different pointing directions of the measurement antenna. The target signal has a substantially sinusoidal component referred to as a single-frequency component. Each power measurement step includes a transposition in the frequency domain of a digital signal, obtained from a signal supplied by the measurement antenna, to obtain a frequency spectrum of the digital signal over a predetermined frequency band having the single-frequency component. The power measurement for the pointing direction being considered is determined based on a maximum value of the frequency spectrum.

The invention relates to a method for deflecting space debris comprising steps of: launching (E2) a thruster (2) by means of a sounding rocket (1) at a target altitude close to that of the one or more debris to be deflected. generating (E3) by the thruster a gas cloud (G) above the sounding rocket.

A spacecraft propulsion system includes at least one chemical thruster operable with a liquid propellant, at least one electric thruster operable with an inert gas, and a first quantity n of pressurant tanks, each of the n pressurant tanks having a substantially identical volume. The propulsion system results from assembling a plurality of subassemblies, such that a first selectable number e of the first quantity of pressurant tanks are manifolded together with the at least one electric thruster, and a second selectable number c of the first quantity of pressurant tanks are manifolded together with the at least one chemical thruster. The first selectable number e is an integer in the inclusive range of 1 to n, and c=n?e.

A constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit, a second spacecraft disposed in a second orbit, and a third spacecraft disposed in a third orbit. Each of the first orbit, the second orbit and the third orbit is substantially circular with a radius of approximately 42,164 km, and has a specified 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, the second orbit has a second RAAN (RAAN2) approximately equal to RAAN1+120?, and the third orbit has a third RAAN (RAAN3) approximately equal to RAAN1+240?. A fourth spacecraft is disposed in a fourth orbit that has a period of approximately one sidereal day, an inclination of less than 2?, a perigee altitude of at least 8000 km, and an eccentricity between approximately 0.4 and 0.66.

A method of deploying a plurality of spacecraft includes coupling the plurality of spacecraft in a stack to a payload adaptor, the stack extending along a longitudinal axis; launching the payload adaptor into an orbit; and decoupling an entirety of the stack from the payload adaptor after the payload adaptor reaches the orbit.

Embodiments herein describe maneuvering a spacecraft using a solar sail when sunlight is not available. Smaller satellites may rely solely on solar sails in order to maneuver to different locations (e.g., different orbits) to adjust for orbital decay, avoid collisions with other satellites, or to avoid space junk. However, solar sails cannot rely on the sun when orbiting on the dark side of a planet (e.g., when in the earth's shadow). When a spacecraft should maneuver but the sun is not available as a power source, the embodiments herein describe identifying other spacecraft within line-of-sight (LOS) of the spacecraft and using these spacecraft to direct lasers (or reflecting sunlight if available) at the spacecraft to maneuver it to a desired path (e.g., a new orbit).

A method for controlling phased transfer of multiple spacecraft from a separation orbit to a target orbit includes, while maintaining an in-phase relationship of the spacecraft relative to each other within the separation orbit, computing, via a control system, respective desired trajectories for a lead spacecraft and two or more follower spacecraft to reach the target orbit. The method includes establishing a constant phase offset between the spacecraft in mean anomaly of the separation orbit. During a series of transfer orbits of the spacecraft from the separation orbit to the target orbit, the method includes applying the desired trajectories via the control system such that the constant phase offset is maintained and the follower spacecraft are simultaneously transferred to the target orbit in-phase with the lead spacecraft. The control system includes a processor and computer-readable storage medium programmed with instructions for performing the method.

A space based magnetic vortex accelerator and methods of use thereof having one or more sections of magnetic material configured as a conduit with a flightpath therethrough for the spacecraft, a magnetic coil field generator electrically connected to said one or more sections of magnetic material configured to generate a space based magnetic field via said one or more sections of magnetic material, a power plant electrically connected to said magnetic coil field generator, said power plant configured to power said magnetic coil field generator, one or more magnetic field receivers affixed to the spacecraft, said one or more magnetic field receivers configured to magnetically engage said space based magnetic field.

Enclosures for facilitating activities in space, and associated systems and methods, are disclosed. A representative system includes a spacecraft having an enclosed interior volume (which can be formed by an inflatable membrane) and one or more unmanned aerial vehicles (UAVs) carried by the spacecraft and positioned to deploy into the enclosed interior volume. The system can include a remote-control system to control the one or more UAVs from a terrestrial location while the spacecraft is in space. A wireless charging system can provide electrical power to the one or more UAVs. A representative method includes configuring one or more controllers to launch a first spacecraft to a first orbit, launch a second spacecraft to a second orbit, move the first spacecraft to the second orbit, dock the first spacecraft with the second spacecraft, and broadcast an event within an interior volume of the first spacecraft to a terrestrial location.

Energy efficient satellite maneuvering is described herein. One disclosed example method includes maneuvering a satellite that is in an orbit around a space body so that a principle sensitive axis of the satellite is oriented to an orbit frame plane to reduce gravity gradient torques acting upon the satellite. The orbit frame plane is based on an orbit frame vector.