Described herein is a thrust system for a vehicle that includes at least three electrical power controllers, at least four electrical switches each configured to receive electrical power from at least one of the at least three electrical power controllers, and at least three thrusters each configured to receive electrical power from at least one of the at least three electrical switches. The at least four electrical switches are operable to switch a supply of electrical power from any of the at least three electrical power controllers to any one of the at least three thrusters.

A method of releasing artificial satellites into Earth's orbit includes providing an orbital transport spacecraft able to move at orbital height and comprising a cargo area, hooking a plurality of satellites in said cargo area, housing said orbital transport spacecraft in a space launcher configured to reach an orbital height, releasing said orbital transport spacecraft at orbital height, when said space launcher reaches orbital height, by imparting a separation thrust to said orbital transport spacecraft, releasing satellites in sequence from the cargo area. The release of each satellite from the cargo area occurs in a respective predetermined direction and upon the orbital transport spacecraft has reached a respective predetermined position.

A system for imparting linear momentum transfer may include a catching mechanism of a target space vehicle and a tether that is configured to impart a linear momentum transfer from the tether to the target space vehicle. The tether may be fixedly or detachably connected to a Kinetic Energy Storage and Transfer (KEST) vehicle that maneuvers and potentially retrieves the tether. Alternatively, the tether may be separate from the KEST vehicle and may be retrieved by a suitable retrieving mechanism, such as a robotic arm.

A vehicle for intercepting a target object orbiting in space is provided, comprising a launching portion for driving the vehicle into an orbit, and an interception portion for intercepting a target object when the vehicle is in orbit, wherein the interception portion comprises means for engaging with the target object and wherein the launching portion is arranged to drive the vehicle into a first elliptical orbit and the vehicle is arranged to adopt a second elliptical orbit when engaged with the target object in which the first elliptical orbit is arranged so as to intersect the orbit of the target object at an interception point, and the second elliptical orbit is such that the vehicle is arranged to move from the interception point towards the Earth's atmosphere when engaged with the target object. A method of controlling a vehicle for intercepting a target object orbiting in space is also provided, comprising controlling the vehicle to be driven into a first elliptical orbit to intersect the orbit of the target object at an interception point and controlling the vehicle to engage with the target object at the interception point and to adopt a second elliptical orbit when engaged with the target object in which the second elliptical orbit is such that the vehicle is arranged to move from the interception point towards the Earth's atmosphere when engaged with the target object.

A satellite system includes a so-called carrier satellite and a so-called piggyback satellite, each one having an Earth face. The piggyback satellite is attached to the carrier satellite by fastening elements that can be released on command. The piggyback satellite includes propulsion elements suitable for maintaining same in orbit, and the carrier satellite includes propulsion elements for performing a change of orbit of the satellite system including the carrier satellite and the piggyback satellite. The piggyback satellite is attached to the Earth face of the carrier satellite in such a way that the Earth face of the piggyback satellite is essentially perpendicular to the Earth face of the carrier satellite.

A method of guidance for placing a satellite on station comprises the following steps carried out during a predefined current cycle: A) determining on the ground a law of orientation of the thrust vector, and a history of state variables and of adjoint state variables of the satellite for the transfer from a starting orbit to a target orbit using optimal control theory, B) determining on the ground a law of evolution of the rotation of the satellite about the thrust vector, on the basis of the orientation law and of the history, C) representing according to a predetermined format the evolution of the state variables and adjoint state variables so as to obtain first parameters, D) representing according to a predetermined format a law of evolution of the rotation so as to obtain second parameters, E) concatenating the first and second parameters so as to obtain a guidance plan for the satellite, F) downloading onboard the guidance plan, G) periodically repeating according to a predefined period which is smaller than the duration of the guidance cycle: g1) reconstructing onboard the satellite a guidance instruction, g2) executing onboard the satellite the instruction by applying a closed control loop, H) measuring on the ground the real orbital trajectory of the satellite, I) repeating steps A) to H) with the trajectory measured at the end of the cycle as starting orbit of the following cycle, until the target orbit is attained.

An orbit transfer method for a spacecraft using a continuous or quasi-continuous thrust propulsion, the method comprises: the acquisition, at least once in each half-revolution of the spacecraft, of measurements of its position and of its velocity; the computation of a thrust control function as a function of the measurements; and the driving of the thrust in accordance with the control law; wherein the control law is obtained from a Control-Lyapunov function using orbital parameters, preferably equinoctial, of the spacecraft, averaged over at least one half-revolution. An embedded driving system for a spacecraft for implementing such a method and a spacecraft equipped with the driving system are provided.

Techniques for deploying a plurality of smallsats from a common launch vehicle are disclosed where a structural arrangement provides a load path between an upper stage of the launch and the plurality of spacecraft. Each spacecraft is mechanically coupled with the launch vehicle upper stage only by the structural arrangement. The structural arrangement includes at least one trunk member that is approximately aligned with the longitudinal axis of the launch vehicle upper stage, a plurality of branch members, each branch member being attached to the trunk member and having at least a first end portion that is substantially outboard from the longitudinal axis; and a plurality of mechanical linkages, each linkage coupled at a first end with a first respective spacecraft and coupled at a second end with one of the plurality of branch members, the trunk member or a second respective spacecraft.

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.

Apparatus and methods for controlling a spacecraft for a transfer orbit. The spacecraft includes a momentum subsystem that stores angular momentum relative to a center of mass of the spacecraft, and a propulsion subsystem that includes electric thrusters. A controller identifies a target spin axis for the spacecraft, determines gimbal angles for electric thruster(s) that so that thrust forces from the electric thrusters are parallel to the target spin axis, and initiates a burn of the electric thruster(s) at the gimbal angles. The controller controls the momentum subsystem to compensate for a thruster torque produced by the burn of the electric thrusters. The momentum subsystem is able to produce a target angular momentum about the center of mass, where a coupling between the target angular momentum and an angular velocity of the spacecraft creates an offset torque to counteract the thruster torque.