An apparatus for providing optimized power balanced variable thrust transfer orbits to minimize an electric orbit raising duration is disclosed. The electric orbit raising includes a first transfer orbit and a target orbit. The apparatus includes control electronics configured to transfer to a second transfer orbit to reach the target orbit. A variable thrust based on a current electric power balance is determined. The control electronics are further configured to execute a maneuver to transfer from the first transfer orbit to the second transfer orbit according to the determined variable thrust and a predetermined maneuver plan. The predetermined maneuver plan includes a set of compound steering parameters. The set of compound steering parameters are based on an optimized variable thrust and an associated electrical power balance to the optimized variable thrust. An optimized series of transfer orbits minimizes the electric orbit raising duration.

A method for autonomously controlling electric power supplied to a thruster of a spacecraft during electric orbit raising includes determining a state of charge of a battery onboard the spacecraft at an entry into an eclipse during each orbit of a plurality of orbits during the electric orbit raising of the spacecraft. The method also includes determining an electric power level used to fire each thruster of a plurality of thrusters during each orbit beginning after the eclipse, based at least on the state of charge of the battery, and that will provide a shortest electric orbit raising duration and minimize thruster propellant usage during electric orbit raising.

A method for providing optimized power balanced low thrust transfer orbits utilizing split thruster execution to minimize an electric orbit raising duration of an apparatus includes monitoring an electric power balance on the apparatus. The method also includes firing a first thruster in response to the apparatus exiting an eclipse and based on the electric power balance. The method additionally includes firing a second thruster at a predetermined time delay after firing the first thruster based on the electric power balance. The method additionally includes ending firing one of the first thruster or the second thruster after a predetermined time duration based on the electric power balance. The method further includes ending firing another of the first thruster or the second thruster in response to the apparatus entering a next eclipse.

Methods and apparatus to methods and apparatus for performing propulsion operations using electric propulsion system are disclosed. An example apparatus includes means to use an electric propulsion system coupled to a frame of a spacecraft, the electric propulsion system including at least a first thruster and a second thruster, the first thruster adjacent a first side of the frame, the second thruster adjacent a second side of the frame, and means to allow at least one of the first thruster or the second thruster to control the spacecraft without using a chemical propulsion system.

A system and method for highly efficient constellations of satellites which give single, double, . . . k-fold redundant full earth imaging coverage, or k-fold coverage for latitudes greater than any selected latitude is given for remote sensing instruments in short periods of time, i.e., continuous coverage, as a function of the parameters of the satellite and the remote sensing instrument for many different types of orbits. A high data rate satellite communication system and method for use with small, mobile cell phone receiving and transmitting stations is also provided. Satellite instrument performance models, full and partial satellite constellation models, and satellite cost models are disclosed and used to optimize the design of satellite systems with vastly improved performance and lower cost over current major satellite systems.

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 first spacecraft and a second spacecraft are configured to be disposed together, in a launch configuration, for launch by a single launch vehicle. In the launch configuration, the first spacecraft is mechanically coupled with a primary payload adapter of the launch vehicle, and the second spacecraft is mechanically coupled with the first spacecraft by way of an inter-spacecraft coupling arrangement. The spacecraft are configured to be deployed, following injection into a first orbit by the launch vehicle, by separating the first spacecraft from the primary payload adapter while the second spacecraft is mechanically coupled with the first spacecraft. A first onboard propulsion subsystem of the first spacecraft includes one or more thrusters configured to execute an orbit transfer maneuver from the first orbit to a second orbit. A propellant line arrangement detachably couples the first onboard propulsion subsystem with a second propellant storage arrangement on the second spacecraft.

A satellite constellation forming system forms a satellite constellation (20) having a plurality of orbital planes (21) in each of which a plurality of satellites fly at the same orbital altitude. A satellite constellation forming unit forms the satellite constellation (20) in which orbital altitudes of the orbital planes (21) are mutually different. Furthermore, in the satellite constellation (20), relative altitude differences between adjacent orbital planes in the plurality of orbital planes are sequentially arranged to be sinusoidal. The satellite constellation forming unit sequentially changes an orbital altitude (23) of each orbital plane of the plurality of orbital planes while maintaining a sinusoidal arrangement of the relative altitude differences between adjacent orbital planes in the plurality of orbital planes.

A method of releasing spacecraft from a rocket includes arranging spacecraft in a stack on a rocket, launching the rocket until it attains orbital velocity, imposing a flat spin on the rocket, and releasing the entire stack from the rocket during the flat spin. In one aspect, the method includes applying a compressive load along a length of the stack, launching the rocket until it attains orbital velocity, imposing a flat spin on the rocket, and releasing the compressive load to allow the entire stack to separate from the rocket during the flat spin.

A system for optimizing a low-thrust trajectory of a spacecraft trajectory for orbital transfer includes an interface to receive data, a memory to store scheduled geostationary transfer orbit (GTO) data and scheduled geostationary Earth orbit (GEO) data and computer-executable programs, and a processor. The processor is configured to provide a two-dimensional (2D) averaged trajectory consisting of a predetermined revolutions by executing the optimal control program using the GTO data and GEO data, arrange N equidistant points on the 2D averaged trajectory to form segments on the 2D averaged trajectory, obtain osculating elements corresponding to the segments by solving optimization problems for the segments, estimate initial guesses of the segments under continuous thrusting conditions in tangential directions at the N equidistant points, solve a minimum energy optimization problem to obtain a minimum energy 2D osculating trajectory by using as initial guess the concatenation of segments, compute a minimum energy three-dimensional (3D) osculating trajectory by linearly decreasing an inclination of the minimum energy 2D osculating trajectory to zero, and generating a minimum fuel 3D osculating trajectory by iteratively solving a cost function while changing a parameter from one to zero.