A spacecraft control system and method for determining the necessary delta-V and timing for impulsive maneuvers to change the plane of an orbit or the size of the orbit of a secondary spacecraft that is in an orbit around a primary spacecraft. The system and method uses an apocentral coordinate system for the relative orbital motion and geometric relative orbital elements to determine the required impulsive velocity change and time to maneuver, for relative orbital changes in which only one of slant or colatitude of the sinilaterating node changes.

A spacecraft control system and method for determining the necessary delta-V and timing for impulsive maneuvers to change the plane of an orbit or the size of the orbit of a secondary spacecraft that is in an orbit around a primary spacecraft. The system and method uses an apocentral coordinate system for the relative orbital motion and geometric relative orbital elements to determine the required impulsive velocity change and time to maneuver, for relative orbital changes in which only one of slant or colatitude of the sinilaterating node changes.

A rapid sequential circular assembly system for space station includes a plurality of circular space station assemblies that are connected to each other in radial profile so that a torus-shaped space station and/or cylindrical shaped space station can be built with a plurality of panels. Each of the plurality of circular space station assemblies includes a panel dispensing unit, a panel transporting cart, a first welding assembly, and a second welding assembly. The panel transporting cart is operatively coupled to a pair of rails of the panel dispensing unit, wherein the panel transporting cart travels between a storage compartment of the panel dispensing unit and the second welding assembly. The first welding assembly is connected along the pair of rails, and the second welding assembly is terminally connected across the pair of rails to seam weld each of the panels that are transported from the panel transporting cart.

A rapid sequential circular assembly system for space station includes a plurality of circular space station assemblies that are connected to each other in radial profile so that a torus-shaped space station and/or cylindrical shaped space station can be built with a plurality of panels. Each of the plurality of circular space station assemblies includes a panel dispensing unit, a panel transporting cart, a first welding assembly, and a second welding assembly. The panel transporting cart is operatively coupled to a pair of rails of the panel dispensing unit, wherein the panel transporting cart travels between a storage compartment of the panel dispensing unit and the second welding assembly. The first welding assembly is connected along the pair of rails, and the second welding assembly is terminally connected across the pair of rails to seam weld each of the panels that are transported from the panel transporting cart.

A satellite orbiting in one of a plurality of orbital planes of a satellite constellation system at an altitude range corresponding to low earth orbit includes at least one processor configured to generate satellite state data, and to generate a navigation signal based on the satellite state data. The satellite includes at least one transmitter configured to transmit the navigation signal for receipt by at least one client device on earth. Each of the plurality of orbital planes includes a corresponding one of a plurality of satellite subsets of a plurality of satellites of the satellite constellation system. Each of the plurality of orbital planes is within the altitude range, and the plurality of orbital planes includes a set of inclined orbital planes at a non-polar inclination.

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.

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.

One embodiment is directed towards a method of navigating a body. The method includes determining a respective measured direction of each of a plurality of celestial objects with respect to the body based on an output of one or more star tracking sensors mounted to the body. Calculating an expected direction of at least one of the plurality of celestial objects with respect to the body based on a current navigation solution for the body. Calculating an updated navigation solution for the body based on the expected direction of the at least one celestial object, the measured direction of the plurality of celestial objects, and an output of one or more inertial sensors mounted to the body.

One embodiment is directed towards a method of navigating a body. The method includes determining a respective measured direction of each of a plurality of celestial objects with respect to the body based on an output of one or more star tracking sensors mounted to the body. Calculating an expected direction of at least one of the plurality of celestial objects with respect to the body based on a current navigation solution for the body. Calculating an updated navigation solution for the body based on the expected direction of the at least one celestial object, the measured direction of the plurality of celestial objects, and an output of one or more inertial sensors mounted to the body.

Embodiments in accordance with the invention address potential co-orbital threats to a spacecraft through the use of a plurality of evasion pattern maneuvers selected to prevent a rendezvous with a potential co-orbital threat from occurring within a finite horizon. Embodiments in accordance with the invention maintain separation from the potential co-orbital threat while minimizing a defending spacecraft's fuel consumption.