An apparatus includes a satellite in the form of a plate having a thickness being smaller than a width of the satellite. The apparatus also includes a plurality of contact points distributed on a face of the satellite, allowing for one or more additional satellites to be stacked upon the satellite.

A method for creating an artificial reentry corridor. Several modules are deployed in a retrograde orbit relative to target debris. Each module releases a gas plume which, in turn creates an artificial reentry corridor. The debris passes through the corridor and becomes decelerated.

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 predicting the orbit of a satellite of a satellite positioning system, including: associating first and second types of satellites with first and second models of celestial mechanics forces, respectively; storing first ephemerides data of a satellite, associated to first time intervals and second ephemerides data associated to second time intervals. Further, the method comprises: calculating reference satellite positions based on the first ephemerides data; estimating first and second satellite positions in the first time intervals by using the second ephemerides data and the first and second forces models, respectively; determining first and second estimate errors by comparing the reference positions with the first and second positions, respectively; and detecting the type of satellite between the first and second types by an analysis of the first and second errors.

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 satellite with a paraboloid mirror fabricated while in space is described. The mirror is formed by solidifying liquid precursor material after its surface assumes a paraboloid shape as a result of compound rotation of the satellite. The mirror is preferably formed from a photopolymer which creates a rigid paraboloid mirror surface upon exposure to a cross-linking radiation source. Optical coating(s) deposition system is described. Several deployable satellite structures, including mirror support are executed in shape memory materials and are deployed by application of heat.

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 and apparatus provides cube-shaped satellite primary structures, each comprised of six identical, or nearly identical, rectangular truss panels and internal struts. The struts, all adjustable in length, connect, and are directed between all cube opposite corners and all cube opposite panel centers. All struts meet at the cube center where they attach rigidly to either a block called the “nucleus fitting” or to a hollow sphere. Each strut attaches to either a ball-socket corner fitting located at the interior corner of the cube, or to a ball-socket panel center fitting located at the panel center interior to the cube.

Systems and methods are provided for calculating satellite access windows for a constellation of imaging satellites. In some implementations, systems and methods are provided for managing execution of native programs on high performance computing systems. In one embodiment a system can determine, for each time interval within a first period of time, a position of each imaging satellite of a constellation of imaging satellites. The system can transform, for each time interval within the first period of time, the position of each imaging satellite from a first coordinate system to a second coordinate system. The system can determine an access window for at least one imagining satellite based at least in part on a determined angle between the vector to the respective location and a determined vector to the respective satellites. The system can schedule the at least one imaging satellite to perform a task within the access window.

A projection unit (12) of an interference power estimation device (1) projects an orbit of a satellite onto a map representing a ground surface. A range acquisition unit (13) determines a plurality of ranges on the map so that the projected orbit is included in the ranges. An altitude calculation unit (14) calculates an altitude of the orbit of the satellite in each of the ranges. A range interference calculation unit (16) calculates, for each of the ranges, an interference power between the satellite at a position determined by a latitude and a longitude of the range and the altitude calculated for the range and a radio station installed on the ground surface. An estimation result calculation unit (17) selects, as an estimation result, a maximum value among the interference powers calculated for each of the ranges.