Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: A geosynchronous Solar Power Satellite System is created by an artificial gravity, closed ecology, multiple use structure in low earth orbit that manufactures modular solar power panels and transmitter arrays. This facility takes empty fuel tanks and expended rocket boosters from launch vehicles that are sent into low earth orbit, and re-manufactures them into structural components. These components are mated to solar cells that are launched from earth. The modular solar panels are transported to geosynchronous orbit by vehicles with ion engines, where the panels are mated to other solar panels to collect power. Structural components are also mated to transmitter elements launched from earth. These are likewise transported to geosynchronous orbit. They are mated to the solar power collecting panels and they beam the collected power back to earth.

Abstract: A system for, and method of recovering a solar-powered spacecraft from an anomaly that renders the attitude of the spacecraft unknown includes maintaining a power-safe attitude by switching between two orthogonal axes using solar panel current sensors. The system and method may also include simultaneously determining spacecraft attitude using a star sensor. The system is applicable to spacecraft operating in a solar wing-stowed configuration.

Abstract: A method of reducing propagation of threading dislocations into active areas of an optoelectronic device having a III-V material system includes growing a metamorphic buffer region in the presence of an isoelectronic surfactant. A first buffer layer may be lattice matched to an adjacent substrate and a second buffer layer may be lattice matched to device layers disposed upon the second buffer layer. Moreover, multiple metamorphic buffer layers fabricated in this manner may be used in a single given device allowing multiple layers to have their band gaps and lattice constants independently selected from those of the rest of the device.

Abstract: The present invention relates to a device making it possible to study the solar corona on a space mission requiring the formation flight of two satellites: an occulting satellite (OCC), the role of which is to create an artificial eclipse of the sun from the point of view of a coronagraph (10) onboard a second satellite, called carrying satellite (COR). The invention presents the advantage of proposing a formation flight device intended for a solar coronagraphy mission comprising fixed solar panels (11a), requiring no deployment, thanks to a dissymmetrical accommodation of the coronagraph (10) reflected in a shifting of said coronagraph (10) to a side of the carrying satellite (COR).

Abstract: The invention relates to an articulated assembly (1) comprising at least two panels (4-7) which are positioned close to a solar generator. The aforementioned panels are articulated in pairs such they can pivot between a stacked configuration, in which the panels are stacked on top of one another, and an unstacked or deployed position, in which the panels are disposed essentially in one plane, said panels being interconnected by means of a hinge. According to the invention, the hinge element is formed by at least one Carpentier coupling (11) which performs the following two functions: (i) in the stacked configuration and during deployment, the coupling generates a permanent driving torque which moves the panels into the unstacked configuration; and (ii), in the unstacked configuration, the coupling provides a mechanical restraint for the panels.

Abstract: An active vibration damping (AVD) control system for spacecraft provides a simple-to-implement and robust formulation that provides a novel damping control method for reducing spacecraft structural vibrations and improving antenna and instrument pointing. The AVD control system comprises an excitation signal generator configured to generate excitation input signals, and a damping model identification unit configured to receive system identification data and configured to produce control model parameters, the system identification data comprising the excitation input signals and information associated with motion of the spacecraft. The AVD control system further comprises an AVD control unit configured to receive the control model parameters, the AVD control unit configured to produce AVD control signals to control one or more actuators of the spacecraft.

Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: An integrated power and attitude control system and method for a vehicle efficiently supplies electrical power to both low voltage and high voltage loads, and does not rely on relatively heavy batteries to supply power during the vehicle initialization process. The system includes an energy storage flywheel, and a solar array that is movable between a stowed position and a deployed position. The energy storage flywheel is spun up, using electrical power supplied from a low voltage power source, to a rotational speed sufficient to provide attitude control. Then, after the solar array is moved to its deployed position, the energy storage flywheel is spun up, using electrical power supplied from a second power source, to a rotational speed sufficient to provide both attitude control and energy storage.

Abstract: A method and apparatus for deploying a fixed and canted solar array of a satellite. The solar array is rotated in a first plane and about a first axis from a stowed position to a first position, rotated about a second axis orthogonal to the first axis from the first position to a deployed position determined by a deployment orbit of the satellite, and locked in the deployed position to prevent further motion of the solar array relative to a satellite body for the operational life of the satellite.

Abstract: A plurality of panels are coupled rotatably by a hinge mechanism. The panels are folded when a rocket loaded with the panels is launched and unfolded in space after the artificial satellite is released. Each of the panels is loaded with at least functional elements such as a solar panel, a battery, an attitude control device and a communication device separately or together. The devices loaded on the panels are electrically connected to each other.

Abstract: The present invention provides a solar cell comprising a substrate, a first buffer layer disposed above the base layer, a second buffer layer disposed above the first buffer layer, a first boron compound layer disposed above the second buffer layer, a second boron compound layer disposed above the first compound layer, and a window layer disposed above the second compound layer, wherein the first compound layer comprises a first type of doping, wherein the second compound layer comprises a second type of doping, wherein the second buffer layer comprises a higher energy bandgap than the first compound layer, and wherein the first buffer layer and the second buffer layer permit a boron content in the first compound layer and the second compound layer to be greater than 3 %.

Abstract: A spacecraft comprises a main body; at least one elongated solar wing extending from the main body, defining a generally flat plane and comprising a pair of longitudinally extending side edges; at least one other component or structure extending from the main body and spaced at a separation distance from the at least one solar wing; and at least one solar trim tab coupled to the at least one solar wing, linearly elongated and extending in the generally flat plane in a direction transversely away from one of the longitudinally extending side edges, and sized and positioned along the longitudinally extending side edge for counteracting or compensating for one or more types of disturbance torques.

Abstract: According to invention, said elements (1.1 to 1.n) are secured to the same side (3) of a flexible inflatable mattress (4) and, when said elements are in the folded state, said mattress (4) is in the deflated state and is folded so that said elements are situated in pairs on either side of a fold (5.1 to 5.n?1) of said mattress.