The present disclosure describes a method of deploying an extensible boom from a housing. Sheets supporting respective arrays of photovoltaic devices are deployed substantially simultaneously so that a first sheet is deployed in a first direction from the housing and a second sheet is deployed in an opposite direction from the housing. Angular momentum imparted by deploying the first sheet is canceled by angular momentum imparted by deploying the second sheet. The housing can be part of a space satellite, such that the first and second sheets are deployed without causing the satellite to move out of its orbit.

Integrated power module device, systems, and methods are provided in accordance with various embodiments. For example, some embodiments include a system that may include one or more integrated power modules. Each integrated power module may include: one or more solar cells; one or more rechargeable energy storage cells; and/or one or more circuits coupling the one or more solar cells with the one or more rechargeable energy storage cells. In some embodiments, each integrated power module is configured such that the one or more rechargeable energy storage cells of the respective integrated power module are coupled with one or more back sides of the one or more solar cells. In some embodiments, at least two of the one or more integrated power modules are coupled with each other at least in parallel or in series.

A blocking diode board (“BDB”) for use with a rollable solar power module (“RSPM”) array is disclosed. The DBD includes a blocking diode, first flat electrical conductor, second flat electrical conductor, first tubular hook, and second tubular hook.

The present invention provides a method for refined attitude control based on output feedback for a flexible spacecraft. The control method comprises the following steps of: a) building a flexible spacecraft dynamical system Σ.sub.1, converting the flexible spacecraft dynamical system Σ.sub.1 into a flexible spacecraft dynamical system Σ.sub.2, and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ.sub.2; b) constructing an external system Σ.sub.3, and describing the rigid-flexible coupling dynamic disturbance through the external system Σ.sub.3; c) configuring a disturbance observer for estimating the value of the rigid-flexible coupling dynamic disturbance; d) configuring a dynamic output feedback H.sub.∞ controller; e) compounding the disturbance observer in step c) with the dynamic output feedback H.sub.∞ controller in step d) to obtain a flexible spacecraft refined attitude control system Σ.sub.6; the flexible spacecraft refined attitude control system Σ.sub.6 compensating for the rigid-flexible coupling dynamic disturbance through the estimated value.

A solar panel includes a front portion. The front portion includes an electrical insulation layer and a front face sheet layer coupled to the electrical insulation layer. The solar panel also includes a cell coupled to the front portion to produce a tile. The solar panel also includes a back portion coupled to the tile. The back portion includes a honeycomb core layer and a back face sheet layer coupled to the honeycomb core layer.

A flexible solar array for extraterrestrial deployment and a method of manufacturing such a flexible solar array are disclosed. A power generating layer, a durable layer, and an ultraviolet radiation blocking layer are disposed such that durable layer is between the power generating layer and the ultraviolet radiation blocking layer.

Apparatus and methods for controlling a spacecraft for a transfer orbit. The spacecraft includes a propulsion subsystem with electric thrusters that are installed with two-axis gimbal assemblies. The spacecraft also includes a controller that identifies a target spin axis for the spacecraft, determines an actual spin axis for the spacecraft during the transfer orbit, determines gimbal angles for the electric thruster(s) that adjust the actual spin axis toward the target spin axis, and initiates a burn of the electric thruster(s) at the gimbal angles.

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 self-deployable array of panels includes a plurality of panels, each panel having a first compressed panel thickness state and a second expanded panel thickness state, and including a spring bias element biased to the second expanded panel thickness state. A plurality of locking hinges hingedly couple each of the panels to an adjoining panel. Each locking hinge is biased to an open position. A release of stored potential energy of both of the spring bias element biased to the second expanded panel thickness state, and the locking hinges biased to the open position causes the self-deployable array of panels to self-deploy from a folded stowed state. A single part offset locking hinge is also described.

A system and method for generating electrical power from a solar power supply using at least one concentrator array having a plurality of photovoltaic cells and corresponding reflector groups to direct light to the photovoltaic. A concentration ratio indicative of a portion of the electrical power capacity to generate to power a spacecraft is determined. The concentration ratio is communicated to a control module on the concentrator array. The control module selects a number of reflectors from the total number of reflectors to orient into a photovoltaic energizing position, where the selected number of reflectors corresponds to a concentration ratio of the total number of reflectors.