Abstract: A satellite dispenser and method of using same are disclosed. In various embodiments, a satellite dispenser as disclosed herein includes a dispenser body defining an interior cavity to accommodate a payload; and a plurality of externally adjustable restraints positioned within the interior cavity and configured to be extended further into the interior cavity by actuation of a manual interface external to the interior cavity.

Abstract: In one embodiment, an apparatus is attached to a feature to be deployed on a satellite. The apparatus includes a first material having an impedance, a second material coupled to the first material configured to provide a current or voltage to the first material causing the first material to generate heat based on the impedance after a launch process of a launch vehicle carrying the satellite has completed, and a third material configured to change state at a transition temperature. A release mechanism is coupled to the third material and holds the feature in an undeployed position on the satellite. The heat generated by the second material causes the third material to change state when the transition temperature range is reached and the release mechanism is released from the third material when the third material is in the second state to deploy the feature.

Abstract: A deployable and tracked solar array mechanism for restraining and releasing deployable solar panel arrays on a spacecraft is provided. The mechanism comprises a solar array drive mountable to the top side surface of the spacecraft. A solar panel array is pivotally attached to the solar array drive and positionable against the spacecraft. At least one notch is formed in at least one of the side edges of the solar panel array. A frame is slidably mounted to the spacecraft around the solar panel array. At least one tab extends from the frame over the solar panel array and releasably restrains the solar panel array. A spring mechanism urges the frame in a direction generally away from the solar array drive assembly.

Abstract: The present invention consists of an active tie-rod device making it possible to hold and smoothly release space appendages. The active tie-rod device includes a fixed base, an active tie-rod screwed into the said fixed base, a retractable release mechanism for the active tie-rod making it possible, in the stowed position, to hold the said active tie-rod in place, and a nut for tensioning the assembly. The active tie-rod includes an internal heater, and consisting at least partially of a material with a high coefficient of thermal expansion, so that, under the action of the internal heater, the active tie-rod can expand, retracting the retractable release mechanism for the active tie-rod, and consequently allowing the smooth release of the said space appendages.

Abstract: Latching elements on a first structure are configured to securely engage corresponding bearings on a second structure, via a lateral member that drives all of the latching elements simultaneously. When engaged, or when disengaged, the system is in a stable state, requiring no active force by the controlling system to maintain the system in each state. Preferably, each latching element is coupled to the lateral member via a pinion that provides a mechanical advantage that substantially reduces the force required on the lateral member to effect the coupling or decoupling. Also preferably, the elements are formed from extruded aluminum forms, thereby providing for a relatively inexpensive and lightweight configuration that is particularly well suited for spacecraft applications.

Abstract: A flexible inflatable hinge includes curable resin for rigidly positioning panels of solar cells about the hinge in which wrap around contacts and flex circuits are disposed for routing power from the solar cells to the power bus further used for grounding the hinge. An indium tin oxide and magnesium fluoride coating is used to prevent static discharge while being transparent to ultraviolet light that cures the embedded resin after deployment for rigidizing the inflatable hinge.

Abstract: The invention relates to devices for releasing the stacking tie-rods that are used for holding deployable structures, in particular solar panels of artificial satellites. The device of the invention is self-releasing in order to enable such deployable structures to be released via the top portion of the stacking column. The invention consists in holding together the members (101) that are to be deployed by means of a tie-rod (102) having a nut (111) screwed onto its free end. The force for holding the stack together is transmitted by modules (106, 107) which are themselves held together by said force. When the stacking tension is reduced, resilient means, such as two flat spring blades (112, 113), cause the modules to move away from the tie-rod, thus enabling the members of the deployable structure to be released.

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 cable restraint/deployment mechanism is disclosed for deploying an object in space from a spacecraft. To deploy the object, the cable restraint/deployment mechanism cuts a cable attached to the spacecraft and the object. For example, a spacecraft may include a solar panel array that is restrained by the spacecraft in a stowed form through the use of a wire rope cable. The cable may be affixed at one end to a deployable object. The other end of the cable is affixed to the spacecraft through a cutting mechanism, such as a pyrotechnically actuated cutting device. The cutting mechanism cuts the cable at a swaged sleeve of the cable. The swaged sleeve of the cable substantially prevents splaying of the cable and cutting debris.