Abstract: A deployable structure that may include a slit-tube longeron and a flat panel coupled with the slit-tube longeron. The slit-tube longeron may include a tubular member having a slit that runs along the longitudinal length of the slit-tube longeron. The deployable structure may be configured to couple with a satellite. And the deployable structure may be configured to transform between a stowed state and a deployed state where the tubular member is substantially straight when the deployable structure is in the deployed state, and the tubular member is wrapped around the satellite when the deployable structure is in the stowed state.

Abstract: Deployable structures are disclosed having collapsible structural members that include a stowed configuration with a volume smaller than the expanded configuration. Such structures can include and/or be used for litters, bridges, support structures, solar arrays, and more. These structures can be easily transported to a new location and deployed from the stowed configuration into a larger functional structure. In some embodiments these structures can use one or more slit-tube longerons that can have two resting states: deployed and rolled.

Abstract: Actuator for a device for the rotary positioning of a movable element for space equipment comprising a part attached to the fixed portion of said equipment, a bearing, a sleeve designed to actuate the movable portion of said equipment and mounted on an output shaft supported by a first race of said bearing, said actuator comprising a mechanically movable input means and being capable of converting a movement of the input means into a rotary movement of the output shaft about the axis of said bearing, characterized in that said second race of said bearing is connected in rotation to the part of said equipment, in that an adjustable stop so as to angularly limit the relative rotation of the two races of the bearing, and in that the rate of rotation of the output shaft in response to a movement of the input means is different depending on whether or not the sleeve is resting on the adjustable stop.

Abstract: A system and method for the initial deployment of a space elevator using a tether, doubly spooled from each end, an inertial mass, beacon satellite and positional thrusters, whereby tether is de-spooled from geosynchronous Earth orbit altitude, using gravity and centripetal force gradient, and timed such that when the tethers fully de-spool or near the point of being fully de-spooled, are released and the free ends, one left to fall inwards to Earth anchor point, and the other left to fall outwards to tether end ballast support point, result in an Earth anchored self supporting tension structure.

Abstract: A device for deploying and aiming structural elements designed to be placed in Earth includes at least one locking/unlocking device making it possible to deactivate the first coupling mode and to activate a second coupling mode allowing a portion of the device to be aimed at a target. The device uses: a plurality of structural elements linked together by articulations, the assembly forming an articulated arm linked to a payload via a root section; the articulations having at least one pivoting connection making it possible to have two consecutive structural elements pivot relative to one another; a motor making it possible to activate at least one pivoting connection of an articulation; and a system for coupling the articulations making it possible to link the pivoting of all of the structural elements comprising a first coupling mode.

Abstract: Fuel less ACS for solar sails, using furl- and unfurl able ballast-sail-foil-segments for simultaneously displacement of center of mass and center of solar radiation pressure into opposite directions to each other. Solar-Sail-Launch-System for direct launch of the System-Sail including already docked in daughter units and payload in the sailcraft's central docking station. Unlike todays launch able solar sail designs the System-Sail features ample solar cell arrays and additional SEP-thruster-units for steering and propulsion, while the solar sail rather serves for longtime fuel less attitude controls and station keeping. The SEP-Sailcraft may also serve as a carrier-ship for daughter-units in asteroid exploration missions and is able to deliver prospector landers back to LEO with furled in foils.

Abstract: A satellite (50) is disclosed having at least one deployable thermal radiator (64). Once deployed, this thermal radiator (64) may be repositioned. In one embodiment, the deployed thermal radiator (64) may be moved about a first axis (74), about a second axis (76), or both to move the deployed thermal radiator (64) from one deployed position to another deployed position.

Abstract: A rotary positioning device of a movable element for space equipment comprising a part attached to a fixed portion of the equipment and a handle actuating the movable portion of the equipment, said handle being rigidly connected to a first race of a ball or roller antifriction bearing and said fixed part being connected to the second race of said antifriction bearing, wherein a link between said fixed part and the second race is an elastic link in rotation about the axis of said antifriction bearing and wherein a stop is rigidly attached to one of the races in order to interact with an element rigidly attached to the other race, so that the relative rotation of the two races of said antifriction bearing is limited angularly.

Abstract: A surveying device and a method for surveying a satellite and its environs is disclosed. The surveying device comprises a docking platform having an interface for physically coupling to a surface of a bus of a spacecraft and an imaging module, releasably coupleable to the docking platform via a semi-rigid tether axially extendable and retractable by command, the imaging module having an imaging sensor for collecting image data, the imaging sensor communicatively coupled with the docking platform via the tether.

Abstract: An apparatus for heat shielding and deceleration of a spacecraft comprises an unfoldable shield of individual panels made of a high temperature resistant fiber reinforced ceramic and pivotally mounted on the outer structure of the spacecraft. An unfolding ring slides axially along the body of the spacecraft. The panels are pivotally connected to the unfolding ring by compression struts acting as toggle levers. Tension springs pull the unfolding ring axially along the spacecraft body to toggle out the compression struts so as to outwardly pivotally deploy the panels. The unfoldable shield can be arranged on the forward end or the rear end of the spacecraft relative to its flight direction upon entry into the atmosphere.

Abstract: Device for deploying and retracting appendages of space systems comprising at least one support (2) carrying an appendage (1) to be deployed, which is mobile about a rotation spindle (4), and an actuator furnished with a collar (10) driving the said support in rotation by way of a coupling means (9) characterized in that the said coupling means is able in one direction of rotation to cooperate with at least one support (2.1, 2.2, 2.3) so as to drive the said support in rotation and to not cooperate with the said supports in the opposite direction of rotation.

Abstract: A deployable structure is disclosed. The deployable structure may include one or more slit-tube longerons; and one or more flat sheets coupled with the one or more slit-tube longerons. The one or more slit-tube longerons and the one or more flat sheets may be stowed by rolling the one or more slit-tube longerons and the one or more flat sheets together into a roll. In one embodiment, at least a portion of the one or more slit-tube longerons may be exposed when stowed. In another embodiment, the one or more slit-tube longerons may be manufactured from a shape memory material. These slit-tube longerons unroll into to a straight configuration when exposed to heat.

Abstract: A self opening hinge comprises a tubular member with a set of circumferentially spaced blades extending longitudinally between opposite end portions. It can be folded by bending the blades to bring the end portions together and when released will naturally return to the straight condition under the spring action of the blades. The blade which will be located on the inside of the fold when the member is in its folded condition is configured with a convex circumferential curvature as viewed from the axis of the member when in its unfolded condition, namely reversed in curvature as compared with the other blades. In this way the stress on the inside blade when the member is folded is reduced as compared with a conventional hinge where all the blades follow a circular tube profile and are concave as viewed from the central axis.

Abstract: A phased array antenna for a telecommunications satellite, that is deployable from a retracted condition to a deployed condition when the satellite is on-station, comprising a base member of hexagonal form, and a plurality of deployable antenna panels stacked one on top of the other on the base member in the retracted condition, each antenna panel being connected to a respective side edge region of the base by means of a respective back flap hinge, and the hinges having pivot points that are offset relative to one another, such that the antenna panels can be hinged sequentially one after the other from the stick to a position in which each panel is adjacent a respective base side edge region to provide an extended flat two-dimensional area when deployed.

Abstract: A device is disclosed for sequencing the deployment of a deployable structure comprising a first body, at least one second body mobile relative to the first body, and at least one third body mobile relative to either the first body or the second body and disposed, in an at least partially folded position of the structure, against or between the first and second bodies. The device comprises a first member fixedly mounted on the third body and preventing its deployment when it is immobilized and a second member mounted on the second body and immobilizing the first member until the second body has effected a selected portion of a kinematic that drives it from an initial position to a final position.

Abstract: In one aspect, a space vehicle includes a structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the structure. The structure includes at least one of an antenna mesh, a light-shielding mesh, an optical reflector mesh and a net. In another aspect, a space vehicle includes an antenna structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the antenna structure. At least one of the at least two equipment components include a solar panel, a propulsion system and an antenna feed. In a further aspect, a space vehicle includes a payload element and at least two spacecraft-support structures attached to the periphery of the payload element. At least one of the at least two spacecraft-support structures includes a propulsion system and a tracking, telemetry and control system.

Abstract: A low-energy locking hinge mechanism operable with a spacecraft to support a deployable device, and to facilitate the deployment and locking of the deployable device, the hinge mechanism comprising: (a) a support base configured to couple to a structure of the spacecraft; (b) means for securing a deployable device to the support base, the deployable device being configured to rotate about a pivot axis; (c) a one-way clutch supported about the support base and operable with the means for securing, the one-way clutch being configured to facilitate one-way rotation of the deployable device from a stowed position to a deployed position, and to lock the deployable device in the deployed position, the one-way clutch having an inherent actuatable locking function configured to prevent backward motion; and (d) an actuation member operable with the one-way clutch to deploy the deployable device from the stowed position to the deployed position, the actuation member comprising an actuation energy sufficient to actuate

Abstract: A spacecraft is provided that includes a spacecraft body, a phased array coupled to a side of the spacecraft body by a first deployment couple, and a reflector coupled to the side of the spacecraft body by a second deployment couple. The first and second deployment couples are configured to permit stowing the reflector and the phased array parallel to the side of the spacecraft body. The reflector and the phased array share a common launch restraint mounting point on the side of the spacecraft body. A spacecraft is also provided that includes a spacecraft body, a mounting platform coupled to a side of the spacecraft body by a deployment couple, and a plurality of phased array assemblies. Each phased array assembly has a face with elements, and is coupled to the mounting platform by a gimbal. The deployment couple and the gimbals are configured to permit stowing the phased array assemblies parallel to the side of the spacecraft body and with the face of each phased array assembly oriented in a first direction.

Abstract: A deployable boom with adjacent panels hinged together typically in an end to end configuration. A first longeron is attached to at least two panels and a second longeron is also attached to at least two panels. Each longeron includes a self locking hinge between adjacent panels.

Abstract: First and second releasably connectable portions of a docking system are moved together. A relatively central concave element of the second portion of the docking system contacts a corresponding relatively central mating convex element of the first portion of the docking system. A plurality of relatively distal coupling elements rigidly connected to one of the first and second portions of the docking system are inserted into a corresponding plurality of relatively distal sockets of the other of the first and second portions of the docking system. The plurality of relatively distal coupling elements are captured with a corresponding plurality of relatively distal latch mechanisms associated with the plurality of relatively distal sockets responsive to inserting the plurality of relatively distal coupling elements into the corresponding plurality of relatively distal sockets.

Abstract: The invention concerns an observation satellite (1) which is intended to be placed in orbit around a celestial body (2), and which comprises a reflecting device (5), a receiving device (6), a linking mechanical system (19), the whole forming a capture system (3) which is suitable to be able to orient the capture system (3) by gravity gradient in an aiming position in which the electromagnetic radiation corresponding to the information to be captured is received. The invention extends to an observation method and a fleet of such satellites (1).

Abstract: Provided is a space object deployment system. Specifically, the system in at least one embodiment, includes at least one deployable object. A deployment chamber is paired with each deployable object, each chamber including: a base opposite from the deployable object; and a spring loaded countermass coupled to the base and slidably disposed within the chamber, the deployment chamber at least partially disposed within a rotatable body transverse to the axis of rotation. At least one coupler is paired to each deployable object and adapted to detachably couple the deployable object to the body and constrain the countermass within the chamber. An associated method of use is provided as well.

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: A mechanical rotational drive has a bar-like torsion element, whereby the torsion element is firmly clamped with a first end in a bearing device. The torsion element can be pivoted with a second free end with respect to the clamped end about the longitudinal axis, forming a rotation axis, of the torsion element between a relaxed position and a tensioned position, whereby the length of the torsion element is reduced with respect to the relaxed position when the torsion element is brought into its tensioned position.

Abstract: A device (D) is devoted to supporting a first (M1) and a second (M2) element of a piece of space equipment, such as a telescope for example. This device (D) comprises at least two deployable flexible arms (O11 to O32) each comprising a first end (EX1) and a second end (EX2) which ends are secured respectively to the first (M1) and second (M2) elements and each designed to adopt at least one initial position folded in curves and an unfolded final position, in which positions they hold the first element (M1) away from the second element (M2) by first and second chosen distances respectively, the second distance being greater than the first.

Abstract: An outer space deployable boom system is formed of a plurality of interconnected longeron members that are linked by diagonal members and batten frames. The diagonal members are arranged in a unique K-type configuration that provides a predictable and orderly folding dynamic as the boom system transitions from a stowed state to a deployed state.

Abstract: The invention relates to a thermal control device intended to dissipate the heat generated by a payload on a spacecraft, comprising a number of surfaces and including means for circulating a refrigerant. An evaporation zone (Z1) comprises means for circulating the refrigerant, a compression zone (Z2), a condensation zone (Z3) comprising at lEast one radiating panel, linked to a part of the means for circulating the refrigerant, including several branches and comprising means to allow or inhibit the circulation of the refrigerant within these branches so as to vary the area of the heat exchange surface in the condensation zone, a pressure reduction zone (Z4) comprising means for circulating the refrigerant. Such a device is particularly well adapted to thermal problems encountered in telecommunications satellites.

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 truss structure for supporting a membrane that is to be deployed in space, where the truss structure maintains the membrane sufficiently flat with uniform tension. The truss structure includes a circular perimeter truss and a transitional truss. The transitional truss includes a series of ties that are coupled to the perimeter truss and the membrane. When the truss structure is deployed, telescopic parts of the perimeter truss that are stowed in expansion cause the transitional truss to expand and pull the membrane into the flat configuration.

Abstract: A method and apparatus for controlling a plurality of solar panels of a spacecraft is described. The method comprises the steps of providing a first step command to a first solar panel, and providing a second step command to a second solar panel at a time of a transient zero-crossing of a dynamic response of the spacecraft body to the first step command, wherein the second solar panel is disposed on an opposite side of the spacecraft from the first solar panel. The apparatus comprises a processor; a first solar panel driver, communicatively coupled to the processor, for providing a first step command to a first solar panel, and a second solar panel driver, communicatively coupled to the processor, for providing a second step command to a second solar panel at a time of a transient zero-crossing of a dynamic response of the spacecraft body to the first step command.

Abstract: A low-energy locking hinge mechanism operable with a spacecraft to support a deployable device, and to facilitate the deployment and locking of the deployable device, the hinge mechanism comprising: (a) a support base configured to couple to a structure of the spacecraft; (b) means for securing a deployable device to the support base, the deployable device being configured to rotate about a pivot axis; (c) a one-way clutch supported about the support base and operable with the means for securing, the one-way clutch being configured to facilitate one-way rotation of the deployable device from a stowed position to a deployed position, and to lock the deployable device in the deployed position, the one-way clutch having an inherent actuatable locking function configured to prevent backward motion; and (d) an actuation member operable with the one-way clutch to deploy the deployable device from the stowed position to the deployed position, the actuation member comprising an actuation energy sufficient to actuate

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.

Abstract: A positioning boom disposed between a space-based observatory platform and its instrument payload provides thermal and dynamic isolation as well as fine pointing and momentum control. The inventive system isolates a sensitive payload from a warm, dynamically noisy spacecraft, which includes a sunshield. Isolation is required in terms of dynamics and heat flow, both in terms of the absolute level and its variance (thermal isolation). The present invention provides intrinsic control over momentum buildup (which is due to the separation of the center of pressure from the center of mass). The space-based platform also provides a view (field of regard) to at least half the sky (in the anti-sun direction).

Abstract: A remotely controlled apparatus for grasping an object such as a satellite in space includes a free-flying grasper unit connected by a selectively extendable cable to an orbital platform. The orbital platform and the grasper unit are preferably each independently propelled by respective propulsion systems including multi-axis maneuvering nozzles respectively provided on the orbital platform and on the grasper unit. The grasper unit preferably includes a grasper unit body and a grasper-arm mechanism having a controllable grasping claw articulately mounted on the end of a controllable telescoping arm that extends from the grasper unit body. The grasper unit preferably further includes a camera and a distance measuring sensor to provide visual and distance feedback data for assisting the remotely controlled maneuvering of the grasper unit and actuation of the grasper-arm mechanism.

Abstract: The invention relates to a pyromechanical separating element having a hermetically sealed pyrotechnic pressure element (2), which is installed in a housing (1) and has a gas-forming pyrotechnic charge, and a detachable latching pin (5) which is separated from the pressure element (2) by a driving volume (19) and is inserted into the housing (1), wherein a first securing point is arranged on the housing (1) and a second securing point is arranged on the latching pin (5), and the latching pin (5) is anchored on the housing (1) by way of an arresting and force-limiting element (8).