Abstract: A structure for changing the attitude of a spacecraft is provided including an elongated member and a plurality of actuators coupled to the elongated member. The elongated member is coupled to the spacecraft. The elongated member has a length and an angular position relative to an axis. The plurality of actuators are configured to deflect the angular position of the elongated member relative to the axis and to alter the length of the elongated member.

Abstract: Femto-satellites are very small satellites that can be deployed in constellations from a larger mothership satellite for distributed measurement. They are too small to accommodate the GNSS receivers that many satellites use for navigation, but they can be located with an electromagnetic beam from the mothership satellite. The mothership satellite scans this beam across a constellation of femto-satellites. When the beam scans across a particular femto-satellite, the femto-satellite transmits an acknowledgement to the mothership satellite, e.g., by retroreflecting the beam or via a separate radio link. The beam can be modulated with commands for the femto-satellite, such as to make a measurement or transmit previously acquired data, as well with commands for determining the femto-satellite's location, such as a time stamp or beam pointing information.

Abstract: A debris removal satellite includes a capture device, a thruster of a chemical propulsion method, and a propellant tank to store chemical fuel. A solar array wing is operable in an orbit at an orbital altitude higher than a congested orbit region congested with satellites forming a satellite constellation. The debris removal satellite is built in advance for future use as a satellite to be launched, and when a debris intrusion alarm to give a warning about intrusion of debris into the congested orbit region is issued, propellant is loaded into the propellant tank and the debris removal satellite is launched by a rocket built in advance for future use as a launch rocket. The debris removal satellite captures capture-target debris at an orbital altitude higher than the congested orbit region, and operates a propulsion device with the capture-target debris being captured.

Abstract: A solar array wing includes an extension mast to be extended from a wound state, and a support member, around which the extension mast is wound, to support the extension mast after the extension mast is extended. The support member is made of a fiber reinforced composite material. A coefficient of linear expansion of the fiber reinforced composite material, a unit of which is for each degree Celsius, in a direction that is orthogonal to an extension direction of the extension mast, is higher than or equal to ?1×10?6 and lower than or equal to 1×10?6.

Abstract: A satellite refueling system has a refueling satellite carrying fuel, a fuel hose connected to the fuel carried by the refueling satellite, with a maneuverable end effector at a deployed end of the fuel hose, the end effector comprising a fuel supply nozzle connected to the fuel hose, a plurality of thrusters on the end effector providing thrust in a plurality of directions, an imaging device on the end effector capturing images in an immediate environment of the end effector, and computerized circuitry operating individual ones of the plurality of thrusters in response to the images captured by the imaging device. The refueling satellite deploys the fuel hose, and the computerized circuitry operates the thrusters to maneuver the end effector, bringing the fuel supply nozzle to a location to deliver fuel.

Abstract: A solar panel array system is disclosed. The system comprises a base structure, a slider assembly, one or more hinges, one or more stands, a primary set of solar panels coupled to a primary set of slider panels slidably extendable and slidably retractable along a first direction, a secondary set of solar panels coupled to a secondary set of slider panels slidably extendable and slidably retractable along a second direction opposite from the first direction, and telescoping legs.

Abstract: A method for assembling devices on a satellite structure is disclosed including attaching the devices to a first face of a plate drilled with a through-hole provided with a fluid connector, on a second face of the plate, depositing a continuous peripheral bead of polymerisable adhesive composition and depositing a plurality of discontinuous and disjointed inner beads of polymerisable adhesive composition, pressing the second face of the plate against a continuous face of the structure, performing suction so as to create a pressure differential between the plate and the structure, at least partially polymerising the polymerisable adhesive compositions so as to form adhesive joints between the plate and the structure, and interrupting the suction.

Abstract: A satellite with deployable optics is provided. The satellite has a frame, an optical axis, and a deployable optical system. The optical system has a mechanical aperture perpendicular to the optical axis, where light collected travels substantially parallel to the optical axis. The optical system has a stored configuration in which it remains within the frame and a deployed configuration in which it extends outside the frame. In some configurations, the light-collecting area of the deployed configuration is larger than the possible light-collecting area of the stored configuration. In a partially deployed configuration, all of the primary mirror segments remain substantially within the frame, and the light-collecting area is smaller than that in the deployed configuration. A method of using the satellite includes setting the satellite to the deployed configuration, detecting whether there is a deployment malfunction, and, if so, setting the satellite to a partially deployed configuration.

Abstract: A solar array structure for a spacecraft includes one or a pair of flexible blanket or other foldable solar arrays and a deployable frame structure. The deployable frame structure includes a T-shaped yoke structure, a T-shaped end structure, and one or more rigid beams, the T-shaped yoke structure connectable to the spacecraft. When deployed, the frame structure tensions the flexible blanket solar array or arrays between the T-shaped yoke structure and the T-shaped end structure. When stowed, the flexible blanket solar array or arrays are folded in an accordion manner to form a stowed pack or packs between the cross-member arms of the T-shaped yoke structure and the T-shaped end structure, also stowed in its own Z-fold arrangement. The cross-member arms of the T-shaped end structure can include a solar array that can provide power before deployment while the flexible blanket solar array is stowed.

Abstract: A satellite includes a first radiator panel, a second radiator panel, a space defined between the first radiator panel and the second radiator panel, and one or more first heat-generating components located in the space. Each of the first heat-generating components is attached to at least one of the first or second radiator panels. The satellite further includes a third radiator panel extending from the space and one or more second heat-generating components located in the space, each of the second heat-generating components is attached to the third radiator panel.

Abstract: A deployable electromagnetic radiation antenna system for extraterrestrial deployment of an electromagnetic radiation directing surface, the deployable electromagnetic radiation antenna system comprising is provided. The deployable electromagnetic radiation antenna system includes one or more deployable support structures, an electromagnetic radiation directing surface, a plurality of actuators, each coupled to one or more of the one or more deployable support structures and the electromagnetic radiation directing surface, and a satellite body, wherein the electromagnetic radiation directing surface is coupled to the satellite body by the one or more deployable support structures and the electromagnetic radiation directing surface is deployable from the satellite body by the one or more deployable support structures.

Abstract: A deployable electromagnetic radiation antenna system is provided. The deployable electromagnetic radiation antenna system includes one or more support structures, an electromagnetic radiation directing lens adapted to pass a beam of electromagnetic radiation, and a satellite body including at least one deployment mechanism, wherein the electromagnetic radiation directing lens is deployable in a first direction away from the satellite body, the electromagnetic radiation directing lens being coupled to the satellite body by the one or more support structures, wherein the at least one deployment mechanism deploys the one or more support structures to deploy the electromagnetic radiation directing lens from an undeployed state to a deployed state by at least forming a substantially planar surface of the deployed electromagnetic radiation directing lens.

Abstract: An antenna system for space applications provides a membrane antenna with one or more flexible membranes. An antenna enclosure stores the membrane antenna during stowage. One or more first deployable support structures extend along a first axis from the antenna enclosure during deployment, at least a first point of the membrane antenna being operably anchored to a point on the first deployable support structures. Deployment mechanisms are operably anchored at a junction with the first deployable support structures. The deployment mechanisms extend one or more second deployable support structures along a second axis from the first deployable support structures during deployment. At least a second point of the membrane antenna is operably anchored to a point on the second deployable support structures. Extension of the first deployable support structures and second deployable support structures unfurls the membrane antenna along both axes to overlap the junction.

Abstract: A solar array system associated with a spacecraft includes a solar array blanket portion moveable from a stowed configuration into a deployed configuration, an extendable frame coupled to the spacecraft and the blanket portion and moveable from at least a collapsed configuration into an extended configuration to move the solar array blanket portion from the stowed configuration into the deployed configuration, and at least one biasing member extending across an exterior portion of a first hinge assembly that is configured to bias at least a portion of the extendable frame into the deployed configuration.

Abstract: End effectors and systems may capture, release, and/or create a mating engagement between the end effector and a target object. Said end effectors are tolerant of positional and rotational misalignment of the target object, and include a plurality of roller wheels, one or more of which is arranged in a non-parallel plane with respect to one or more other roller wheels. A first roller wheel configured to rotate in a first plane, a second roller wheel configured to rotate in a second plane, and a third roller wheel configured to rotate in a third plane may be arranged such that the end effector is configured to engage a passive receptacle of the target object, to capture the target object. Rotating the roller wheels in the opposite direction may cause the target object to be released or launched, by urging the passive receptacle off of or away from the roller wheels.

Abstract: An additive manufacturing device for an aerospace truss includes a raw material input unit, a longitudinal beam forming unit, a longitudinal beam traction unit, a cross beam forming unit and a truss support unit. The raw material input unit stores pre-impregnated wires and pre-impregnated tapes, and a motor drives rollers to convey the pre-impregnated wires and the pre-impregnated tapes forward; the longitudinal beam forming unit is composed of three sets of forming molds, and the pre-impregnated tapes form V-shaped longitudinal beams through heating molds; a stepper motor used in the longitudinal beam traction unit drives three sets of roller traction devices through steering gears to pull formed longitudinal beams; the cross beam forming unit is composed of a motion module and a printing module, and a truss cross beam is printed through a 3D printing method of molten deposition.

Abstract: A panel deployment apparatus includes a guide provided on a first hinge axis of a main hinge; a guide provided on a second hinge axis of a side hinge; and a deployment stop arm that stops deployment of a second side panel by sequential engagement with the guide and the guide, the deployment stop arm being fixed to the second side panel, wherein the second main panel is deployed in a state of the deployment stop arm being engaged with the guide and, when the deployment stop arm is disengaged from the guide and then engaged with the guide, the first side panel is deployed in a state of the deployment stop arm being engaged with the guide and, when the deployment stop arm is disengaged from the guide, the second side panel is deployed.

Abstract: An apparatus includes an integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring. In a spacecraft application, a spacecraft appendage may be coupled with a deployment mechanism, the deployment mechanism including at least one integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring.

Abstract: A panel assembly for use with a spacecraft includes a payload, a radiator panel and a heat pipe. The payload is configured to generate waste heat during operation. The radiator panel is spaced apart from the payload and is configured to dissipate waste heat. The heat pipe is coupled to the payload and the radiator panel. The heat pipe includes a compliant portion to permit the radiator panel to move relative to the payload. Further the heat pipe is configured to transfer waste heat from the payload to the radiator panel.

Abstract: Provided is a parachute device capable of reliably opening a parachute.

Abstract: A satellite observation system and method of deploying a satellite system are disclosed. The system includes a plurality of observation satellites comprising one or more sensors, each of the plurality of observation satellites configured with at least a solar array and a mechanical stabilization element. Each of the plurality of observation satellites is constructed without positioning components. The plurality of observation satellites is positioned in a dawn/dusk sun-synchronous orbital plane about a celestial body such that the one or more observation sensors are oriented toward the celestial body.

Abstract: An emergency deorbit device provided in a satellite flying on an orbit around the earth includes a propulsion module generating thrust for separating the satellite from the orbit, a reception unit receiving a repeat signal repeatedly sent at an interval from an sending unit of a satellite bus in the satellite, a detection unit outputting a detection signal when the reception unit does not receive the repeat signal in a set time period or when the reception unit receives a deorbit command from the sending unit or from an outside of the satellite, an activation device performing, in response to the detection signal, processing for activating the propulsion module, and a power supply device provided separately from a power supply device of the satellite bus and supplying electric power to the reception unit, the detection unit, and the activation device.

Abstract: A method for estimating collision between a satellite in orbit and at least one piece of space debris having a time of closest approach to the satellite is disclosed including: obtaining the reference orbit of the satellite; determining an ephemeris of state transition data representative of the trajectory of the reference orbit; communicating the reference orbit and the ephemeris of state transition data to the satellite. The method includes the steps on board the satellite of: determining the true orbital position of the satellite; propagating the true orbit; calculating a probability of collision between the satellite and the piece of debris.

Abstract: The described features generally relate to adjusting a native antenna pattern of a satellite to adapt communications via the satellite. For example, a communications satellite may include an antenna having a feed array assembly, a reflector, and a linear actuator coupled between the feed array assembly and the reflector. The feed array assembly may have a plurality of feeds for communicating signals associated with a communications service, and the reflector may be configured to reflect the signals transmitted between the feed array assembly and one or more target devices. The linear actuator may have an adjustable length, or otherwise provide an adjustable position between the feed array assembly and the reflector. By adjusting the position of the feed array assembly relative to the reflector, the communications satellite may provide a communications service according to a plurality of native antenna patterns.

Abstract: Embodiments provide a spacecraft airlock system. Embodiments provide a method and apparatus for attaching space exposed payloads to a space station. The spacecraft airlock system provides a defined volume of space payload to the international space station. The airlock further includes a means of attaching to a space station, a closed structure attached to said means, said means of attaching is capable of robotic manipulation, and a cooling system for cooling payload components within said closed structure.

Abstract: A propeller apparatus of an air mobility is arranged such that a propeller is configured to fold when the air mobility crashes, thereby preventing additional accidents caused by fragments generated by the propeller hitting the ground, and enabling the propeller to be reused by unfolding the folded propeller after the air mobility lands.

Abstract: Embodiments in accordance with the invention address potential co-orbital threats to a spacecraft through the use of a plurality of evasion pattern maneuvers selected to prevent a rendezvous with a potential co-orbital threat from occurring within a finite horizon. Embodiments in accordance with the invention maintain separation from the potential co-orbital threat while minimizing a defending spacecraft's fuel consumption.

Abstract: A tilting device includes: an upper movable plate arranged on a side on which an object T to be tilted is arranged; an intermediate movable plate connected to the upper movable plate in a rotatable manner via an upper rotation shaft; a first actuator; a first lower movable plate connected to the intermediate movable plate in a rotatable manner via a first lower rotation shaft; a second actuator; and a first connecting member that is connected to the upper movable plate in a rotatable manner via a first connecting rotation shaft at an end portion on a side of a first direction, and that is connected to the first lower movable plate in a rotatable manner via a second connecting rotation shaft at an end portion on a side of a second direction opposite to the first direction.

Abstract: A satellite refueling system has a refueling satellite carrying fuel, a fuel hose connected to the fuel carried by the refueling satellite, with a maneuverable end effector at a deployed end of the fuel hose, the end effector comprising a fuel supply nozzle connected to the fuel hose, a plurality of thrusters on the end effector providing thrust in a plurality of directions, an imaging device on the end effector capturing images in an immediate environment of the end effector, and computerized circuitry operating individual ones of the plurality of thrusters in response to the images captured by the imaging device. The refueling satellite deploys the fuel hose, and the computerized circuitry operates the thrusters to maneuver the end effector, bringing the fuel supply nozzle to a location to deliver fuel.

Abstract: Method of installation of a PV array with planar PV modules of square/rectangular configuration, each module defining square/rectangular edge and comprising a pair of parallel end and side edges, the modules being connected along adjacent end edges and foldable relative to each other about the connected end edges between a closed condition and an open condition, whereby in the closed condition, the modules are stacked together on a movable carriage on which the modules can be transported, the modules comprising a leading module, a trailing module and two or more intermediate modules, and in the open condition, the modules are laterally displaced from the closed condition about the end edge connections to collect electromagnetic radiation, the method including securing the leading module and moving the carriage relative to the leading module so that the carriage moves away from the leading module, allowing the PV array to unfold from the carriage.

Abstract: Deployable tile aperture devices, systems, and methods are provided in accordance with various embodiments. Some embodiments include a device that may include multiple aperture tiles that may be coupled with each other such the multiple aperture tiles have a stacked stowed configuration and a flat deployed configuration. Some embodiments include one or more tension chords configured to deploy the multiple aperture tiles when tension is applied to the one or more tension chords. The flat deployed configuration may include at least one side edge portion of each aperture tile from the multiple aperture tiles making contact with another side edge portion of another aperture tile from the multiple aperture tiles. The flat deployed configuration may form one or more continuous face surfaces formed from the multiple aperture tiles. The one or more tension chords may pass through at least a portion of one or more of the multiple aperture tiles.

Abstract: A subsea vehicle capable of supporting inspection of underwater objects while underway includes a body that provides a capability to allow the subsea vehicle to submerge underwater and follow or position near an object while maintaining an orientation to the object appropriate for inspection of, and safety requirements for, the object. The vehicle includes a set of deployable, semi-rigid arms to support the movement of inspection sensor probes near or lightly touching the inspection target with the probes. A controller helps tracks the intended inspection object using various sensor inputs along with a priori knowledge of the object to drive and position the subsea vehicle such that the appropriate orientation to the inspection target is maintained.

Abstract: Methods, systems, and devices for foldable tube with unitary hinge are provided in accordance with various embodiments. Some embodiments include a device that may include one or more foldable tube sections configured to facilitate the deployment of an object with respect to an end portion of at least one of the one or more foldable tube sections; each of the one or more foldable tube sections may include one or more unitary hinge regions. The unitary hinge regions may include one or more longitudinal hinge regions configured to enable folding that is transverse to a primary axis of a respective foldable tube section. The unitary hinge regions may include a lateral hinge region configured to enable flattening that is parallel to a primary axis of a respective foldable tube section. Some embodiments may include a wire harness disposed within an interior portion of one or more foldable tube sections.

Abstract: An electric folding mechanism and a motorized vehicle thereof, the electric folding mechanism is connected between a first part and a second part capable of being folded together, which comprises a folding motor (61), a transmission shaft (62) with two ends driven by the folding motor (61), and two hinge assemblies (63) respectively disposed at each end of the transmission shaft (62); each hinge assembly (63) comprises a first hinge component (631) and a second hinge component (632) rotatably connected to the first hinge component (631); both the first hinge component (631) and the folding motor (61) are attached to the first part, and the second hinge component (632) with two ends is attached to the second part.

Abstract: In an example embodiment, an attitude, orbit, and de-orbit control system (AODCS) for a satellite is provided. In an example embodiment, the AODCS system comprises one or more selectively retractable booms. The one or more selectively retractable booms are collectively configured to provide a selectively adjustable drag during de-orbiting of a satellite over a predefined de-orbiting time.

Abstract: An assembly for deployment and retraction of a panel for outer space environment usage, includes a member secured to the panel with the member having an axis of rotation. Mass component connected to the panel, with panel wrapped about member with mass component positioned a first radial distance from the axis of rotation. Mass component has a mass per unit area greater than a mass per unit area of the panel adjacent to the mass component. With a rate of rotation of the mass component and member being same, the mass component rotates about the axis of rotation at a second radial distance from the axis of rotation, wherein the second radial distance is greater in dimension than the first radial distance. With a rate of rotation of the mass component and the member changed to be different, radial distance of the mass component to the axis of rotation decreases.

Abstract: The single axis solar tracking assembly comprises a plurality of spine sections (2) connected to one another by universal joints (10), supporting legs (3) rotatably supporting the spine sections (2) on the ground, a plurality of brackets attached to each spine section (2), a plurality of rib members (4) connected to the brackets, and solar panels (5) secured to the rib members (4). The rib members (4) are moveable between a folded position and a deployed position. An auxiliary shaft is located between adjacent universal joints (10) connected to adjacent spine sections (2). First and second crossed spine hinge axes of the universal joints are arranged so that the spine sections (2) can be fold at right angles to the auxiliary shaft and the spine sections (2) can be fold in a zig zag fashion into a compact storage position in which the solar panels are arranged in adjacent parallel planes.

Abstract: Multiple-input-multiple-output (MIMO) antenna devices and methods of using and fabricating the same are provided. A MIMO antenna device can include a plurality of substrates each having an antenna element. The substrates can be provided in connected series and can be attached to a framework. The substrates can have alternating style antenna elements, such that a first substrate can have a straight-fed patch and a second substrate adjacent to the first substrate can have a bent-fed patch and a third substrate adjacent to the second substrate, and on an opposite side of the second substrate than the first substrate is, can have a straight-fed patch and so on.

Abstract: A deployment system for solar panels has the panels spring loaded to deploy, but held folded against a satellite framework by trigger bars engaging slots or notches in hinge bodies holding the solar panels. An actuator moves the trigger bars to disengage from the hinge bodies to release the panels to deploy.

Abstract: The present disclosure relates to deployable structures and methods of use thereof. In particular, deployable structures with non-cylindrical or irregular shapes and methods of use thereof are disclosed. Non-cylindrical combustion elements can be used to rigidize such non-cylindrical or irregular shapes. The use of gaseous oxidizers along with deployable structures is also disclosed.

Abstract: The description relates to hinged devices, such as hinged computing devices. One example can include first and second portions and a hinge assembly secured to the first and second portions. The hinge assembly can define first, second, and third hinge axes. The hinge assembly can include a first timing cord that synchronizes rotation around the first and second hinge axes and a second timing cord that synchronizes rotation around the second and third hinge axes.

Abstract: Multiple-input-multiple-output (MIMO) antenna devices and methods of using and fabricating the same are provided. A MIMO antenna device can include a plurality of substrates each having an antenna element. The substrates can be provided in connected series and can be attached to a framework. The substrates can have alternating style antenna elements, such that a first substrate can have a straight-fed dipole and a second substrate adjacent to the first substrate can have a bent-fed dipole and a third substrate adjacent to the second substrate, and on an opposite side of the second substrate than the first substrate is, can have a straight-fed dipole and so on.

Abstract: A digital processing device with high input/output connectivity and modular architecture comprises a first plurality of input ports, a second plurality of output ports, and a third plurality of at least four basic elementary modules. The third plurality of the elementary modules is split up according to a partitioning of at least two sub-assemblies of module(s), at least two of which form different islets comprising at least two modules.

Abstract: The invention relates to a space wing, produced by means of a diaphragm forming a polygonal surface provided with an inflatable structure which includes ribs extending over the diaphragm along diagonals of the diaphragm and passing through a central point of the diaphragm. The inflatable structure includes at least one film strip, the perimeter of which adheres onto the diaphragm such as to form an inflatable space with the diaphragm.

Abstract: An antenna is made from strips of a shape-memory alloy or other resilient material acting as a spring with attached branches that constitute individual monopole antennas. In the folded state, the antenna looks like a strip roll and can be placed on a satellite. When in orbit, the antenna unfolds after the roll retention mechanism is released and orderly unfolds unrolling from a support frame or otherwise extends. The proposed design of monopole branches utilizes conductors of minimum length and achieves maximum directivity. Each monopole branch is connected to the signal receiver/transmitter by signal conduit elements. A system may include at least two such unfolding antennas thus achieving even greater operational effectiveness in regard to signal steerability, interference suppression and reduced moment of the satellite inertia. To prevent problems, additional measures are used that prevent unwinding of inner layers of the roll before the outer layer is extended.

Abstract: A solar collector may include a frame for supporting a plurality of photovoltaic (PV) panels. The frame may be adapted to removably attach to a base. The solar collector may include a first panel assembly, including at least one of the plurality of PV panels, pivotally attached to the frame about a first axis. The solar collector may also include a second panel assembly, including at least one of the plurality of PV panels, pivotally attached to the first panel assembly. The second panel assembly may collectively move with the first panel assembly about the first axis and to pivot relative to the frame, and to pivot about a second axis that is substantially parallel to and radially offset from the first axis, to move between a deployed position and a retracted position.

Abstract: A foil deployment mechanism comprises a first drum rotatable about a first longitudinal axis, and a second drum rotatable about a second longitudinal axis. The foil deployment mechanism further comprises a cable, the cable comprising a first section which extends from a lower part of the second drum to an upper part of the first drum, a second section which is wound around a part of the first drum facing away from the second drum, a third section which extends from a lower part of the first drum to an upper part of the second drum, and a fourth section which is wound around a part of the second drum facing away from the first drum, wherein the first section and the third section intersect each other between the first drum and the second drum when being viewed along the first longitudinal axis.

Abstract: A system, method, and computer-readable storage devices for a 6U CubeSat with a magnetometer boom. The example 6U CubeSat can include an on-board computing device connected to an electrical power system, wherein the electrical power system receives power from at least one of a battery and at least one solar panel, a first fluxgate sensor attached to an extendable boom, a release mechanism for extending the extendable boom, at least one second fluxgate sensor fixed within the satellite, an ion neutral mass spectrometer, and a relativistic electron/proton telescope. The on-board computing device can receive data from the first fluxgate sensor, the at least one second fluxgate sensor, the ion neutral mass spectrometer, and the relativistic electron/proton telescope via the bus, and can then process the data via an algorithm to deduce a geophysical signal.

Abstract: A deployable truss with modified primary orthogonal joints. The construction of these joints causes the center-hinged primary chords on opposite sides of a truss bay to fold inward in a plane orthogonal to the folding planes of the side diagonals while the two secondary chords fold in planes orthogonal to the plane of the in-folding primary chords. This provides for stiffness and stability during deploy and retract. The unique joint configuration permits the truss to deploy one bay at a time in a stable manner while having lateral bending stiffness, and the truss thus can extend and retract in a sequential manner. It can deploy integral flat panels nested between the secondary folding chords, or use cross bracing in lieu of panels. The truss can be triangular, square or rectangular in cross-section. A powered support frame may be used in conjunction with the truss.

Abstract: An assembly for positioning photovoltaic panels on a structure includes a main frame connectable to the structure, and a subassembly connectable to the main frame to support a photovoltaic panel, in which the subassembly is rotatable with respect to the main frame to place the subassembly in a deployed position. The assembly may further include a stop-arm connectable to the main frame to support the subassembly when in the deployed position.