Abstract: An artificial satellite comprises a satellite structure, an onboard control system including an onboard controller, and a memory system. The memory system is physically coupled to the satellite structure and independently powerable with respect to the onboard controller. The memory system is also arranged to communicatively couple with the onboard controller, and to store data which specifies one or more launch-specific parameters for configuring at least one of the satellite components. The onboard control system is adapted to operate in a transfer phase in response to the satellite separating from a payload dispenser, and to autonomously control, while operating in the transfer phase, one or more aspects of at least one satellite component at least partly based on the data, and particularly the one or more launch-specific parameters specified by or derived from the data.

Abstract: A spacecraft includes a semi-rigid solar array and a main body structure, the main body structure configured as a convex polyhedron and including an aft and a forward face disposed opposite to the aft face and at least four side faces disposed between and approximately orthogonal to the aft face and the forward face. The solar array includes a number of panels linked together with flexible couplings. In an undeployed configuration, panels of the solar array cover at least two adjacent side faces, the flexible couplings providing an articulable joint approximately aligned with a line along which the two adjacent side faces are joined and connecting a first panel of the solar array and a second panel of the solar array, the first panel being proximal to a first side face and the second panel being proximal to a second side face.

Abstract: Provided herein are various improvements to launch vehicle payload systems, such as employed to launch and deploy secondary payloads into orbit. In one example, a system includes a fairing configured to encase a payload within an envelope of a primary fairing of a launch vehicle, and a mount system configured to adapt a mounting port for the payload to a mounting port associated with the launch vehicle. The system also includes a fairing door configured to be commanded open for deployment of the payload after the primary fairing has open.

Abstract: An artificial satellite includes two pointing mechanisms. The pointing mechanisms respectively include main body side gimbals, deployed booms, thruster side gimbals, and thruster groups. The main body side gimbal connects the deployed boom to a satellite main body and adjusts a direction of the deployed boom. The thruster side gimbal connects the thruster to the deployed boom and adjusts the direction of the thruster. Each gimbal is a two-axis gimbal.

Abstract: An artificial satellite includes two pointing mechanisms. The pointing mechanisms respectively include main body side gimbals, deployed booms, thruster side gimbals, and thruster groups. The main body side gimbal connects the deployed boom to a satellite main body and adjusts a direction of the deployed boom. The thruster side gimbal connects the thruster to the deployed boom and adjusts the direction of the thruster. Each gimbal is a two-axis gimbal.

Abstract: A method and apparatus provides cube-shaped satellite primary structures, each comprised of six identical, or nearly identical, rectangular truss panels and internal struts. The struts, all adjustable in length, connect, and are directed between all cube opposite corners and all cube opposite panel centers. All struts meet at the cube center where they attach rigidly to either a block called the “nucleus fitting” or to a hollow sphere. Each strut attaches to either a ball-socket corner fitting located at the interior corner of the cube, or to a ball-socket panel center fitting located at the panel center interior to the cube.

Abstract: A networked electronic ordnance system is provided. The system includes a first plurality of pyrotechnic devices connected to a first network bus. The system further includes a first bus controller connected to the first network bus. The system further includes a second plurality of pyrotechnic devices connected to a second network bus. The system further includes a second bus controller connected to the second network bus. The system further includes a bus interface circuit connected to the first bus controller by a first electrical connection and connected to the second bus controller by a second electrical connection.

Abstract: The present invention achieves technical advantages as a vehicle cabin, comprising a floor system formed of removable and reconfigurable floor segments extending from a vehicle frame, and a ceiling system formed of removable and reconfigurable ceiling segments disposed in the ceiling of the vehicle cabin. Floor segments that can be positioned to form a floor channel, therebetween. A fixture interface adapted to releasably secure a fixture to the floor system. The fixture can be slidably repositioned within the vehicle along the floor channel. One of the ceiling segments includes one or more of a lighting module, a ventilation outlet, a ventilation inlet, a display screen, or a touch control panel. The ceiling segments can be hexagonal.

Abstract: A self-contained payload module and method of deployment of a payload includes a housing that is configured to engage a propulsive payload adapter hub, at least one deployable payload that is arranged in an interior cavity of the housing and deployable using the propulsive payload adaptor hub, a payload electronics system arranged in the interior cavity, a thermal control system in communication with the at least one payload and arranged in the housing, and at least one antenna that is arranged on the housing and configured for wideband communication.

Abstract: A system includes a satellite assembly. The assembly includes a plurality of satellite devices, where each satellite device includes a processor device, an impulse actuator, and a memory storing instructions. The instructions are executable by the processor device to access a sequence of waypoints. The sequence of waypoints identifies a first waypoint and is associated with formation of a distributed structure including the plurality of satellite devices. The instructions are further executable by the processor device to initiate movement of the satellite device toward the first waypoint at least by initiating activation of the impulse actuator.

Abstract: A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

Abstract: A robotic apparatus including a plurality of rigid body sections that move relative to each other by one or more multi-degree of freedom joints. The robotic apparatus can traverse a fixed frame by attaching its distal ends to the frame and moving the rigid body sections relative to each other.

Abstract: A method for constructing an inflatable environment on top of or beneath a surface of an extraterrestrial object includes spraying Regishell onto an airform or piping the Regishell into a sandwich membrane layer of the airform. When performing the spraying of the Regishell, the method further includes combining basalt material with the Regishell and applying the combination of the basalt material and Regishell to a reinforcement layer, the reinforcement layer being internal to the airform to strengthen the inflatable environment. When performing the piping of the Regishell into the sandwich membrane, the method further includes using the sandwich membrane layer as a permeable membrane or drilling one or more holes in the sandwich membrane layer forming vents to create the permeable membrane, and releasing the gas from the sandwich membrane layer from the vents to cure and conform the Regishell as a rigid shape and structurally sound layer.

Abstract: Small cubesat systems may be available for a lower cost, have a higher reliability, and be much simpler to use than conventional cubesats. A complete turnkey system solution may be provided, including the ground station and remote field units. The satellite, ground station, and field unit may be provided separately as kits that are ready to go out-of-the-box as soon as they arrive. This enables universities, researchers, and hobbyists to obtain and deploy their own functional satellites. Furthermore, theoretical designs and functionality may be rapidly prototyped and demonstrated, which allows for proof-of-concept without needing to build a larger, more expensive satellite system and hope that the new design or functionality works as intended.

Abstract: Systems, methods, and apparatus are disclosed for monitoring a performance of mechanical joints. Methods may include receiving, from a first plurality of sensors, a first plurality of measurements associated with a first mechanical joint included in the plurality of mechanical joints. The methods may also include generating, at a trend engine, at least one failure detection metric based on the first plurality of measurements. The failure detection metric may characterize a prediction of an operational failure associated with the first mechanical joint.

Abstract: The present invention concerns a method for making a microsatellite, comprising providing: modules of a first type configured to house electronic boards of a microsatellite; modules of a second type configured to house devices and systems of a microsatellite; and modules of a third type comprising first and second interface means configured to be coupled to a launch vehicle and to external appendages of a microsatellite, respectively; said modules of a third type being designed to cause a body of a microsatellite to have a predefined height; wherein all the modules of the first, second and third types are configured to be stacked regardless of the type. The method further comprises making a body of a microsatellite by stacking modules of different types, wherein the stacked modules include at least one module of the second type and at least one module of the third type.

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: An exemplary satellite includes a pair of payloads coupled to one another by a flexible boom, where the flexible boom is configured to enable easy manual engagement of the pair of payloads with an associated attachment hub, and also to provide a passive release force for deploying at least one of the payloads in a direction outwardly from the attachment hub. Potential energy for enabling the passive release is provided as stored strain energy in the flexible boom when flexed for attachment of the payloads to the attachment hub. The strain energy is released upon release of at least one of the payloads from the attachment hub, which release may be by way of a non-complex, non-exotic attachment mechanism. Additional payloads may be connected in series to the pair of payloads, with a flexible boom connecting adjacent payloads. The additional payloads may be released from the attachment hub via a non-complex or complex attachment mechanism.

Abstract: A payload interposer (PIP) system and its control software provide an interface between a space vehicle and a payload. The PIP board facilitates power and communications between a command and data handler (CDH) of the space vehicle and the payload. A microcontroller of the PIP board may control operation of the payload, format messages between the space vehicle and the payload, and extract data from the payload for downlink via the space vehicle.

Abstract: Disclosed is a modular, human-crewed interplanetary spacecraft that is assembled in cislunar space. It is primarily comprised of a hollowed-out asteroid; five expandable habitation modules, one of which is expanded inside the asteroid cavity; two docking and airlock nodes; two landing craft suitable for exploring celestial bodies; structural support members; truss structures; robotic arms; a propulsion module; and shielding curtains that are filled with pulverized asteroidal material and attached to the truss structure. This configuration provides substantial radiation and meteoroid shielding. Upon completion of their mission, the crew will use the robotic arms to disconnect and mate (1) the asteroid containing the control module, (2) the forward docking and airlock node, and (3) the propulsion module. This crew-return vehicle will return to cislunar space.

Abstract: A payload dispenser and a method for populating same with satellites is disclosed. The payload dispenser comprises a shell. A plurality of rail assemblies are each configured to receive plural satellites. The satellite laden rail assemblies are then coupled to the shell.

Abstract: A networked electronic ordnance system is provided. The system includes a first plurality of pyrotechnic devices connected to a first network bus. The system further includes a first bus controller connected to the first network bus. The system further includes a second plurality of pyrotechnic devices connected to a second network bus. The system further includes a second bus controller connected to the second network bus. The system further includes a bus interface circuit connected to the first bus controller by a first electrical connection and connected to the second bus controller by a second electrical connection.

Abstract: A habitation module that provides an artificial gravity environment. In one embodiment, the habitation module includes a rotating structure configured to rotate about an axis in relation to a stationary structure, and attachment assemblies configured to rotatably couple the rotating structure to the stationary structure. The attachment assemblies comprise bearings installed coaxially between a cylindrical housing of the rotating structure and a cylindrical body member of the stationary structure, and dynamic radial seals installed coaxially between the cylindrical housing and the cylindrical body member. The attachment assemblies further comprise a static seal assembly that includes an end cap movable between a locked position and an unlocked position.

Abstract: A habitation module that provides an artificial gravity environment. In one embodiment, the habitation module includes a stationary structure and a rotating structure that rotates about an axis in relation to the stationary structure. The rotating structure includes a cylindrical body that is substantially hollow and closed at both ends, where the ends represent gravity chambers of the habitation module. There are center openings in the cylindrical body that are coaxially aligned on opposing sides of the cylindrical body along an axis of rotation. First and second hub members are permanently affixed to the cylindrical body on the opposing sides of the cylindrical body about the center openings. Radial seals that span gaps between the rotating structure and the stationary structure.

Abstract: Disclosed is a satellite payload device, including a plurality of sensors configured to detect images of a satellite, a video distribution and storage unit configured to collect, compress, store, and transmit the images, and a control computer configured to request a portion of the plurality of sensors from which to receive sensor data that is routed to the video distribution and storage unit.

Abstract: A satellite comprises at least one first communication module, a repeater module, one or more antennas, at least one propulsion module and an avionics module, the repeater module comprising: a north first panel when the satellite is in an orbital position, an east second panel, a south third panel and a west fourth panel arranged respectively opposite the north first panel and the east second panel, the north panel and the south panel having widths XmcR in the east-west direction, the east panel and the west panel having a width or YmcR in the north-south direction; characterized in that: the value of the ratio XmcR/YmcR for the repeater module lies within the interval [0.84; 1.12], and the repeater module consists of a number of floors, and the value of Xmck varies as a function of the floor k and of its position in the repeater module.

Abstract: A networked electronic ordnance system is provided. The system includes a first plurality of pyrotechnic devices connected to a first network bus. The system further includes a first bus controller connected to the first network bus. The system further includes a second plurality of pyrotechnic devices connected to a second network bus. The system further includes a second bus controller connected to the second network bus. The system further includes a bus interface circuit connected to the first bus controller by a first electrical connection and connected to the second bus controller by a second electrical connection.

Abstract: A side-by-side dual-launch spacecraft arrangement is provided. The arrangement may include a dual-launch adaptor, a first spacecraft, and a second spacecraft. The first spacecraft and the second spacecraft may be mounted on the dual-launch adaptor and may be arranged side by side on the dual-launch adaptor. An aspect ratio of each of the first and second spacecraft may be within a range of 0.55 and 0.8.

Abstract: Various material compositions of a heat shield for a spacecraft are described. The heat shield can be formed by multi-dimensional weaver or three-dimensional (3-D) printer. Furthermore, the heat shield can be configured with a superconducting coil.

Abstract: The invention relates to a satellite having a structure that includes a circular launching interface ring and at least two main planar apparatus-holding walls that are substantially parallel to the axis of the ring and rigidly connected to one another and to the ring. Each wall directly bears on the ring by a base of the wall, and each wall is rigidly and directly attached to the ring by at least one somewhat point-by-point connection on at least one point or area of tangency or intersection with the base of the wall and the ring.

Abstract: A system includes a gasification feed injector. The gasification feed injector includes a tip portion that has an outer wall and a feed outlet, and an injection axis of the gasification feed injector extends through the feed outlet of the tip portion. The gasification feed injector also has at least one fastener extending radially from the outer wall relative to an injection axis of the gasification feed injector through the feed outlet of the tip portion. An attachment location of the at least one fastener is located away from the feed outlet of the tip portion. The system also includes a heat shield, which includes a tile radially disposed on the at least one fastener and radially secured to the outer wall via fastening the tile to the at least one fastener. In addition, the tile wraps around a distal end portion of the tip portion.

Abstract: A gravity acceleration station for producing gravity acceleration and creating conditions for living under a permanent effect of gravity acceleration more than 1 g for prolonged periods of time. The station comprises a base and a hollow torus, rotating around a central vertical axis. A support of the station and motors for rotation of the station are located peripherally, along with the perimeter of the torus. That feature allows variable size of the station with diameter more than 100 meters, larger area for location of objects, and gradual increase of gravity acceleration from the center of the station along the radius. Due to a mechanism for altering the angle of deviation of the premises of the station, the value of the net acceleration can be changed according to the needs while keeping direction perpendicular to the floor of the premises. The station can be located on the ground or underground.

Abstract: An inflatable module space station is disclosed. The module has an avionics architecture designed for the module serving as a space station. The module also has at least one hundred and eighty cubic meters of internal habitable volume and at least one window. The module is capable of supporting at least two humans in a zero gravity environment for at least one month. There is also a propulsion bus designed for a space station and attached to the module. Furthermore, there is a docking mechanism designed for a space station and attached to the module. The module is capable of orbiting a mass like a planet or moon. The space station can be one or multiple expandable modules.

Abstract: A system and method for assembling a spacecraft in orbit using orbiting modules. Each module has a function such as fuel, transport, communication and payload. A command and control system and logic assembles the modules for missions. After use the modules may be disassembled and parked in orbit. The assembly of modules for a mission is controlled by a logic that assesses the mission requirement, module status and capability and matches resources. The referenced command and control system and logic is used to maneuver vehicles and modules and controls missions. Communications between and among modules and signal sources are facilitated by a language protocol that has a library of commands and responses accessible by signals using divergent communications languages. The protocol also converts common programming language to a language compatible for use by a recipient module, logic or communication satellite or ground station.

Abstract: Payload adapters for launch vehicles include a ring structure having an opening formed therein and at least a portion of an antenna assembly coupled to a circular sidewall of the ring structure at least partially within the opening of the ring structure. Satellite assemblies may include a payload adapter comprising a ring structure having a circular opening formed therein and an antenna assembly including a parabolic reflective dish formed within the circular opening. Launch stack systems may include a primary payload and a secondary payload adapter including an antenna dish integrally formed with a ring structure of the secondary payload adapter. Methods of forming a satellite assembly include positioning at least a portion of the antenna assembly within an opening formed in a ring-shaped payload adapter for a launch vehicle and coupling the at least a portion of the antenna assembly to the ring-shaped payload adapter.

Abstract: A spacecraft deployment mechanism employing a pair of bi-stable tape springs attached to a cylinder such that rotation of the cylinder extends or retracts the tape springs in opposite directions. The tape springs are fabricated such that they are stable when wrapped around a cylinder (stowed configuration) or when deployed and have a higher energy state when stowed than when deployed. A plurality of bi-stable tape spring mechanisms may be combined to create a deployment device that does not impart rotational torques to the spacecraft during deployment. Deployment may be accomplished by release of the stored energy in the coiled or packaged state progressing to the lower energy deployed state. Alternatively, a motor may be employed to rotate the cylinders for deploy or retrace the tape springs.

Abstract: A spacecraft carbon nanotube shield is disclosed. Shield segments are produced in a facility in space. The segments are transported from the facility to a vicinity of a spacecraft hull. The segments are assembled over the hull to substantially cover an area of the hull.

Abstract: An apparatus and method are disclosed for dual Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) port small satellite designs. The apparatus and method provide payload volume for larger satellites. In one or more embodiments, the apparatus and method include a plurality of small satellite components, a payload adaptor ring, and at least one pivoting hinge system. In at least one embodiment, the plurality of small satellite components includes at least one payload, one bus, and/or one solar panel. The small satellite components are mounted on the payload adaptor ring. At least one pivoting hinge system connects together at least two of the small satellite components. Upon deployment of the small satellite components from the payload adaptor ring, at least one pivoting hinge system combines together at least two small satellite components, thereby creating at least one single larger satellite.

Abstract: A polygonal support structure is disclosed herein. The polygonal support structure has a plurality of rigid segments configured to be joined at longitudinal ends by a joint affixed to at least one end of each rigid segment. Each joint is configured to rigidly, rotatingly or hingedly join each rigid segment to an adjacent rigid segment to form the polygonal support structure. A plurality of flexible members are configured and disposed to provide tension between each joint and joints spaced two segments therefrom, in each direction, about the polygonal support structure. One or more joints may be configured to join and un-join two adjacent rigid segments and each rigid segment and each joint may be configured to enable stacking of rigid segment into a compact configuration for transport. In at least one aspect, a polygonal support structure may be configured to support assemblies in a substantially weightless environment.

Abstract: A high capacity satellite having multiple aspect ratios is configured to have at least two modules. Each module includes a first panel and a third panel facing, in an on-orbit configuration, respectively, toward north and south and having a width in an east-west direction (EW width); and a second and a fourth panel oriented, in the on-orbit configuration, respectively, east and west and having a width in the north-south direction (NS width). Each module has a respective aspect ratio of EW width to NS width, a first module is configured with a first aspect ratio, a second module is configured with a second aspect ratio; and the second aspect ratio is substantially larger than the first aspect ratio. At least one antenna reflector is disposed, during launch, proximate to at least one of the first and third panel of the second module.

Abstract: Solar power satellite system for transmitting microwave energy to the earth and a method of arranging the solar power satellite system about the sun for same. The solar power satellite system comprises a space-based power generation unit disposed in a planetary orbit about the sun. A photovoltaic cell on the space-based power generation unit collects solar energy that is then converted to microwave energy to be beamed to the earth. A ground-based rectenna receives the microwave energy and converts the microwave energy to electricity that is transmitted to an end user. The solar power satellite system and method provides electrical power on earth day or night and regardless of atmospheric conditions. Also, surface area of the solar panel on the space-based power generation unit orbiting about the sun is much less than the surface area required of a ground-based solar panel or a solar panel in earth orbit.

Abstract: An air barrier having a scuff coating is used in connection with the core of a spacecraft to form an inflatable, or expandable, human habitat module. The air barrier is comprised of a number of segments joined together to form a enclosure that has at least one opening. The joined segments are generally gas impermeable, flexible, and have a scuff coating on both sides of the segments.

Abstract: An inflatable airlock for use with a spacecraft is disclosed. The airlock has a substantially cylindrical shaped layered shell comprised of an outer meteoroid shield layer, a restraint layer under the meteoroid shield, and an air barrier layer under the restraint layer. There is a door arrangement attached to the airlock and the airlock is adapted to being attached to a spacecraft bulkhead that also includes a door. When inflated, a person can open the spacecraft bulkhead door and pass through from the spacecraft into the airlock or vice versa. When the atmosphere is removed from the airlock, a person can pass from the airlock into space or vice versa.

Abstract: A spacecraft architecture and accompanying standard allows for the creation of a spacecraft using an assortment of modules that comply with the standard. The standard preferably includes both mechanical and electrical compatibility criteria. To assure physical/mechanical compatibility, the structure of each module is constrained to be compatible with any other compatible module. To minimize the interference among modules, the extent of each module in select dimensions is also constrained. To assure functional compatibility, a common communication format is used to interface with each module, and each public-function module is configured to respond to requests for function capabilities that it can provide to other functions. Each module is preferably designed to provide structural support to the assemblage of modules, and an anchor module is provided or defined for supporting the entire assemblage and coupling the assemblage to other structures, such as a launch vehicle.

Abstract: Structures and methods are disclosed regarding deployable structures with expandable longerons adjustably coupled with supporting structures such that an angle between the supporting structures can be adjusted. Such structures can include and/or be used for solar arrays, bridges, support structures, 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 longerons that can have two resting states: deployed and rolled.

Abstract: A plurality of sufficient closely-packed functional units which are interconnected to form a non-planar payload module assembly 40. Advantageously, the payload modular structure 40 is compact with few piece parts, has low output losses, and is robust enough for supporting complete payloads with or without reflectors. The payload modular structure has utility in various space-based applications as well as in various terrestrial applications.

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: 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 connecting node comprises a polyhedral structure comprising a plurality of panels joined together at its side edges to form a spherical approximation, wherein at least one of the plurality of panels comprises a faceted surface being constructed with a passage for integrating with one of a plurality of elements comprising a docking port, a hatch, and a window that is attached to the connecting node. A method for manufacturing a connecting node comprises the steps of providing a plurality of panels, connecting the plurality of panels to form a spherical approximation, wherein each edge of each panel of the plurality is joined to another edge of another panel, and constructing at least one of the plurality of panels to include a passage for integrating at least one of a plurality of elements that may be attached to the connecting node.

Abstract: A deployable structural assembly is provided along with associated deployment mechanisms and associated methods of forming and deploying the structural assembly. In one exemplary embodiment, the structural assembly includes a plurality of structural side elements formed from furlable truss structures. The truss structures may include spaced apart longitudinal members, members extending substantially transversely between the longitudinal members, and diagonal members extending between the longitudinal members at an acute angle relative thereto. A plurality of interlocking elements are located along the edges of the longitudinal members and are configured to engage interlocking elements of an adjacent structural side element. In one embodiment the truss structures may be formed to include at least one material layer having the longitudinal members, the cross-members and the diagonals formed as an integral structure.