A hinge includes internal on-axis stopping mechanisms that cause the hinge to shear and break at an on-axis weakened region of the hinge when rotation of the hinge reaches a predetermined angle with a specified torsional load. The on-axis configuration is compact, has minimal impact on the outer mold line (OML) of the object to which it is mounted both pre and post detachment and allows for accurate tailoring of the torsional load that will detach the hinge.

A device for docking with a satellite including: a mobile, satellite receiving platform (10), suitable for resting against a protruding element of the satellite, a device for capturing the protruding element, and a device (20) for damping and positioning the receiving platform, including: a set (21) of link arms (210) connecting the satellite receiving platform to a member that is fixed relative to a spacecraft carrying the docking device, the set of link arms being suitable for enabling the platform to move in six degrees of freedom, and a set of magnetic dampers (25), suitable for damping the contact between the satellite and the mobile receiving platform in proportion to the relative speed between the satellite and the mobile receiving platform. A method for docking and undocking this device to/from a satellite.

The present technology is directed to thrusting rails for launch vehicles, and associated systems and methods. Certain embodiments of the thrusting rails can include a first rail housing portion having a cavity and a second rail housing portion having a projection positioned at least partially within the cavity. The rails can include a bellows positioned within the cavity and an elongated tube positioned within the bellows. The elongated tube can include a vent opening in a lateral wall of the elongated tube. A shield can be positioned between the vent opening at the bellows and a sleeve can be positioned within the elongated tube. The sleeve can be constructed from a fibrous material and positioned to retain an ordnance within the elongated tube.

A spacecraft is disclosed, comprising a deployable spacecraft body (110) comprising a plurality of sub-systems (321-324) for controlling operations of the spacecraft, and a plurality of panels (101, 102) and a plurality of hinges (112-115) each connecting adjacent ones of the plurality of panels, the hinges being arranged to permit the plurality of panels to be folded into a stowed configuration and unfolded into a deployed configuration, wherein the plurality of sub-systems are fixed to and supported by one or more of the plurality of panels. By forming the body of the spacecraft from a deployable structure, the overall size of the spacecraft can be significantly reduced in the stowed configuration. In some embodiments, a plurality of the spacecraft in the stowed configuration can be combined into a modular spacecraft assembly prior to launch, with data and power connections between the plurality of stowed spacecraft being used to transfer power from, and data to, a payload monitoring unit on the launch vehicle.

A coupling/uncoupling device includes a columnar member configured by a divided pair of semicircular members, and a pair of coupling members forming a circular holding section that holds the columnar member. The circular holding section includes a circular inner surface section with which an outer circumferential surface of the columnar member is in slide contact, the circular inner surface section being formed from inner circumferential surfaces of a semicircular section in one coupling member and a semicircular section in the other coupling member. The columnar member is configured to be rotatable in both clockwise and counterclockwise directions along the circular inner surface section about a reference position where a contact surface of the pair of coupling members and a dividing surface of the columnar member are flush with each other.

A coupling/uncoupling device includes a columnar member configured by a divided pair of semicircular members, and a pair of coupling members forming a circular holding section that holds the columnar member. The circular holding section includes a circular inner surface section with which an outer circumferential surface of the columnar member is in slide contact, the circular inner surface section being formed from inner circumferential surfaces of a semicircular section in one coupling member and a semicircular section in the other coupling member. The columnar member is configured to be rotatable in both clockwise and counterclockwise directions along the circular inner surface section about a reference position where a contact surface of the pair of coupling members and a dividing surface of the columnar member are flush with each other.

A reinforcing element for a structural profile, in particular for a round, oval or elliptical structural tube. The reinforcing element comprises: a fiber structure which has a hollow-cylindrical, helically wound mesh of fiber strands and forms an inner shell surface formed to receive the structural profile; and a matrix material into which the fiber strands are respectively embedded and which is formed to be shrinkable by heating so that the fiber structure can be fastened to the structural profile with the inner shell surface by heating the matrix material. Also provided are a structural arrangement with such a reinforcing element, an aircraft or spacecraft with such a structural arrangement, as well as a method for producing such a structural arrangement.

Omni-directional, extensible grasp mechanisms are disclosed. Such grasp mechanisms may be used as a robotic end effector for docking, grasping, and manipulating space structures, or to interconnect other structures or vehicles. Novel interconnected lattice structures may enable large arrays to be assembled. The grasp mechanisms may be used to create structures from parallel docking linkages. This may enable reconfiguration of multiple docked space vehicles and/or structures without the use of propellant. The grasp mechanisms have the ability to make and break connections multiple times, enabling a nondestructive and reversible docking process.

A carrier of a spacecraft can include multiple latch assemblies that are linked together and held in place by one notched bolt or main shaft, which is out of the main load path. The latch assemblies secure the payload (e.g., CubeSat device) within a carrier by interfacing with pin assemblies on the payload. To deploy the payload, the shape memory alloy actuator is fired which causes a series of springs and preload forces to rotate latches of the latch assemblies out of the path of the pin assemblies. Once the latches are out of the way, the payload is deployed by a spring-loaded pusher assembly and guided through deployment using rollers. Each latch assembly is preloaded in tension using a corresponding preload lug of a receiver assembly.

A carrier of a spacecraft can include multiple latch assemblies that are linked together and held in place by one notched bolt or main shaft, which is out of the main load path. The latch assemblies secure the payload (e.g., CubeSat device) within a carrier by interfacing with pin assemblies on the payload. To deploy the payload, the shape memory alloy actuator is fired which causes a series of springs and preload forces to rotate latches of the latch assemblies out of the path of the pin assemblies. Once the latches are out of the way, the payload is deployed by a spring-loaded pusher assembly and guided through deployment using rollers. Each latch assembly is preloaded in tension using a corresponding preload lug of a receiver assembly.