A Hold Down Release Mechanism, HDRM, interface for attachment of a spacecraft to an adjacent structure of a launch vehicle or another spacecraft, wherein the HDRM interface is configured for forming part of a single or multi-point releasable attachment of the spacecraft to said adjacent structure. The HDRM interface includes first and second connector parts, wherein one of the first and second connector parts is fastened to said adjacent structure, and other is fastened to the spacecraft; wherein the first connector part has a tapered projection with a non-circular external surface; wherein the second connector part has a tapered recess with a non-circular interior surface configured for form-lockingly receiving the tapered projection, for enabling transfer of torsion and shear load between the first and second connector parts, when the tapered projection is inserted in the tapered recess and the first and second connector parts are pressed together.

The system (S) to separate at least two mechanical elements (E1, E2), comprises a holding device comprising a connecting screw integral with a mechanical element (E1) and held by a nut segmented in portions (4n) arranged between the connecting screw and an outer envelope integral with the other mechanical element (E2), the system (S) comprising a force generator generating a thrust force (F) in a longitudinal direction (X-X) and comprising an energy accumulator comprising a sealed chamber with a colloid of a porous matrix and liquid, the chamber deforming in the longitudinal direction (X-X) to adapt to the change from a compressed to a decompressed state in order to generate the force (F), and an activatable actuating element to, either apply a predetermined pressure to the chamber to keep the colloid in the compressed state, or not apply the predetermined pressure to allow the change in state of the colloid.

Technology is disclosed herein for a spacecraft launch restraint and dispensing structure. The dispensing structure has a number of trusses and a central structure. When the trusses are in a support position, each spacecraft may be supported at one point by the central structure and at two points by one or more of the trusses. Therefore, each spacecraft may be supported at three points, thereby providing a stable support for each spacecraft. The spacecrafts do not touch each other and do not bear the weight of other spacecrafts. In a deployment position, the trusses extend away from the satellites and do not support the satellites; however, the satellites initially remain connected to the central structure. In the deployment position, the trusses are out of an ejection path such that the satellites can be ejected in a desired sequence from the central structure.

Stacked satellite dispensing systems are described herein. The disclosed systems have diagonal struts that stabilize satellite stacks horizontally and vertically without adding performance-reducing mass. The diagonal struts increase the number of bracing points and improve stability. The improved stability can allow for the satellite stack to be made heavier and taller, such as by having more satellites than a dispensing system with vertical struts. The diagonal struts, which provide the improved stability, can also allow for sub-stacks to be used. The sub-stacks include batches of satellites retained by the stacked satellite dispensing system. Therefore, the stacked satellite dispending system can release single satellites batches at once, rather than all the satellites at once.

Provided herein are various enhancements to spacecraft or other vehicles, including spacecraft docking mechanisms and vehicle mating systems. In one example, a vehicle mating mechanism includes a latch assembly of a vehicle having soft capture elements and hard capture elements. The soft capture elements accept a bar element of a mating vehicle and retain the bar element to within an envelope that provides a soft capture with the mating vehicle. The hard capture elements move within the envelope to engage the bar element and draw the bar element toward the vehicle to provide a hard capture with the mating vehicle. The vehicle mating mechanism can also include a cup-cone interface element of the vehicle that mates with a cup-cone interface element of the mating vehicle during the hard capture to establish an alignment between the vehicle and the mating vehicle.

A device for mounting a load to a carrier is described. The device includes a first support configured for fixation to the carrier; a second support configured for fixation to the load. The first and second supports are spaced away from each other viewed along a main load bearing axis of the device. The device further includes an outer shell extending along the main load bearing axis, and an inner member that is within the outer shell. The device includes a damping material connecting the inner member to at least the outer shell, wherein one of the outer shell and the inner member connects the first support and the second support to each other.

A representative spacecraft system includes a launch vehicle elongated along a launch vehicle axis and having at least one stage carrying a corresponding rocket engine. The representative system further includes an annular support structure carried by the at least one stage and positioned to support a cargo spacecraft having a service module and a cargo module. The cargo module of the cargo spacecraft is positioned along the launch vehicle axis in a direction distal from the support structure, and at least a portion of the service module of the cargo spacecraft positioned within an annulus of the support structure

A space vehicle electromechanical system may employ an architecture that enables convenient and practical testing, reset, and retesting of solar panel and antenna deployment on the ground. A helical antenna winding fixture may facilitate winding and binding of the helical antenna.

Disclosed is artificial satellite including: a mounting structure supporting equipment-bearing walls; a launcher-adapter rigidly connected to the mounting structure; a first radiator; and at least one first system for transporting heat by a fluid, including at least one duct having a first heat-exchange section and a second heat-exchange section, the second heat-exchange section being capable of being in thermal contact with the first radiator. The first heat-exchange section is in thermal contact with at least one portion of the launcher-adapter. Also disclosed is a method for filling a tank of propellant gas of the artificial satellite.

A habitation module that provides an artificial gravity environment. In one embodiment, the habitation module includes a stationary structure including a hub having a plurality of portals spaced radially around an outer cylindrical surface of the hub, and a rotating structure that attaches to the outer cylindrical surface of the hub using rotatable attachment members to rotate about an axis in relation to the hub. The rotating structure includes a platform that attaches to the rotatable attachment members and is configured to revolve around the outer cylindrical surface of the hub on the rotatable attachment members. The rotating structure also includes a gravity chamber that attaches to the platform, and projects radially from the axis. A drive mechanism is configured to rotate the rotating structure about the axis in relation to the hub to simulate a gravitational force within the gravity chamber.