A solar array structure for a spacecraft is based on a modular approach, allowing for arrays to be designed, and designed to be modified, and manufactured in reduced time and with reduced cost. The embodiments for the solar array are formed of multiple copies of a bay of a multiple strings of solar array cells mounted on semi-rigid face-sheet structural elements. The bays are then placed into frame structures made of tubes connected by nodes to provide an easily scalable, configurable, and producible solar array wing structure. This allows for rapid turnaround of program specific designs and proposal iterations that is quickly adaptable to new/future PhotoVoltaic (PV) technologies and that can create uniquely shaped (i.e., not rectangular) arrays, allowing for mass production with simple mass producible building blocks.

A retaining and releasing device for spaceships retains and releases a body to/from a support structure. A releasing bolt keeps the releasable body and the support structure connected in a position for locking the retaining and releasing device. A locking device for locking the releasing bolt in the support structure has a case and a segmented ring. The case is movable axially relative to the releasing bolt and includes an end with a through-slot. A releasing system includes an actuator, a rod having the same orientation as the releasing bolt, with a locking pin, cooperating with the slot, and a compression spring. The actuator locks the rod in a position for locking the retaining and releasing device. The compression spring is in contact with the case of the locking device and disposed around the rod to rotate the rod. The device includes a system for removing the releasing bolt.

Artificial satellites are dependent on solar arrays to produce the power needed to support the functional components of the satellite. The present invention is directed to a hinge-locking mechanism designed to maximize the photovoltaic real estate of a solar array of an artificial satellite. The hinge-locking mechanism facilitates the deployment of the solar array upon the artificial satellite entering orbit. The hinge-locking mechanism utilizes a tapered pin with an internal spring resting against an asymmetrically oval shaped cam of a common pivot point. The hinge-lock trades kinetic energy of a solar panel for spring compression to permanently lock the solar array in place upon deployment while reducing the shock load and maximizing deployed stiffness.

Artificial satellites are dependent on solar arrays to produce the power needed to support the functional components of the satellite. The present invention is directed to a hinge-locking mechanism designed to maximize the photovoltaic real estate of a solar array of an artificial satellite. The hinge-locking mechanism facilitates the deployment of the solar array upon the artificial satellite entering orbit. The hinge-locking mechanism utilizes a tapered pin with an internal spring resting against an asymmetrically oval shaped cam of a common pivot point. The hinge-lock trades kinetic energy of a solar panel for spring compression to permanently lock the solar array in place upon deployment while reducing the shock load and maximizing deployed stiffness.

A multiple hold down and release device for restraining and releasing a payload from a spacecraft performs methods to release a payload from a spacecraft and install a multiple hold down and release device. At a periphery, the device includes hold-down assemblies. Each assembly has a hold-down bracket configured to attach to the payload and a hold-down support, both having conical mating surfaces. Each hold-down support includes a torsion spring(s) around an axis. In pairs of connecting levers, each lever includes a first end coupling to a corresponding hold-down support such that each corresponding lever is articulated with the corresponding hold-down support through an axis; and a second end opposite the first end. Restraint cables join second ends of each pair of connecting levers and are arranged radially with a central area where the restraint cables cross. A central release actuator is configured for cutting the restraint cables.

Devices for the assembly of structures are described. More specifically, devices designed for the assembly of structures in space are described, with a particular focus on antenna structures and methods for their assembly and deployment in space. Even more specifically, it addresses the construction of mesh reflector antennas in space. The invention includes devices, systems, and methods designed to facilitate the creation and deployment of these antennas in a space environment by leveraging advanced assembly techniques.