Enclosures for facilitating activities in space, and associated systems and methods, are disclosed. A representative system includes a spacecraft having an enclosed interior volume (which can be formed by an inflatable membrane) and one or more unmanned aerial vehicles (UAVs) carried by the spacecraft and positioned to deploy into the enclosed interior volume. The system can include a remote-control system to control the one or more UAVs from a terrestrial location while the spacecraft is in space. A wireless charging system can provide electrical power to the one or more UAVs. A representative method includes configuring one or more controllers to launch a first spacecraft to a first orbit, launch a second spacecraft to a second orbit, move the first spacecraft to the second orbit, dock the first spacecraft with the second spacecraft, and broadcast an event within an interior volume of the first spacecraft to a terrestrial location.

A release apparatus includes a base member and a channel having a first portion and a second portion. A first rod positioned within the first portion includes a first end portion having a first coupling device and a second end portion coupled to a first portion of a panel assembly. A second rod positioned within the channel's second portion includes a first end portion having a second coupling device such that the second coupling device is positioned proximate to the first coupling device. The second rod includes a second end portion coupled to a second portion of the panel assembly. First and second coupling devices rotate such that a linear force is generated between the first and second rods, enabling the first rod second end portion and the second rod second end portion to simultaneously release the first and second portions of the panel assembly, respectively.

A spacecraft, reconfigurable from a launch configuration to an on-orbit configuration, includes a main body structure, a manipulator, a first deployable rigid reflector and an attachment arrangement, including at least one hold-down assembly (HDA). In the launch configuration, the HDA is in a fully engaged configuration such that the attachment arrangement mechanically attaches the first reflector with the spacecraft main body structure and prevents relative motion between the first reflector and the spacecraft main body. Reconfiguring the spacecraft from the launch configuration to the on-orbit configuration includes (i) actuating the HDA from the fully engaged configuration to a partially engaged configuration; (ii) grasping and moving the first reflector, with the manipulator, a distance in the first direction; and (iii) moving the first reflector from a first position proximate to the attachment arrangement to a second position proximate to a deployed position associated with the on-orbit configuration.

A boom deployment mechanism is disclosed. The boom deployment mechanism may include a boom, a root plug, and a locking mechanism. In some embodiments, the boom may have a proximal end and a distal end. The boom may have a deployed configuration where the boom has a tubular shape with a slit that extends along the longitudinal length of the boom from the proximal end of the boom to the distal end of the boom. The boom may have a stowed configuration where the boom is flattened and rolled. In some embodiments, the locking mechanism may be configured to secure the proximal end of the boom to the root plug when the boom is in a deployed configuration.

A spacecraft includes a first deployable module and a second deployable module and is reconfigurable from a launch configuration to an on-orbit configuration. In the launch configuration, the first deployable module is adjacent to the second deployable module. The first deployable module includes a first solar array, the first solar array being rotatable, in the on-orbit configuration, about a first axis of rotation, and the second deployable module includes a second solar array, the second solar array being rotatable, in the on-orbit configuration, about a second axis of rotation, the second axis of rotation being separated by a substantial distance from the second axis of rotation.

A landing device for a low gravity lander having a main body. The landing device comprises a number of leg-like rods attached to the main body, wherein, in a deployment position of the rods, each of the number of rods is inclined with regard to a plane of a first side surface of the main body such that the rods substantially extend in a direction of movement of the low gravity lander. Furthermore, the number of rods is made such that they bend or buckle under forces within a predetermined range by an impact due to a landing on a landing surface, thereby absorbing an impact momentum.

Some embodiments of the invention include a boom deployment system. The boom deployment system, for example, may include a housing, a spool, a first boom, and a second boom. The spool may be disposed within the housing and configured to rotate around an axis that is fixed relative to the housing. The first boom and/or the second boom may have a cylindrical shape in a deployed configuration, a flattened shape in a stowed configuration, and a slit that extends along the longitudinal length of the boom in the deployed configuration. The first boom and/or the second boom may be stowed in the stowed configuration flattened and wrapped around the spool. The first boom and/or the second boom may transition from the stowed configuration to the deployed configuration as the spool rotates around the axis.

One aspect of the present invention involves a hold-down and release mechanism (HDRM) that includes a hold-then-release mechanism configured to hold an element until the hold-then-release mechanism is commanded to release the element; and an integral sensor configured to sense an amount of a parameter of interest associated with the HDRM and to communicate the sensed amount of the parameter of interest to a receiving device. The parameter of interest comprises one of tensile preload, temperature, vibration, shock, and time from application of an HDRM firing current to HDRM release of the element.

A flexible pivot includes a first stage including a first cylinder and interface structure and a second stage including a second cylinder and interface structure in axial alignment with those of the first stage. Flexible connecting members are arranged for connecting the first and the second stages. Each flexible connecting member includes a pair of legs and a cross member joining the legs, each leg extending in a direction transverse to the axis of the cylinders, the legs being attached to the first and the second cylinders respectively. The first cylinder and the first interface structure are concentric. Flexible spokes are attached to the first cylinder by one end and to the first interface structure by the other. Each spoke extends in a direction transverse to the axis of the cylinders. Finally, the second stage includes flexible connection unit arranged to connect the second cylinder to the second interface structure.

A sparse-isogrid columnar lattice structure including rigid ring frames connected by a mirrored symmetric double helix pattern comprised of first shell hinge elements in a first helical pattern and second shell hinge elements in a second helical pattern oriented in an opposite direction to the first helical pattern and congruent thereto. The helical axes of the first and second helical patterns intersect the respective centers of the ring frames. The first and second shell hinge elements are configured to stow in a stored energy state when the ring frames are collapsed toward one another along the helical axis, and the first and second shell hinge elements are configured to release the stored energy to deploy to a restored state and extend the ring frames apart from each other along the helical axis when deployed to form a stable rigid axial column in a restored state.