A deployable thin membrane apparatus for use with a spacecraft is provided. The apparatus includes a flexible membrane structure and a deployment mechanism for transitioning the membrane structure from an undeployed state towards a deployed state in which the membrane can perform a function needed by a spacecraft. In one embodiment, the deployment mechanism includes a plurality of pantographs that each engage the membrane structure, a rotatable disk structure that engages and coordinates the transition of the pantographs from an undeployed state towards a deployed state, and a spring system for providing energy that is used to rotate the disk structure.

Radiation shielding technologies for radioisotope battery-powered vehicles to protect computer systems and humans and increase the energy efficiency, mass efficiency, and duration capability of the vehicle during operation. A radioisotope power system includes a radioisotope power unit that emits a plurality of radiation particles. The radioisotope power system further includes a radiation shield configured to block a first radiation particle of the plurality of radiation particles. The radioisotope power system further includes a decoupling device configured to decouple the radiation shield from a vehicle. The radioisotope power unit can include one or more radioisotopes for power, propulsion, or both power and propulsion of the vehicle. The one or more radioisotopes can include an alpha emitting isotope, a beta emitting isotope, a gamma emitting isotope, or a combination thereof. The one or more radioisotopes can be for heat generation. The vehicle can be a spacecraft or an aircraft.

A deployment system pivotably deploys panel elements between a stowed configuration and a deployed configuration. Actuator(s) associated with each pair of panel elements include a first bracket mountable to one panel element, a second bracket mountable to the other panel element, and a linear motion to rotary motion converter (LMRMC) for pivoting the first bracket with respect to the second bracket responsive to a predetermined datum linear displacement being applied to the LMRMC. An actuation cable, coupled to each actuator, can be displaced linearly with respect thereto between a first position, corresponding to the stowed configuration, and a second position, corresponding to the deployed configuration, responsive to operation of the drive unit, such as to apply at least a corresponding datum linear displacement to the respective LMRMC of each actuator. The drive unit is configured for selectively displacing the actuation cable between the first position and the second position.

A spacecraft with retractable solar sails is a spacecraft that is designed as a semi-autonomous spacecraft that can maneuver and propel using radiation pressure. The spacecraft includes a spherical hull and several propulsion mechanisms. The spherical hull houses the propulsion mechanisms, the payload, and other systems that enable the semi-autonomous operation of the spacecraft. The spherical hull also enables the semi-autonomous operation of the propulsion mechanisms. The propulsion mechanisms enable the propulsion and maneuvering of the spacecraft using radiation pressure in deep space. Each of the propulsion mechanisms includes a solar sail assembly, a sail receptacle, a pneumatic actuator, and a quantity of fluid. The solar sail assembly supports several retractable solar sails. The sail receptacle allows the retraction of the solar sail assembly into the spherical hull. The pneumatic actuator allows the deployment of the solar sail assembly. The quantity of fluid enables the operation of the pneumatic actuator.

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.

The present invention relates to articulating spacecraft chassis and methods of making and using same. The present invention relates to spacecraft chassis and methods of making and using same. Such spacecraft chassis have a dynamic movement capability that allows the spacecraft to alter its structure while still maintaining industry volumetric launch standards. This capability increases opens up a wide range of achievable volumetric states and increases the ability to meet mission requirements by introducing a new tunable parameter. In addition, the judicious selection of certain dynamic movement parameters can result increased payload capabilities and improved maneuverability.

An on-orbit assembly auxiliary device for space structures includes a plurality of first link rods, one end of the first link rod is hinged with a plurality of second link rods, and an opening assembly for expanding the second link rod is provided on the second link rod; the opening assembly includes a plurality of first hinge rods and third hinge rods, and each of the second link rods is provided with a chute; a first slider and a second slider are slidably provided in each chute; every two adjacent first sliders are rotatably connected through the first hinge rod; every two adjacent second sliders are connected through the third hinge rod; and the first slider close to the first link rod is rotatably connected to the first link rod through the first hinge rod. A rotating assembly for providing rotational force is provided on the top second link rod.

A deployable solar array is mounted on a rocket in a stowed position and deployed in space. The deployable solar array includes a plurality of frame body units. Each frame body unit includes a frame body and a film. The frame body defines a frame shape. The film is attached to the frame body. The film appears as a mounting surface in an opening that is formed by the frame shape defined by the frame body, with a plurality of solar cells mounted on the mounting surface. According to the deployable solar array, effects that are advantageous for low cost, space saving, and mass productivity are obtained, compared to the widely used conventional rigid panel type solar arrays.

A system, including: a satellite; a first solar array including a first solar panel; a second solar array including a second solar panel; a boom having a first end and a second end opposite the first end, where the first end connects to the satellite; and a bracket assembly, where the bracket assembly has a first, second, and third brackets, where the first bracket connects the first solar array to the third bracket, where the second bracket connects the second solar array to the third bracket, and where the third bracket connects the first bracket and the second bracket to the second end of the boom; where the bracket assembly is configured to reorient the first solar array and the second solar array between a stowed position to a deployed position, and where in the deployed position, the first and second solar arrays are oriented at a predetermined angle.