Systems for insulating a space vehicle or space borne container from an external environment as well as cryogens from heat sources. Such systems also protect the vehicle from the high dynamic pressures, the high heat loads encountered in atmospheric flight, and provide storage capability that strongly limits, or effectively eliminates, cryogenic boil-off losses once in space. Such systems include an expandable structure having a plurality of contiguously adjacent expandable layers. The layers are connected by a plurality of tension connectors between successive layers. For launch and flight the layers can be restrained in a collapsed position. Whereupon exiting a free stream environment, the layers are expanded where they can lock into place or otherwise remain in an expanded state. The expansion creates separation between the layers with minimal conduction paths providing near theoretically perfect multi-layer insulation and extremely effective debris protection.

To reduce space debris and decrease risks for future space flights and currently operating satellites, NASA requires all satellites to have an end of life deorbiting plan to prevent satellites from having long and indefinite orbit lifespan. Accordingly, disclosed herein are systems and methods for deploying a deorbiting drag device to dramatically decrease the orbit lifespan of satellites. One of the methods comprises: providing power, using a photovoltaic panel, to a central processing unit (CPU) of the satellite; determining, using a health sensor, a health status of the satellite by monitoring activities of the CPU; and releasing a deorbiting drag device based on the health status by diverting power from the photovoltaic panel to a release actuator.

Exemplary embodiments described herein include innovative engagement devices. Exemplary engagement devices may include on or more tape spring systems. The tape spring system may include a continuous or segmented bi-stable tape spring. The tape spring can be stowed in a rolled up configuration, extended to a deployed configuration, and then triggered to return to a retracted configuration.

The present disclosure relates to deployable structures and methods of use thereof. In particular, deployable structures with non-cylindrical or irregular shapes and methods of use thereof are disclosed. Non-cylindrical combustion elements can be used to rigidize such non-cylindrical or irregular shapes. The use of gaseous oxidizers along with deployable structures is also disclosed.

A landing inflation system may include a compressed gas source, an airbag assembly fluidly coupled to the compressed gas source and configured to receive compressed gas, and a command processor configured to: (1) generate an inflation initiation command signal in response to an initiation signal, (2) generate an inflation cessation command signal in response to a nominal fill pressure signal, and (3) generate a vent command signal in response to an impact pressure signal.

An ordered, 3-dimensional, micro-scale, open-cellular truss structure including interconnected hollow polymer tubes. The hollow micro-truss structure separates two fluid volumes which can be independently pressurized or depressurized to control flow, or materials properties, or both. Applications for this invention include deployable structures, inflatable structures, flow control, and vented padding.

A landing inflation system may include a compressed gas source, an airbag assembly fluidly coupled to the compressed gas source and configured to receive compressed gas, and a command processor configured to: (1) generate an inflation initiation command signal in response to an initiation signal, (2) generate an inflation cessation command signal in response to a nominal fill pressure signal, and (3) generate a vent command signal in response to an impact pressure signal.

A spacecraft includes: a body a surface of revolution connected with the body, and a heat engine positioned at the center of the surface of revolution. The surface of revolution has a substantially open hollow torus shape with a transverse cross-section of a circle and two diametrically opposite portions each having a curvature extending from the circle. A first portion of the diametrically opposite portions has an opening and forms a solar concentrator for concentrating solar radiation in the direction of the heat engine. The first portion forms a primary solar radiation reflector. A second portion of the diametrically opposite portions is coaxial with the first portion and forms a secondary solar radiation reflector. The opening is configured so that the solar radiation passes in the direction of the center of the surface of revolution after reflection at the primary and secondary reflectors.

The invention relates to a space wing, produced by means of a diaphragm forming a polygonal surface provided with an inflatable structure which includes ribs extending over the diaphragm along diagonals of the diaphragm and passing through a central point of the diaphragm. The inflatable structure includes at least one film strip, the perimeter of which adheres onto the diaphragm such as to form an inflatable space with the diaphragm.

A lunar base supply method which enables supplying a base on a surface of the moon more easily than carrying materials from the earth to the moon and building the base on the surface of the moon. The lunar base supply method includes the steps of: (a) carrying materials to be used to build a structure to sky of the earth, and causing the materials to revolve along an orbit around the earth; (b) building the structure by using the materials on the orbit around the earth; and (c) moving the built structure from the orbit around the earth toward the moon, and causing the built structure to make a soft landing onto a surface of the moon and thereby to be available as a base.