Abstract: A pressurized structure panel for forming a pressurized space pressurized inside includes a panel structure having a panel body to receive the pressure, a rib provided in the panel body, and a hollow part formed by the panel body and the rib, a radiation shielding material provided in the hollow part, and a debris bumper provided outside the panel structure and provided spaced apart from the panel structure by a certain spacing. The radiation shielding material includes a material containing hydrogen atoms in molecules.

Abstract: Segmented shields are disclosed comprising material combinations and rolled configurations to impede and mitigate through-and-through penetration damage of a structure incorporating the segmented shields from the damage resulting from the impact of high-velocity micro-meteoroids and orbital debris particulate.

Abstract: The present invention relates to a composite material based on aluminium or copper and tin oxide-functionalized carbon nanotubes, to the method for producing same and to a cable comprising said composite material as the electrically conductive element.

Abstract: A multilayer particle shield for a spacecraft includes an inboard exterior layer configured to be disposed proximal to the spacecraft, an outboard exterior layer configured to be disposed distal from the spacecraft and at least one interior layer disposed between the inboard exterior layer and the outboard exterior layer, wherein the interior layer includes a semi-rigid, porous, compressible spacer.

Abstract: Segmented shields are disclosed comprising material combinations and rolled configurations to impede and mitigate through-and-through penetration damage of a structure incorporating the segmented shields from the damage resulting from the impact of high-velocity micro-meteoroids and orbital debris particulate.

Abstract: There is provided an oriented wire electrostatic radiation protection system for a spacecraft. The system has a wire management system, and first and second wires coupled to the wire management system. A first wire orientation apparatus orients the first wire in a first wire direction toward, and in parallel alignment with, an approach path of approaching solar particles. A second wire orientation apparatus orients the second wire in a second wire direction opposite to the first wire direction. The system has a control system, and a power supply to charge the first wire to a positively-charged wire and to charge the second wire to a negatively-charged wire. When the approaching solar particles travel alongside the positively-charged wire toward the spacecraft, the positively-charged wire deflects the approaching solar particles away from the spacecraft, via electrostatic repulsion, and the positively-charged wire creates a radiation protection shielded region around the spacecraft.

Abstract: A living atmosphere control system is provided with a primary living area, a green cell containing plants consisting of Boston fern, aloe vera, areca palm, peace lily, and garden mums, at least one inlet configured for transporting oxygen from said green cell into said living area, and at least one outlet configured to transport carbon dioxide from said living area to said dehumidifying coils.

Abstract: A spacecraft for removing space debris is disclosed. The spacecraft includes a satellite bus, a shield member foldable on an outer side face of the satellite bus and disposed facing towards space debris to reduce a movement speed of the space debris, and a support member configured to support the shield member with respect to the satellite bus, in which the shield member includes a central panel configured to overlap one face of the satellite bus, a plurality of first panels connected to peripheral sides of the central panel and radially extended, and a plurality of second panels located between the first panels.

Abstract: A method and apparatus for cultivating vegetation at an arid location includes rooting immature vegetation in a mat combined with a super absorbent polymer (“SAP”) and, in embodiments, fertilizer, sand, and/or soil; placing the mat at the arid location; covering the mat with a perforated, transparent or semi-transparent cover; collecting urine in a water purification system; and providing purified water extracted from the urine to the mat. A water barrier can be placed below the mat. The cover can be placed on or suspended above the mat. SAP, seeds, and/or additional water barriers can be placed between mats in a stack. The opacity of the cover can be increased to emulate shade from natural vegetation. A water distribution system can be included for continued support of the vegetation, and can include a water reservoir and/or at least one solar still.

Abstract: A method and system for visualizing infrared radiation emitted by a source, the visualizing system including a substrate, at least one solid temperature-sensitive layer containing thermochromic pigments having at least one transition temperature TH associated with a visible color change, and at least one intermediate layer placed between the substrate and the solid layer containing thermochromic pigments; the substrate, the at least one intermediate layer and the solid temperature-sensitive layer containing thermochromic pigments forming a multilayer suitable for absorbing infrared electromagnetic radiation and for inducing a local temperature increase dT; and the at least one intermediate layer being suitable for transferring, by thermal conduction, the local temperature increase dT to a region of the solid temperature-sensitive layer, so as to induce a local change in the visible color of the solid temperature-sensitive layer of thermochromic pigments.

Abstract: Vehicles with control surfaces and associated systems and methods are disclosed. In a particular embodiment, a rocket can include a plurality of bidirectional control surfaces positioned toward an aft portion of the rocket. In this embodiment, the bidirectional control surfaces can be operable to control the orientation and/or flight path of the rocket during both ascent, in a nose-first orientation, and descent, in a tail-first orientation for, e.g., a tail-down landing. Launch vehicles with fixed and deployable deceleration surfaces and associated systems and methods are also disclosed.

Abstract: A vanadium oxide thermo-sensitive film material with a high temperature coefficient of resistance (TCR) contains a rare earth element of Yttrium serving as a dopant in a preparation process. The vanadium oxide thermo-sensitive film material includes a substrate and a yttrium-doped vanadium oxide film layer. The yttrium-doped vanadium oxide film layer includes three elements of vanadium, oxygen and yttrium, wherein the atomic concentration of yttrium is at a range of 1%-8%, the atomic concentration of vanadium is at a range of 20-40% and the residue is oxygen. The method for preparing the vanadium oxide thermo-sensitive film material with high TCR includes a reactive magnetron sputtering method using a low-concentration yttrium-vanadium alloy target as a sputtering source or a reactive magnetron co-sputtering method using dual targets including a high-concentration yttrium-vanadium alloy target and a pure vanadium target as a co-sputtering source.

Abstract: Computer-aided method to predict the location of an impact on an aircraft of debris shed off from the own aircraft comprising the following steps: a) providing the nominal position on the aircraft of said debris before its detachment; b) represent the debris by a body of a size and density appropriate to its characteristics; c) calculating a predetermined number of trajectories in three dimensions of said body in a predetermined fluid field when it is detached from the aircraft using an analytical model for calculating said trajectories and randomly varying one or more of the following initial conditions: the initial position of the body; the dimensions of the body; the damping coefficients of the angular velocity; the initial angles of roll, pitch and yaw; d) calculating the points of impact of said trajectories in the aircraft.

Abstract: Disclosed is a method for clearing space debris in geocentric orbit, wherein, in geocentric orbit, a spatial region having a greater resistive force than atmosphere in the geocentric orbit is provided, space debris passing through the spatial region is decelerated by the resistive force, and the decelerated space debris is dropped from the geocentric orbit.

Abstract: Heat exchange device with a boiling surface comprising a porous surface layer arranged on a solid substrate, the porous surface layer comprises a porous wall structure defining and separating macro-pores that are interconnected in the general direction normal to the surface of the substrate and have a diameter greater than 5 ?m and less than 1000 ?m wherein the diameter of the pores gradually increases with distance from the substrate wherein the porous wall structure is a continuous branched structure.

Abstract: A variable length light shield is disclosed for an electro-optical sensor within a nose cone. The light shield includes a base, a telescopic shade supported by the base, and a ring rotatably supported about the base. The light shield also includes a guide tube disposed proximate the ring with an end extending away from the ring about a side of the telescopic shade. The light shield further includes an extension spring supported by the guide tube with an end coupled to the telescopic shade. Additionally, the light shield includes a cable extending through the guide tube and the extension spring, with one end of the cable coupled to the ring and another end of the cable coupled to the telescopic shade. The extension spring is configured to exert a force on the telescopic shade to extend the telescopic shade. Rotation of the ring causes retraction of the telescopic shade.

Abstract: A device for thermal monitoring a piece of equipment, which is integrated on a craft placed in a forced vacuum environment, an outside part of the equipment projecting outside of a wall of the craft and being subjected to a solar radiation flux, includes an absorbent screen that is suitable for being placed between the outside part of the equipment and the wall of the craft, whereby this absorbent screen has—on at least one portion of its front face, designed to be placed on the side removed from the wall of the craft—an absorptivity ?SOLAR that is the largest possible in the field of solar radiation, coupled to a low emissivity ?IR in the infrared spectrum. The absorbent screen is made of a very heat-conductive material and has—on at least a portion of its rear face, suitable for being oriented toward the wall of the craft—a high emissivity ?IR in the infrared spectrum, typically greater than or equal to 0.7.

Abstract: A window system for a vehicle comprising a pressure and thermal window pane, a seal system, and a retainer system. The pressure and thermal window pane may be configured to provide desired pressure protection and desired thermal protection when exposed to an environment around the vehicle during operation of the vehicle. The pressure and thermal window pane may have a desired ductility. The seal system may be configured to contact the pressure and thermal window pane to seal the pressure and thermal window pane. The retainer system may be configured to hold the seal system and the pressure and thermal window pane.

Abstract: A rocket is provided and includes booster stages at a rear of the nose cone, the booster stages being configured for propelling the nose cone in a propulsion direction and a divert control system housed entirely in the nose cone for controlling an orientation of the propulsion direction.

Abstract: A hydrogen storage tank for a hydrogen fueled aircraft. The tank has a wall made of layers of aerogel sections around a hard shell layer, sealed within a flexible outer layer, and having the air removed to form a vacuum. The periphery of each layer section abuts other sections of that layer, but only overlies the periphery of the sections of other layers at individual points. The wall is characterized by a thermal conductivity that is lower near its gravitational top than its gravitational bottom. The tank has two exit passageways, one being direct, and the other passing through a vapor shield that extends through the wall between two layers of aerogel. A control system controls the relative flow through the two passages to regulate the boil-off rate of the tank.

Abstract: The invention relates to a surface section that is exposed to an ion flow, in particular for a drive arrangement in a spacecraft, comprising an electrostatic ion accelerator arrangement. According to the invention, in order to reduce erosion, an intermediate potential energy surface is provided, the surface being advantageous in that it allows a magnetic field that is essentially parallel to the surface section to be formed.

Abstract: A thin membrane structure for stabilizing a spacecraft is provided. The spacecraft has a first end and a second end. The spacecraft has a center of pressure and a center of gravity. The thin membrane structure comprises a module secured to the spacecraft and a structure having a polygonal base and a plurality of triangular side panels meeting in a point opposite the base with the point secured to the module. Prior to deployment, the structure is entirely contained within the module. Subsequent to deployment, the structure expands to a pyramidal shape with the distance between the center of pressure and the center of gravity increasing to a length greater than approximately one-half the length of the spacecraft.

Abstract: A thermal protection system for a vehicle may include a substantially rigid, relatively thin outer aeroshell, a relatively low density insulation layer, and a resiliently compressible conformal layer. The vehicle may include a substructure. The aeroshell may be configured to be fastened to the substructure. The insulation layer may be disposed against the aeroshell. The conformal layer may be disposable against the insulation layer. The conformal layer may be compressively preloaded against the substructure when the aeroshell is fastened to the substructure.

Abstract: A passive non-ablating thermally protective structure body is disclosed. An improved-reliability heat pipe edge comprises a non-porous material formed into a wedge-shaped chamber, and a porous layer. The porous layer extends from interior surfaces of the non-porous material to provide capillary wicking.

Abstract: A spacecraft may include a module structure having a plurality of module sides. The spacecraft may include a central cylinder extending through a center of the spacecraft. The central cylinder may be the only closed cross-section extending along a longitudinal axis of the spacecraft.

Abstract: A spacecraft is described which uses a magnetic field source to generate a shield magnetic field to protect the spacecraft from energetic charged particles. The field may be perturbed to increase the effectiveness of the protection. Injection of material into the shield cavity to enhance the local plasma density may also be used.

Abstract: A design method for a thermal control film advantageously used in cosmic space, especially for a multilayer film, includes: setting at least two candidate materials for forming the multilayer film and the number of layers of the multilayer film; and designing a combination of the candidate material and a layer thickness by using a genetic algorithm stored in a computer-readable medium using an evaluation function, and recording the obtained results on a computer-readable medium. The evaluation function includes: ?s=1?Rs, wherein Rs represents a reflectance of solar energy in a wavelength range used for design, within a solar radiation wavelength range; and ?H: ?H=1?RIR, wherein RIR represents a reflectance of heat radiation energy in a wavelength range used for design, within a heat radiation wavelength range.

Abstract: A telecommunication satellite with geostationary orbit comprises an upper module, a lower module, and a lateral module, disposed in a storage configuration between the upper module and the lower module, and deployed to an operational configuration of the satellite in the orbit by a rotation in relation to an axis Z oriented towards the earth in the operational configuration. The lateral module comprises two substantially plane and mutually parallel main surfaces, termed dissipative surfaces, able to dissipate by radiation a quantity of heat generated by facilities of the satellite; the dissipative surfaces being, in the operational configuration, held in a manner substantially parallel to the plane of the orbit, making it possible to limit the solar flux received by the dissipative surfaces and to optimize the quantity of heat dissipated by the lateral module.

Abstract: Radiation shields include an outer composite wall, an inner composite wall, and an enclosed fuel chamber disposed between the inner and outer composite walls. A fuel may be contained within the enclosed fuel chamber. Spacecraft include one or more such radiation shields, and may further include a fuel cell and a conduit providing fluid communication between the fuel chamber of the radiation shield and the fuel cell. Methods of forming a radiation shield system include forming an enclosed fuel chamber between an outer composite wall and an inner composite wall of a radiation shield, and providing fuel within the enclosed fuel chamber.

Abstract: A vehicle for launching from a gun such as a gas gun and having a housing; preferably incorporating a precessional attitude control system; and utilizing a flared base, fins, or active use of the attitude control system during passage through the atmosphere. Subtly canting the fins can produce desired spinning of the vehicle. The propulsion system can employ liquid, hybrid, or solid fuel. A removable aero-shell assists atmospheric flight with thermal protection being provided by anticipated ablation, an ablative aero-spike, or transpirational cooling. And a releasable sabot enhances the effectiveness of the launch.

Abstract: There is provided an integrated articulating thermal isolation system having an articulating actuator assembly, a support frame assembly, and a thermal blanket assembly. The thermal blanket assembly has a pair of non-articulating blankets positioned on exterior surfaces of the support frame assembly and a pair of articulating blankets disposed between the non-articulating blankets. Each articulating blanket has a plurality of stiffener elements. The system further has a harness assembly, a deployable assembly interface element, and a vehicle interface element. When the articulating actuator assembly is actuated, the articulating blankets are guided by the support frame assembly from a stowed position to a deployed position through motion of the articulating actuator assembly.

Abstract: Disclosed herein, in certain embodiments, is a method of altering the stability of unstable space debris. In some embodiments, the method further comprises changing the orbit of the unstable space debris.

Abstract: Fiber-based debris interceptors are used to intercept and/or contain space debris. The debris interceptors may be made up of fibers that are formed in space from a material supply on a space vehicle. The fibers for the debris interceptors may be formed by extrusion, with a heat source, such as a mirror to focus sunlight, used to heat the material of the material supply that is carried on the space vehicle. The debris interceptor may be separated from the space vehicle, and used to remove debris from an orbit, or otherwise prevent debris entering an orbit from damaging a satellite or other spacecraft that travels in that orbit. The debris interceptor may be deployed prior to later launch of a valuable spacecraft, in order to “cleanse” the intended orbit of debris. Debris objects may pass through the debris interceptor, but in so doing may lose energy so as to de-orbit.

Abstract: A spacecraft magnetic shield apparatus is disclosed. One embodiment has an armature attached at one end of a spacecraft and an electromagnet attached to the other end. A computer determines the position of the armature and electromagnet in relation to the spacecraft, and a power source powers the electromagnet. In operation, the positioning of the electromagnet is generally in line with incoming charged particles or ion radiation and powering of the electromagnet creates a magnetic field that deflects a percentage of the incoming charged particles or ion radiation from directly impacting the spacecraft.

Abstract: A vehicle for launching from a gas gun having a housing; preferably incorporating a precessional attitude control system; and utilizing a flared base, fins, or active use of the attitude control system during passage through the atmosphere. Subtly canting the fins can produce desired spinning of the vehicle. The propulsion system can employ liquid, hybrid, or solid fuel. A removable aero-shell assists atmospheric flight with thermal protection being provided by anticipated ablation, an ablative aero-spike, or transpirational cooling. And a releasable sabot enhances the effectiveness of the launch.

Abstract: A missile includes a separable shroud that covers a nose portion of the fuselage of the missile. The shroud covers and protects a seeker window and a seeker at the nose of the missile. The shroud is configured to remain coupled to the missile during and immediately after launch of the missile, and to separate during flight under the action of aerodynamic forces. Toward that end parts of the shroud are initially coupled together by a retainer, which allows the parts to separate during flight. The retainer may include one or more tension bands that break at a certain tension, and/or one or more weakened parts of the shroud, which break during flight. Parts of the shroud may include inward protrusions that make contact with an ogive portion of the nose of the fuselage.

Abstract: The chassis of a spacecraft has one side exposed to ultraviolet photons and another side shielded from the ultraviolet photons. An electrically conductive surface is disposed on the exposed side of the chassis and is electrically isolated from the chassis. A field-emission array device has a gate, an emitter array, a first terminal electrically connected to the gate, and a second terminal electrically connected to the emitter array. The first terminal electrically couples the gate to the electrically conductive surface on the exposed side of the chassis, and the second terminal electrically couples the emitter array to a surface on the shielded side of the chassis. The emitter array discharges electrons from the surface on the shielded side when a voltage difference in excess of a threshold voltage develops between the gate and the emitter array because of differential charging of the exposed and shielded sides of the chassis.

Abstract: An orbital debris removal assembly is provided. The orbital debris removal and asset protection assembly includes a plurality of spaced layers of netting. Each netting layer includes a plurality of sub-layers of meshed composite fibers that are configured and arranged to break up debris that comes in contact with the assembly.

Abstract: An orbital debris removal assembly is provided. The orbital debris removal and asset protection assembly includes a plurality of spaced layers of netting. Each netting layer includes a plurality of sub-layers of meshed composite fibers that are configured and arranged to break up debris that comes in contact with the assembly.

Abstract: A blanket assembly includes a shear fabric and an attachment frame wrapped in the shear fabric. A thermal protection blanket is carried by the shear fabric.

Abstract: A cryogenically cooled radiation shield device as well as an associated method are provided in order to shield an area, such as a space vehicle capsule, from radiation. The radiation shield device may have embedded cryogen storage. The radiation shield device may include inner and outer coil shells that extend about the area to be shielded. Each coil shell includes coils formed of a superconductive material and disposed within respective first conduits. The radiation shield device may also include a first storage tank configured to store a first cryogen liquid and disposed in fluid communication with the first conduits. The first storage tank is disposed between the inner and outer coil shells. The radiation shield device may further include a second conduit at least partially surrounding the inner and outer coil shells that is at least partially filled with a second cryogen liquid, different than the first cryogen liquid.

Abstract: The invention relates to the protection of spacecraft from debris and to de-orbiting devices of the atmospheric drag type, and to debris sweeping apparatus for the removal of debris from the space environment. The debris shielding apparatus for a spacecraft has a shield unit including a shielding surface for impeding incident debris. The shield unit is attached to the spacecraft body and has a drive mechanism for positioning the shield unit in relation to the spacecraft body. The drive mechanism is capable of moving the shield unit between a stowed first position and a deployed second position. In the deployed second position the plane of the shielding surface is at an angle to the spacecraft body.

Abstract: This invention relates to technology used for creating an enhanced artificial magnetosphere or electromagnetic shield for use in both manned and unmanned spacecraft. The invention includes an Interference Generating Pattern (IGP) which is tuned to the high-frequency radiation of X-rays and gamma rays and a conformal magnetic field. This technology will reduce the exposure of astronauts or other space travelers, as well as radiation-sensitive equipment, to the environmental hazards present therein. The net result will be reduced radiation intensity in order to create a space radiation-free environment as to render space travel safe.

Abstract: A portable thermal-control system adapted to support space-related research and exploration. Embodiments of the present invention assist in preventing overheating of small payloads being transported from an orbiting space vehicle to a planetary surface by small atmospheric-entry vehicles. Other embodiments of the present invention provide thermal control within an extra-vehicular activity (EVA) suit. Each embodiment utilizes at least one phase-change material, cooled significantly below the freezing temperature, to absorb heat.

Abstract: An aircraft insulation apparatus is disclosed having a variable thermal insulating member adjacent the inner fuselage wall and adjacent an aircraft heat source. The variable thermal insulating member is operable to provide variable thermal insulation between the heat source and the exterior of the inner fuselage wall by being changeable between a first state and a second state, the first state providing a first level of thermal insulation and the second state providing a second level of thermal insulation.

Abstract: A spacecraft magnetic shield apparatus is disclosed. One embodiment has an armature attached at one end of a spacecraft and an electromagnet attached to the other end. A computer determines the position of the armature and electromagnet in relation to the spacecraft, and a power source powers the electromagnet. In operation, the positioning of the electromagnet is generally in line with incoming charged particles or ion radiation and powering of the electromagnet creates a magnetic field that deflects a percentage of the incoming charged particles or ion radiation from directly impacting the spacecraft.

Abstract: Radiation shields include an outer composite wall, an inner composite wall, and an enclosed fuel chamber disposed between the inner and outer composite walls. A fuel may be contained within the enclosed fuel chamber. Spacecraft include one or more such radiation shields, and may further include a fuel cell and a conduit providing fluid communication between the fuel chamber of the radiation shield and the fuel cell. Methods of forming a radiation shield system include forming an enclosed fuel chamber between an outer composite wall and an inner composite wall of a radiation shield, and providing fuel within the enclosed fuel chamber.

Abstract: A device for protecting an optical instrument of a satellite comprises a body on which the optical instrument is mounted, said protection device having a folded-down position and a deployed position, and further comprising a plurality of rigid panels in the deployed position, the device forming a tube, the section of which is a polygon, arranged so as to protect the optical instrument, the longitudinal axis of the tube being substantially parallel with the focal axis of the optical instrument, the faces of the tube of polygonal section being formed by panels, in the folded-down position, said panels being held against the body of the satellite.

Abstract: An attachment system for mounting a tile of a thermal protection system to a substructure comprises a post assembly configured to interconnect the tile to the substructure. The post assembly is configured to allow relative sliding movement between the tile and the substructure along a plane which is generally parallel to the tile. The attachment system facilitates installation and removal of the tile from the substructure for inspection, maintenance and repair of the tile and/or the substructure.

Abstract: The present invention relates to a solar protection device for space instrument. This invention is particularly suited to electromagnetic measuring instruments pointing in a direction close to that of the sun. To ensure the continuous availability, the performance and the non-degradation of collecting instruments directed towards the sun, the invention proposes the use of solar baffling thermally and mechanically decoupled from the platform supporting the collecting instruments of the equipment.