Abstract: A method, apparatus, and system include a cooling system comprising a payload refrigeration unit, control refrigeration unit, and a signal interface. The payload refrigeration unit has a set of payload cooling components that operate to cool a payload. The control refrigeration unit has a set of control circuit cooling components in a control circuit. The signal interface connecting the payload is located in the payload refrigeration unit in the control circuit located in the control refrigeration unit.

Abstract: The present disclosure relates to a thermal emissivity control system. The system may have a segmented array that makes use of a thermally conductive base layer configured to be connectable to an external heat generating subsystem, with the base layer including a thermally emissive surface. The array may also have a plurality of actuation elements at least one of positioned on or adjacent to the thermally emissive surface. A plurality of movable shutter elements is disposed adjacent one another in a grid pattern, and controlled in movement by the actuation elements to create gaps of controllably varying dimension therebetween. The shutter elements control at least one of a magnitude of, or direction of, thermal radiation through the gaps.

Abstract: A spacecraft sunshade is provided. The sunshade includes a surface that is maintained in a sun facing orientation. Adjustments to a position of the sunshade are made in a plane that is transverse to a line of sight to the sun, in order to block sunlight from being directly incident on an instrument associated with the spacecraft. The sunshade can include photovoltaic elements on the sun-facing surface of the sunshade. In addition, the sunshade can be formed from an opaque material, and further from a material that absorbs heat from the sun and reradiate that heat to the instrument. The sunshade can perform stray light blocking, electrical power generation, and radiational heating functions.

Abstract: Integrated power module device, systems, and methods are provided in accordance with various embodiments. For example, some embodiments include a system that may include one or more integrated power modules. Each integrated power module may include: one or more solar cells; one or more rechargeable energy storage cells; and/or one or more circuits coupling the one or more solar cells with the one or more rechargeable energy storage cells. In some embodiments, each integrated power module is configured such that the one or more rechargeable energy storage cells of the respective integrated power module are coupled with one or more back sides of the one or more solar cells. In some embodiments, at least two of the one or more integrated power modules are coupled with each other at least in parallel or in series.

Abstract: A method to improve thermal performance of vascular composites by using a two-phase working fluid for isothermalization includes the steps of: manufacturing a vascular composite structure optimized for a design point; manufacturing a thermal back end sized for the application; integrating the vascular composite into a fluid loop; and evacuating and filling the fluid loop with working fluid to an amount resulting in two-phase operation at the design point.

Abstract: A handheld power tool with a rotatable tool, in which the handheld power tool is a saw blade or milling cutter. The handheld power tool includes a sensor device for detecting a mechanical vector quantity. The mechanical vector quantity includes a force, an acceleration, a velocity, a deflection, a deformation and/or a mechanical stress and the mechanical vector quantity depends on a force emanating from the handheld tool. The handheld power tool also includes a control device which is communicatively coupled to the sensor device and is adapted to recognize an event and/or a state of the power tool according to a direction and/or change of direction of the mechanical vector quantity detected by the sensor device.

Abstract: Inner panels including at least one built-in heat pipe connected in a circumferential direction are provided. In a heat pipe panel including the built-in heat pipe, apparatuses are mounted on the outer side of the plural inner panels connected in the circumferential direction to diffuse generated heat of the apparatuses to the circumferential direction of the inner panels. Webbed panels including a built-in heat pipe horizontally arranged and having heat radiation surfaces are radially arranged at corners of the inner panels as well as a heat pipe is horizontally built in and heat radiation surfaces are arranged also on outer panels facing the inner panels to thermally connect the heat pipes to one another.

Abstract: A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. The leading edge assembly includes a plurality of structural layers and a plurality compliant layers alternately stacked with each other to facilitate thermal expansion and movement between the plurality of structural layers, while also providing a thermal break between the plurality of structural layers.

Abstract: Systems, methods, and apparatus for dual condenser loop heat pipes for satellites with sun-normal radiators are disclosed. In one or more embodiments, a disclosed method for a satellite thermal management system comprises heating, in an evaporator, a liquid to convert the liquid to a vapor. The method further comprises passively circulating within tubing, from the evaporator, the vapor to a first radiator not illuminated by a sun and to a second radiator illuminated by the sun. Also, the method comprises converting the vapor to the liquid when the vapor is within the first radiator not illuminated by the sun. Further, the method comprises passively circulating within the tubing, from the first radiator not illuminated by the sun, the liquid to the evaporator.

Abstract: Disclosed is an artificial satellite including one face from the North face and the South face supporting at least one main radiator having an outer face turned towards space and an inner face opposite the outer face. The satellite includes a bearing structure carrying the North face, the South face, the East face and the West face. At least one part of the at least one main radiator protrudes from at least one face from the East face and the West face. The inner face of the at least one protruding part is covered with a high infrared emissivity material. The value of the dimension of the at least one protruding part is between 19% and 50% of the value of the distance between the East and West faces.

Abstract: An apparatus including a battery pack comprising a plurality of individual batteries arranged around a battery cavity such that each individual battery is in thermal contact with at least one neighboring individual battery. A variable-conductance heat pipe (VCHP) having an evaporator end and a condenser end, is positioned so that at least part of the evaporator end being positioned in the battery cavity and in thermal contact with each of the plurality of individual batteries. The apparatus includes a thermally insulating cover having an inside and an outside, wherein the battery pack and the part of the VCHP evaporator end in the battery cavity are positioned inside the thermally insulating cover and at least part of the condenser end of the VCHP is outside the thermally insulating cover, and wherein the VCHP is substantially the only thermal path between the battery pack and the outside. Other implementations are disclosed and claimed.

Abstract: Radiator panels include two spaced-apart face-sheets including an inside face-sheet and an outside face-sheet that are constructed of a fiber reinforced composite material, a honeycomb core positioned between the two spaced-apart face-sheets, and one or more heat pipes extending through the honeycomb core. Spacecraft include a body and two radiator panels operatively coupled to the body opposite each other.

Abstract: A system and apparatus are provided for preventing heat loss in a spacecraft during the upper stage launch phase and transfer orbit. A thermal control panel can be provided that can be positioned adjacent to exposed areas of radiator panels on the spacecraft in a collapsed configuration. The outer surface of the panel can have a high absorptivity, and the inner surface can have a high emissivity. During upper stage launch phase and transfer orbit, the panel can tend to emit heat toward the radiator panels on an inboard side of the spacecraft, thus reducing heat loss from the radiator panels to reduce the need for onboard electronic heaters and thereby preserve battery power for other onboard systems.

Abstract: A cryogenically cooled radiation shield device and method are provided to shield an area, such as the capsule of a space vehicle, from radiation. A cryogenically cooled radiation shield device may include at least one first coil comprised of a superconducting material extending about an area to be shielded from radiation. The cryogenically cooled radiation shield device also includes a first conduit extending about the area to be shielded from radiation. The at least one first coil is disposed within the first conduit. The cryogenically cooled radiation shield device also includes a second conduit extending about the area to be shielded from radiation. The first conduit is disposed within the second conduit. The cryogenically cooled radiation shield device also includes a first cryogen liquid disposed within the first conduit and a second cryogen liquid, different than the first cryogen liquid, disposed within the second conduit exterior of the first conduit.

Abstract: An instrument mounting system for a geosynchronous host satellite provides independent heat rejection capability for a hosted instrument. The system may include an instrument mounting structure that is thermally coupled to a radiator component via heat pipes that may include one or both of rigid and flexible elements. In some embodiments, the system is mounted on the hosting satellite so as to provide satellite components and the hosted instrument with a clear field of view while rejecting heat in one or both of the north and south directions. The system may also provide structural support for components of one or both of the hosted instrument or the hosting satellite.

Abstract: A heat transfer assembly can include an equipment panel having an east end and a west end. East and west radiator panels are coupled to the east and west ends, respectively, of the equipment panel. The assembly also includes a plurality of flexible heat pipes each having a first rigid tube thermally coupled to the east radiator panel, a second rigid tube coupled to the equipment panel, a third rigid tube thermally coupled to the west radiator panel, a first flexible tube sealingly coupled between the first and second rigid tubes, and a second flexible tube sealingly coupled between the second and third rigid tubes. The equipment panel is configured to retain one or more equipment modules in thermal contact with the second rigid tube of at least one of the plurality of flexible heat pipes.

Abstract: A space probing apparatus is provided that can further enhance its thermal protective function for electronic apparatuses and batteries that are apparatuses each having a narrow range of permissible temperature. The space probing apparatus includes, as a vehicle body 1 that is an apparatus body, an inner housing 13 that accommodates an electronic apparatus 11 and a battery 12, and an outer housing 15 that accommodates the inner housing 13 supporting the inner housing 13 afloat therein. A thermal insulating layer 16 is formed between the inner housing 13 and the outer housing 15 using a space therebetween. A heat dissipating body 17 is included in an upper portion of the inner housing 13. A lid 18 is included in an upper portion of the outer housing 15. The whole outer faces of the inner and the outer housings 13 and 15 are each covered with a thermal insulating material 20.

Abstract: A geostationary earth orbit (GEO) spacecraft is disclosed that includes a body with north, east, south, and west sides and a north-south axis. The spacecraft has at least one deployable radiator rotatably coupled to the body. The deployable radiator has a stowed position proximate to one of the east and west sides and a deployed position that is greater than 90 degrees from the north-south axis in a direction away from the respective one of the east and west sides.

Abstract: A thermal management system and method for active cooling of high speed seeker missile domes or radomes comprising bonding to an IR dome or RF radome a heat pipe system having effective thermal conductivity of 10-20,000 W/m*K and comprising one or more mechanically controlled oscillating heat pipes, employing supporting integrating structure including a surface bonded to the IR dome or RF radome that matches the coefficient of thermal expansion the dome or radome material and that of said one or more mechanically controlled oscillating heat pipes, and operating the heat pipe system to cool the IR dome or RF radome while the missile is in flight.

Abstract: A cooling device for regulating the temperature of a heat source of a satellite. The cooling device comprises at least one fluid loop is formed by an evaporator comprising a tank, at least one condenser, two conduits connecting the evaporator to the condenser, a heat transfer fluid flowing inside the fluid loop. The cooling device further comprises a device for pressurizing the fluid loop or a thermal damper. The thermal damper comprises a variable volume leak-tight chamber having a volume which varies on the basis of the operating temperature of the fluid loop so as to provide a substantially constant temperature inside the fluid loop.

Abstract: A thermal module for use on a spacecraft to control thermal loads coming from a heat source is disclosed. The thermal module includes a two-phase loop system and a heat rejection system. The two-phase loop system includes a thermal collector, a heat flow regulator, a by-pass line, and a condenser. The condenser and the heat rejection system are thermally coupled, such that the heat flow regulator redirects part of the thermal loads from the heat source to the condenser, from which the heat rejection system directs said thermal loads to a heat sink. The temperature of the heat source is regulated by bypassing another part of the thermal loads back to the thermal collector through the by-pass line in a proportional manner, to avoid overcooling the heat source. The heat rejection system is designed based on the hottest possible conditions for the spacecraft mission.

Abstract: The invention relates to a fuselage of an air- or spacecraft. The fuselage comprises at least one shell element and one insulation element. The insulation element is configured as a passive, watertight insulation element and can be mounted on an inner side in a completely air- and watertight manner. The invention furthermore relates to a respective aircraft or spacecraft.

Abstract: A system and apparatus are provided for preventing heat loss in a spacecraft during the upper stage launch phase and transfer orbit. A thermal control panel can be provided that can be positioned adjacent to exposed areas of radiator panels on the spacecraft in a collapsed configuration. The outer surface of the panel can have a high absorptivity, and the inner surface can have a high emissivity. During upper stage launch phase and transfer orbit, the panel can tend to emit heat toward the radiator panels on an inboard side of the spacecraft, thus reducing heat loss from the radiator panels to reduce the need for onboard electronic heaters and thereby preserve battery power for other onboard systems.

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: The essence of the invention is a method and apparatus for thermally protecting the external surface of the structure of a space vehicle during atmospheric reentry. The method, in detail, uses a recently discovered fundamental electrodynamic combustion mechanism for alkali metals to burn the metal onto the surface of the outer wall of the vehicle. In this mechanism, a static electric field that is always inherent to the burning alkali metal surface interacts electrodynamically with the ionized particles streaming behind the atmospheric front shock wave that propagates in front of the vehicle such that the electrodynamic Lorentz force deflects the particles perpendicularly away from the original direction of the ionized atmospheric flow. Thus, convective heat flux into the vehicle's structure is prevented. An apparatus is also presented for coating the alkali metal onto the porous outer wall of the vehicle structure via transpiration from the interior of the liquid lightweight alkali metal.

Abstract: According to one embodiment, an apparatus includes a sensor having one or more thermally sensitive components. The sensor is gimbal mounted on a space or air-borne vehicle and includes a component. The component is configured to at least partially adjust a center-of-gravity of the sensor and to at least partially receive and store thermal energy from the one or more thermally sensitive components. The Apparatus also includes a radiator configured to dissipate thermal energy to the ambient environment. The component is thermally coupled between the radiator and the one or more other thermally sensitive components. The radiator is configured to receive thermal energy from the component.

Abstract: A spacecraft is provided. The spacecraft comprises a plurality of heat generating electrical components. A first thermal radiator panel and a second thermal radiator panel are provided on the spacecraft, each panel being thermally coupled to the heat generating electrical components. Heat pipes are also provided. At least one first heat pipe is externally attached to the first thermal radiator panel and at least one second heat pipe is externally attached to the second thermal radiator panel. The at least one first heat pipe is thermally coupled to the at least one second heat pipe.

Abstract: A cryogenically cooled radiation shield device and method are provided to shield an area, such as the capsule of a space vehicle, from radiation. A cryogenically cooled radiation shield device may include at least one first coil comprised of a superconducting material extending about the area to be shielded. The cryogenically cooled radiation shield device also includes a first inner conduit extending about the area to be shielded from radiation. The at least one first coil is disposed within the first inner conduit. The cryogenically cooled radiation shield device also includes a first outer conduit extending about the area to be shielded from radiation. The first inner conduit is disposed within the first outer conduit. The cryogenically cooled radiation shield device also includes a first cryogen liquid disposed within the first inner conduit and a second cryogen liquid, different than the first cryogen liquid, disposed within the first outer conduit.

Abstract: A heat transfer assembly for a spacecraft is disclosed. The assembly includes an equipment panel having an east end and a west end. East and west radiator panels are coupled to the east and west ends, respectively, of the equipment panel. The assembly also includes a plurality of flexible heat pipes each having a first rigid tube thermally coupled to the east radiator panel, a second rigid tube coupled to the equipment panel, a third rigid tube thermally coupled to the west radiator panel, a first flexible tube sealingly coupled between the first and second rigid tubes, and a second flexible tube sealingly coupled between the second and third rigid tubes. The equipment panel is configured to retain one or more equipment modules in thermal contact with the second rigid tube of at least one of the plurality of flexible heat pipes.

Abstract: A deployable radiator arrangement for cooling a geostationary earth orbit spacecraft is provided. In some aspects, the geostationary earth orbit spacecraft may comprise first and second deployable radiators mounted on an east or west surface of the spacecraft when stowed. The first and second deployable radiators are configured to rotate into a north and south facing position, respectively, when deployed. The geostationary earth orbit spacecraft may further comprise first and second fixed radiators disposed on a north and south surface of the spacecraft, respectively. The first and second deployable radiators are thermally coupled to the first and second fixed radiators, respectively.

Abstract: According to an embodiment, a satellite is disclosed comprising an anti-nadir panel having a first side and a second side. In one or more embodiments, the first side of the anti-nadir panel is mounted to an anti-nadir side of the main body of the satellite. The satellite further comprises at least one battery pack mounted to the second side of the anti-nadir panel, where at least one battery pack comprises at least one battery cell.

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: A geostationary earth orbit (GEO) spacecraft is disclosed that includes a body with north, east, south, and west sides and a north-south axis. The spacecraft has at least one deployable radiator rotatably coupled to the body. The deployable radiator has a stowed position proximate to one of the east and west sides and a deployed position that is greater than 90 degrees from the north-south axis in a direction away from the respective one of the east and west sides.

Abstract: A deployable radiator arrangement for cooling a geostationary earth orbit spacecraft is provided. In some aspects, the geostationary earth orbit spacecraft may comprise first and second deployable radiators mounted on an east or west surface of the spacecraft when stowed. The first and second deployable radiators are configured to rotate into a north and south facing position, respectively, when deployed. The geostationary earth orbit spacecraft may further comprise first and second fixed radiators disposed on a north and south surface of the spacecraft, respectively. The first and second deployable radiators are thermally coupled to the first and second fixed radiators, respectively.

Abstract: Systems and methods reduce development, integration and testing time of a spacecraft implemented with one or more reconfigurable processor boards. Mission spacecraft design tools design, configure, simulate and analyze the spacecraft based upon constraints of the mission requirements. The mission spacecraft design tools utilize a database of defined module specifications and firmware components. One or more RPBs are programmed based upon the configuration and interfaced to other end item components defined by the configuration to form the spacecraft.

Abstract: An active IR signature target simulation system which includes: a platform having a known temperature profile and a preplanned flight trajectory; a heat source carried onboard the platform, capable of producing heat following a controlled heating plan; and a control unit connected to the heat source and configured for initiating the heat source, thereby modifying the known temperature profile and generating a desired temperature profile required for sensing a simulated target IR signature during a simulation time window at a desired line of sight from the platform to an observing sensor, wherein the platform is an exo-atmospheric missile which includes an inner shell and a perforated outer shell, and wherein the heat source is a thermally active material disposed between at least a part of the inner shell and a part of the outer shell.

Abstract: A structural panel for a satellite comprises an outer skin intended to be outside the satellite, a core comprising at least one integrated heat pipe mounted in fixed contact with said outer skin, and an inner skin intended to be inside the satellite, said structural panel being equipped with generic heat exchangers which are adapted to be associated with a heat control circuit using the circulation of a liquid coolant, said circuit being situated outside the panel.

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 high capacity satellite having multiple aspect ratios is configured to have at least two modules. Each module includes a first panel and a third panel facing, in an on-orbit configuration, respectively, toward north and south and having a width in an east-west direction (EW width); and a second and a fourth panel oriented, in the on-orbit configuration, respectively, east and west and having a width in the north-south direction (NS width). Each module has a respective aspect ratio of EW width to NS width, a first module is configured with a first aspect ratio, a second module is configured with a second aspect ratio; and the second aspect ratio is substantially larger than the first aspect ratio. At least one antenna reflector is disposed, during launch, proximate to at least one of the first and third panel of the second module.

Abstract: A device for controlling heat flow in a spacecraft which always has the same face facing towards the earth, a pair of mutually parallel opposed North and South faces perpendicular to the North-South axis of the earth and two pairs of mutually parallel opposed East/West and Earth/Anti-earth faces, heat-dissipating or -transmitting equipment being provided on the internal walls of the North and South faces, this device comprising a plurality of heat pipes connecting the North and South faces of said craft to another pair of opposed faces of said craft and forming a heat-pipe loop in thermal-conduction contact with one another.

Abstract: A spacecraft radiator system designed to provide structural support to the spacecraft. Structural support is provided by the geometric “crescent” form of the panels of the spacecraft radiator. This integration of radiator and structural support provides spacecraft with a semi-monocoque design.

Abstract: An integrated instrumentation system includes a body comprising a thermal protection system (TPS) material (e.g., an ablatable material), one or more sensors embedded within the body, and a processor (e.g., an FPGA or the like) communicatively coupled to the plurality of sensors. The processor is configured to acquire sensor signals from the plurality of sensors and produce digital sensor data associated therewith. The sensors may include, for example, recession sensors, pressure transducers, thermocouples, and accelerometers.

Abstract: A plurality of sufficient closely-packed functional units which are interconnected to form a non-planar payload module assembly 40. Advantageously, the payload modular structure 40 is compact with few piece parts, has low output losses, and is robust enough for supporting complete payloads with or without reflectors. The payload modular structure has utility in various space-based applications as well as in various terrestrial applications.

Abstract: A spacecraft battery thermal management system is provided that includes a battery, a first radiator panel and a second radiator panel. A first face of the first radiator panel is arranged to face a first direction and a first face of the second radiator panel is arranged to face a second direction opposite the first direction. A first heat pipe thermally couples the battery and the first radiator panel and is configured to control the transfer of heat between the battery and the first radiator panel. A second heat pipe thermally couples the battery and the second radiator panel and is configured to control the transfer of heat between the battery and the second radiator panel. Solar cells are optionally arranged on the faces of the first and/or second radiator panels.

Abstract: A heat shield includes a ceramic composite heat shield panel having a generally concave first surface and a generally convex second surface and a pair of thickened panel edge portions provided in the heat shield panel. A heat shield assembly is also disclosed.

Abstract: An integrated instrumentation system includes a body comprising a thermal protection system (TPS) material (e.g., an ablatable material), one or more sensors embedded within the body, and a processor (e.g., an FPGA or the like) communicatively coupled to the plurality of sensors. The processor is configured to acquire sensor signals from the plurality of sensors and produce digital sensor data associated therewith. The sensors may include, for example, recession sensors, pressure transducers, thermocouples, and accelerometers.

Abstract: A satellite (50) is disclosed having at least one deployable thermal radiator (64). Once deployed, this thermal radiator (64) may be repositioned. In one embodiment, the deployed thermal radiator (64) may be moved about a first axis (74), about a second axis (76), or both to move the deployed thermal radiator (64) from one deployed position to another deployed position.

Abstract: Thermal module (1) for use on a spacecraft to control thermal loads coming from a heat source (3) comprising a two-phase loop system (5) and a heat rejection system (13), the two-phase loop system (5) comprising a thermal collector (7), a heat flow regulator (10), a by-pass line (6) and a condenser (11), the condenser (11) and the heat rejection system (13) being thermally coupled, such that the heat flow regulator (10) of the two-phase loop system (5) redirects part of the thermal loads from the heat source (3) to the condenser (11), from which the heat rejection system (13) directs said thermal loads to a heat sink (14), the temperature of the heat source (3) being regulated by bypassing another part of the thermal loads back to the thermal collector (7) through the by-pass line (6) in a proportional manner, to avoid the heat source (3) overcooling, being the heat rejection system (13) designed based on the hottest possible conditions for the spacecraft mission.

Abstract: An afterbody device for a spacecraft fitted with at least one rocket engine at the rear of the craft includes at least one movable cover element designed to take a first position, masking and reducing the vehicle's rear drag, where it prolongs the vehicle's fuselage around at least one part of a rocket engine nozzle of the vehicle and extends beyond the rear of the vehicle's fuselage, and to take a second position fully deployed, increasing the vehicle's aerodynamic drag.

Abstract: A thermal control system for a space body uses an active control means to prevent laser beam radiation damage to a heat radiator. A conventional louver and louver actuator are coupled to an active overdrive actuator that closes the louver when hostile laser radiation is present. An extended bimetallic coil spring with a heater therein rotates opposite to the louver actuator in an increasing temperature environment. A cam of the overdrive actuator engages a louver arm when hostile laser radiation is present, otherwise, the louver can move freely within the cam.