HEAT EXCHANGER FOR A GAS TURBINE ENGINE PROPULSION SYSTEM

A propulsion system including a gas turbine engine is disclosed herein. The propulsion system further includes a heat exchanger arranged outside the gas turbine engine and adapted to cool fluid from the gas turbine engine. 1. A propulsion system for an aircraft , the propulsion system comprisinga gas turbine engine including an engine core and a fan coupled to the engine core, the fan configured to discharge pressurized bypass air that is passed around the engine core, anda nacelle surrounding a portion of the gas turbine engine, the nacelle including a strut that extends away from the gas turbine engine and a cooling unit housed in the strut, the cooling unit fluidly coupled to the gas turbine engine to cool fluid or gas from the gas turbine engine and return the cooled fluid or gas to the gas turbine engine,wherein the cooling unit includes a duct, a heat exchanger positioned within the duct, and a diverter valve that is movable within the duct from a first position arranged to direct pressurized bypass air moving through the duct into contact with the heat exchanger to a second position arranged to divert pressurized bypass air around the heat exchanger without contacting the heat exchanger.2. The propulsion system of claim 1 , wherein the cooling unit is positioned radially-outward of the engine core and radially-inward of an outer shroud included in the nacelle.3. The propulsion system of claim 1 , wherein the duct includes a divider extending along the length of the duct to divide the duct into an outer flow portion and an inner flow portion positioned radially-inward of the outer flow portion.4. The propulsion system of claim 1 , wherein the diverter valve includes a first actuator and a plate coupled to the first actuator claim 1 , the first actuator operable to pivot the plate relative to the duct between the first and the second positions.5. The propulsion system of claim 4 , wherein the diverter valve includes a second actuator coupled to the plate claim ...

Adaptive trans-critical CO2 cooling systems for aerospace applications

A cooling system includes a heat exchanger through which a refrigerant flows, and which rejects heat to a fluid, an evaporator, a first circuit having an expansion device, a second circuit having an expansion machine coupled to a compressor, and a set of valves arranged to direct the refrigerant through the first circuit, the second circuit, or both the first and second circuits based on ambient conditions.

ENHANCED HEAT SINK AVAILABILITY ON GAS TURBINE ENGINES THROUGH THE USE OF COOLERS

A cooling assembly for a gas turbine engine including a heat source at a first temperature, a heat sink at a second temperature, and a heat pump coupled to the first heat source and the first heat sink. The heat pump is configured to convey a quantity of heat from the heat source through the heat pump and to the heat sink. 1. A cooling assembly for a gas turbine engine comprising:a heat source at a first temperature;a solid heat sink at a second temperature; anda heat pump coupled to the first heat source and the first heat sink, wherein the heat pump is configured to convey a quantity of heat from the heat source through the heat pump and to the solid heat sink.2. The cooling assembly of claim 1 , wherein the solid heat sink is configured to allow a fuel to pass therethrough claim 1 , and wherein the second temperature is greater than the first temperature.3. The cooling assembly of claim 1 , wherein the heat source is a full authority digital engine control (FADEC).4. The cooling assembly of claim 1 , wherein the heat source is a heat exchanger having a material passing therethrough.5. The cooling assembly of claim 4 , wherein the material is one of a heat transfer fluid.6. The cooling assembly of claim 1 , wherein the heat pump is a thermionic cooler.7. The cooling assembly of wherein the heat pump is a thermoelectric cooler.8. The cooling assembly of claim 1 , further comprising a thermoelectric generator electrically coupled to the heat pump claim 1 , wherein the thermoelectric generator is configured to power the heat pump.9. The cooling assembly of claim 8 , wherein the thermoelectric generator is configured to have a first side coupled with a first fluid stream and a second side coupled with a second fluid stream claim 8 , and wherein a temperature difference between the fluid streams enables the thermoelectric generator to power the heat pump.10. A heat transfer system for a gas turbine engine comprising:a first heat sink;a first cooler coupled to the first ...

FUEL HEAT EXCHANGER WITH LEAK MANAGEMENT

A cooling air system for use in a gas turbine engine includes a microchannel fuel-air heat exchanger. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine. 1. A fuel injector for a gas turbine engine , the injector comprisinga nozzle configured to discharge fuel into a combustion chamber included in the gas turbine engine,a stem coupled to the nozzle, and configured to conduct fuel to the nozzle, anda microchannel fuel-air heat exchanger integral with the stem and including a fuel passageway fluidly coupled with the nozzle to pass fuel and a cooling air passageway arranged in thermal communication with the fuel passageway to transmit heat to fuel within the fuel passageway, the fuel-air heat exchanger including a leak management system arranged to guide leakage from at least one of the fuel and cooling air passageways to discourage hazardous condition formation, wherein each of the fuel and cooling air passageways are microchannel passageways.2. The fuel injector of wherein the leak management system includes a leakage capture channel arranged for capturing and guiding leakage.3. The fuel injector of claim 2 , wherein the leakage capture channel is arranged at least partly between fuel passageway and the cooling air passageway.4. The fuel injector of claim 2 , wherein the fuel-air heat exchanger includes an interface wall disposed between and providing thermal communication between the fuel and air passageways claim 2 , the leak capture channel formed at least partially within the interface wall.5. The fuel injector of claim 4 , wherein the interface wall includes at least one communication bridge for thermal communication between the fuel and air passageways claim 4 , the communication bridge extending through the leak capture channel.6. The fuel injector of claim 5 , wherein the at least one communication bridge is formed as a pin arranged to conduct heat ...

FLUID COOLING SYSTEM INTEGRATED WITH OUTLET GUIDE VANE

A fan module for a gas turbine engine is disclosed herein. The fan module includes a fan, a plurality of outlet guide vanes, and a fluid cooling system. The fan is adapted to rotate about a central axis to pass air at least in part aftward along the central axis and around an engine core of the gas turbine engine. The outlet guide vanes are spaced aft of the fan along the central axis and configured to receive the air passed aftward along the central axis by the fan. The fluid cooling system is configured to transfer heat from a fluid to the air from the fan to cool the fluid.

Aircraft system

An aircraft capable of operating at a variety of speeds includes a power plant and an auxiliary turbine. The auxiliary turbine can be a ram air turbine used to expand and cool an airflow and provide work. The cooled airflow from the auxiliary turbine can be used in a heat exchange device such as, but not limited to, a fuel/air heat exchanger. In one embodiment the cooled airflow can be used to exchange heat with a compressor airflow being routed to cool a turbine. Work developed from the auxiliary turbine can be used to power a heating device and rotate a device to add work to a shaft of the aircraft power plant. In one form the aircraft power plant is a gas turbine engine and the work developed from the auxiliary turbine is used to heat a combustor flow or to drive a shaft that couples a turbine and a compressor.

Gas turbine engine flow duct having integrated heat exchanger

A gas turbine engine flow duct comprising a flow duct disposed along an engine centerline of the gas turbine engine and defining a stream flow passage, and first and second rows of heat exchangers disposed along the engine centerline of the gas turbine engine and integrated in the flow duct in fluid communication with the stream flow passage of the flow duct.

Ram air thermal management system

An aircraft may have a heat generating component and an engine, at least one of which generates a heat load, and a thermal management system to cool the heat load. The engine may have a duct and an engine fan configured to draw an inlet air stream into an inlet portion of the duct, where at least a portion of the inlet air stream may be used as an engine air stream. The thermal management system may include a cooling circuit configured to circulate a fluid through the heat load such that at least a portion of it may be transferred to the fluid, a heat exchanger configured to enable heat transfer between the fluid and a cooling air stream, and a pumping device. The pumping device may be configured to draw the cooling air stream through the heat exchanger and into a portion of the engine air stream.

AIRCRAFT VEHICLE THERMAL MANAGEMENT SYSTEM AND METHOD

An air vehicle power and thermal management system includes an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system and an electrically powered load. The controller is configured to direct the power to create a first duration cooling power to the cooling system to cool the engine fuel cooling medium over a first power time period. The controller is configured to shift the power to reduce the first duration cooling power to create a load electrical power to drive the electrically powered load over a second power time period. By operation of the controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period. 1. An air vehicle power and thermal management system comprising:an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system to cool an engine fuel cooling medium and an electrically powered load to drive the electrically powered load; direct the power to create a first duration cooling power from the gas turbine engine to the cooling system to cool the engine fuel cooling medium, wherein the first duration cooling power is directed over a first power time period, and', 'shift the power to reduce the first duration cooling power to create a load electrical power to the electrically powered load to drive the electrically powered load over a second power time period that follows the first power time period, and', 'wherein by operation of the aircraft controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period by ...

AIRCRAFT VEHICLE THERMAL MANAGEMENT SYSTEM AND METHOD

An air vehicle power and thermal management system includes an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system and an electrically powered load. The controller is configured to direct the power to create a first duration cooling power to the cooling system to cool the engine fuel cooling medium over a first power time period. The controller is configured to shift the power to reduce the first duration cooling power to create a load electrical power to drive the electrically powered load over a second power time period. By operation of the controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period. 1. An air vehicle power and thermal management system comprising:an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system to cool an engine fuel cooling medium and an electrically powered load to drive the electrically powered load; direct the power to create a first duration cooling power from the gas turbine engine to the cooling system to cool the engine fuel cooling medium, wherein the first duration cooling power is directed over a first power time period, and', 'shift the power to reduce the first duration cooling power to create a load electrical power to the electrically powered load to drive the electrically powered load over a second power time period that follows the first power time period, and', 'wherein by operation of the aircraft controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period by ...

Heat exchanger integrated with a gas turbine engine and adaptive flow control

One embodiment of an engine may include an enclosure surrounding an engine having an engine centerline, and the enclosure defining a passage for a cold-side airflow. The engine may also include one or more contiguous heat exchangers having a cold side inlet surface receiving a cold-side airflow. The heat exchanger may be disposed within the passage, such that a surface normal relative to the cold side inlet surface is offset by at least 30 degrees from the engine centerline.

HEAT EXCHANGER INTEGRATED WITH FUEL NOZZLE

A cooling air system for use in a gas turbine engine includes a fuel-air heat exchanger. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine. 1. A fuel injector for a gas turbine engine , the injector comprisinga nozzle configured to discharge fuel into a combustion chamber included in the gas turbine engine,a stem coupled to the nozzle and configured to conduct fuel to the nozzle, anda fuel-air heat exchanger integral with the stem to provide cooled air at a discrete circumferential location when the fuel nozzle is used in the gas turbine engine, the fuel-air heat exchanger including a fuel passageway coupled fluidly with the nozzle and a cooling air passageway located along the fuel passageway such that heat is transferred from air flowing through the cooling air passageway to the fuel flowing through the fuel passageway when the fuel nozzle is used in the gas turbine engine.2. The fuel injector of claim 1 , wherein the fuel-air heat exchanger includes at least one of an insert arranged in the fuel passageway and an insert arranged in the cooling air passageway.3. The fuel injector of claim 2 , wherein the insert arranged in the fuel passageway is corrugated and/or the insert arranged in the cooling air passageway is corrugated.4. The fuel injector of claim 1 , wherein the fuel-air heat exchanger includes a plurality of pins positioned in the cooling air passageway such that air moving through the cooling air passageway flows over the pins.5. The fuel injector of claim 1 , wherein the fuel-air heat exchanger includes a plurality of plates that are etched and bonded together to form the fuel passageway and the cooling air passageway.6. The fuel injector of claim 1 , wherein the fuel-air heat exchanger includes a plurality of nested tubes having different diameters that cooperate to define the fuel passageway and the cooling air passageway.7. The fuel ...

COOLED GAS TURBINE ENGINE COMPONENT

A gas turbine engine component is disclosed having a cooling fluid passageway that provides relatively cool fluid to a surface of the gas turbine engine component. The cooling fluid passageway can be shaped in cross section to reduce a stress present in the gas turbine engine component. One form of the shape is non-circular. The gas turbine engine component can be formed such that an overhanging material otherwise formed by the intersection of a cooling fluid passageway and a surface of the gas turbine engine component is absent. The gas turbine engine component can also have a depression formed near the surface of the gas turbine engine component such that the cooling fluid passageway exits into an upstream portion and a downstream portion of the depression. 1. An apparatus comprising:a film cooled gas turbine engine component having a passage defined by a passage surface for flowing a cooling air to a hot side, the passage extending at an acute angle to the hot side surface and the passage surface having a hot side surface side; andan edge break formed in the gas turbine engine component and having an edge break surface extending from the hot side surface to the hot side surface side of the passage, the side of the passage opposite the hot side surface side continuing to the hot side, the edge break representing an overhanging material absent from the gas turbine engine component when an edge break surface is formed.2. The apparatus of claim 1 , wherein the edge break surface forms an angle of 45 degrees relative to the hot side surface.3. The apparatus of claim 1 , wherein the edge break surface extends to lateral edges of the passage.4. The apparatus of claim 1 , wherein the passage includes a non-circular cross section.6. The apparatus of claim 5 , wherein in the pair {m claim 5 ,n} claim 5 , n does not equal 2 and m does not equal 1.7. The apparatus of claim 1 , which further includes a ramp formed between a bottom edge of the edge break surface relative to ...

GAS TURBINE ENGINE WITH MICROCHANNEL COOLED ELECTRIC DEVICE

A gas turbine engine includes an electrical device and a microchannel cooling system in communication with the electrical device to remove heat. 1. A gas turbine engine for use in an aircraft , the engine comprisinga low pressure spool including a fan arranged at a forward end of the engine, a low pressure turbine rotor arranged at an aft end of the engine, a low pressure drive shaft extending along an axis and rotationally coupling the fan to receive driven rotation from the low pressure turbine rotor,a high pressure spool including a compressor rotor, a high pressure turbine rotor, and a high pressure drive shaft extending along the axis and rotationally coupling the compressor rotor to receive driven rotation from the high pressure turbine rotor, andan electric device including a stator having an annular core, a rotor rotationally coupled to the low pressure drive shaft and disposed about the stator in electromagnetic communication, and a microchannel cooling system arranged radially inward of the stator in thermal communication with the annular core to pass coolant for removing heat from the stator.2. The gas turbine engine of claim 1 , wherein the microchannel cooling system includes a housing and a network of micropassageways within the housing.3. The gas turbine engine of claim 2 , wherein the micropassageways include inlet passageways for receiving coolant and outlet passageways for discharging heated coolant.4. The gas turbine engine of claim 3 , wherein each inlet passageway is connected with at least one of the outlet passageways by at least one transfer section to pass coolant in thermal communication with the annular core.5. The gas turbine engine of claim 3 , wherein the inlet and outlet passageways are arranged in alternating sequence in the circumferential direction.6. The gas turbine engine of claim 1 , wherein the stator includes electrical windings disposed radially outward of the annular core.7. The gas turbine engine of claim 1 , wherein the ...

Cooling System In A Hybrid Electric Propulsion Gas Turbine Engine For Cooling Electrical Components Therein

A cooling system is provided in a hybrid electric propulsion gas turbine engine for cooling electrical components therein. The cooling system includes an electrical component disposed in proximity to a power generation component in the hybrid electric propulsion gas turbine engine. The cooling system further includes a vapor chamber having an evaporator portion and a condenser portion, wherein the evaporator portion is disposed adjacent to and in thermal communication with the electrical component to transfer heat away from the electrical component. The vapor chamber includes biphasic working fluid therein that transitions between liquid and gaseous states as the working fluid flows proximal to the condenser portion and the evaporator portion respectively.

COOLING SYSTEM IN HYBRID ELECTRIC PROPULSION GAS TURBINE ENGINE

A cooling system in a hybrid electric propulsion gas turbine engine is provided for cooling electrical components therein. The cooling system includes an electrical component disposed in proximity to an aircraft power generation component in the hybrid electric propulsion gas turbine engine such that the electrical component is thermally heated by the aircraft power generation component. A loop heat pipe structure is in thermal communication with the electrical component to transfer heat away from the electrical component. Wherein the loop heat pipe includes an evaporator portion, a condenser portion, a first pipe to supply a biphasic working fluid in a liquid state to the evaporator portion, and a second pipe to return the biphasic working fluid in a gaseous state to the condenser portion. 1. A cooling system in a hybrid electric propulsion gas turbine engine for cooling electrical components therein , the cooling system comprising:an electrical component disposed in proximity to an aircraft power generation component in the hybrid electric propulsion gas turbine engine such that the electrical component is thermally heated by the aircraft power generation component; an evaporator portion;', 'a condenser portion;', 'a first pipe to supply a biphasic working fluid in a liquid state to the evaporator portion; and', 'a second pipe to return the biphasic working fluid in a gaseous state to the condenser portion., 'a loop heat pipe structure in thermal communication with the electrical component to transfer heat away from the electrical component, wherein the loop heat pipe comprises2. The cooling system of claim 1 , wherein the condenser portion is disposed in proximity to a heat sink section to cool the working fluid from the gaseous state to the liquid state.3. The cooling system of claim 2 , wherein the heat sink section comprises at least one of engine oil claim 2 , engine fuel claim 2 , fan stream air claim 2 , ram stream air claim 2 , an engine nacelle claim 2 , ...

Integrated aircraft cooling system

An integrated aircraft cooling system comprising a turbine engine and a plurality of secondary cooling streams, wherein the turbine engine core exhaust stream drives the plurality of secondary cooling streams. A core exhaust stream outlet has a periphery and a composite secondary outlet positioned around the periphery of the core exhaust stream outlet, and the composite secondary outlet is segregated between the plurality of secondary streams. Each of the secondary flows are in fluid isolation from each other upstream of the composite secondary outlet.

ENHANCED HEAT SINK AVAILABILITY ON GAS TURBINE ENGINES THROUGH THE USE OF COOLERS

A cooling assembly for a gas turbine engine may include a heat source, a heat sink, and a heat pump coupled to the heat source and the heat sink, wherein the heat pump may be configured to convey a quantity of heat from the heat source to the heat sink. The heat pump may be mounted to an inner surface of an inner fan casing, and the heat sink may be mounted to an outer surface of the inner fan casing such that heat may convey from the heat sink to a bypass air stream passing over the inner fan casing through or past the heat sink. 120-. (canceled)21. A cooling assembly for a gas turbine engine comprising:a heat source;a heat sink; anda heat pump coupled to the heat source and the heat sink, wherein the heat pump is configured to convey a quantity of heat from the heat source to the heat sink;wherein the heat pump is mounted to an inner surface of an inner fan casing, the heat sink is mounted to an outer surface of the inner fan casing such that heat conveys from the heat sink to a bypass air stream passing over the inner fan casing through or past the heat sink.22. The cooling assembly of claim 21 , wherein the heat source is a heat exchanger having a material passing therethrough.23. The cooling assembly of claim 22 , wherein the material is a heat transfer fluid or gas.24. The cooling assembly of claim 23 , wherein the heat transfer fluid or gas is one of compressor air claim 23 , oil claim 23 , fuel claim 23 , or coolant.25. The cooling assembly of claim 21 , wherein the heat pump is a generator.26. The cooling assembly of claim 21 , wherein the heat sink is a surface cooler heat exchanger.27. A heat transfer system for a gas turbine engine comprising:a surface cooler heat exchanger;a generator; anda heat exchanger heat source coupled to the generator, wherein the heat exchanger heat source is configured to remove heat from a material passing therethrough, and wherein the generator is configured to cause a quantity of heat to pass from the heat exchanger heat ...

Gas turbine engine flow duct having integrated heat exchanger

A gas turbine engine flow duct comprising a flow duct disposed along an engine centerline of the gas turbine engine and defining a stream flow passage, and first and second rows of heat exchangers disposed along the engine centerline of the gas turbine engine and integrated in the flow duct in fluid communication with the stream flow passage of the flow duct.

THERMAL MANAGEMENT SYSTEM WITH DUAL-USE SERIAL THERMAL ENERGY STORAGE FOR SYSTEM SIZE REDUCTION

Thermal management systems for cooling high-power, low-duty-cycle thermal loads by rejecting heat from the thermal loads to the ambient environment are provided. The thermal management systems include a two-phase pump loop in fluid communication with a vapor compression system loop, evaporators disposed in parallel between the two-phase pump loop and the vapor compression system loop, and a thermal energy storage loop including a cold-temperature tank and a warm-temperature tank thermally coupled to the two-phase pump loop and the vapor-compression system loop. Methods of transferring heat from one or more thermal loads to an ambient environment are also provided. 1. A thermal management system , comprising:a thermal energy storage (“TES”) loop comprising a TES medium disposed in the TES loop, a TPPL condenser, a first tank, a first liquid pump, a first TES evaporator, a second tank, and a second liquid pump; anda primary fluid flow path comprising a primary fluid disposed in the primary fluid flow path, a two-phase pump loop (“TPPL”), a vapor compression system (“VCS”) loop, an accumulator, and a first-TES-evaporator branch; andwherein the TPPL is configured to cool a primary thermal load, the TPPL comprising a TPPL liquid pump and the TPPL condenser, the TPPL condenser configured to transfer heat from the primary fluid in the TPPL to the TES medium;wherein the VCS loop is configured to transfer heat from the primary fluid in the primary fluid flow path to an ambient environment via a VCS condenser;wherein the accumulator is configured to separate the primary fluid received from the TPPL and the VCS loop into a vapor-phase primary fluid and a liquid-phase primary fluid;wherein the first-TES-evaporator branch comprises the first TES evaporator and is in fluid communication with the TPPL downstream of the accumulator and with the VCS loop upstream of the compressor, wherein the first TES evaporator is configured to transfer heat from the TES medium to the VCS loop; ...

FUEL INJECTION ASSEMBLY FOR GAS TURBINE ENGINE

Fuel injection assemblies for a gas turbine engine includes a fuel injector and heat exchanger in communication via a fuel line cavity to provide heated fuel for combustion. 1. A fuel injection assembly for a gas turbine engine comprisinga high pressure casing defining a high pressure area through which pressurized fluids are passed to a combustion chamber of the gas turbine engine,at least one fuel injector including a head mounted to the high pressure casing of the gas turbine engine, and a stem extending from the head into the high pressure area, and a nozzle connected with the stem to inject fuel into the combustion chamber,at least one fuel-cooled heat exchanger mounted to the high pressure casing and arranged in fluid communication to receive fuel from a fuel supply of the gas turbine engine, the at least one heat exchanger including a core adapted to pass fuel in thermal communication with a heat source to receive heat, anda fuel cavity defined by the high pressure casing, the fuel cavity arranged to receive heated fuel from the at least one heat exchanger and to communicate heated fuel with the head of the at least one fuel injector for injection into the combustion chamber.2. The fuel injection assembly of claim 1 , wherein the fuel cavity is formed as a depression within a wall of the high pressure casing.3. The fuel injection assembly of claim 2 , wherein the high pressure casing includes a cap plate enclosing the fuel cavity.4. The fuel injection assembly of claim 3 , wherein the fuel cavity includes a conduit in communication with each of the at least one heat exchanger and the at least one fuel injector to communicate fuel therebetween claim 3 , the conduit extending within the fuel cavity of the high pressure casing between the at least one heat exchanger and the at least one fuel injector.5. The fuel injection assembly of claim 4 , wherein the cap plate includes at least one opening in communication with the conduit to pass fuel.6. The fuel injection ...

AIRCRAFT SYSTEM

An aircraft capable of operating at a variety of speeds includes a power plant and an auxiliary turbine. The auxiliary turbine can be a ram air turbine used to expand and cool an airflow and provide work. The cooled airflow from the auxiliary turbine can be used in a heat exchange device such as, but not limited to, a fuel/air heat exchanger. In one embodiment the cooled airflow can be used to exchange heat with a compressor airflow being routed to cool a turbine. Work developed from the auxiliary turbine can be used to power a heating device and rotate a device to add work to a shaft of the aircraft power plant. In one form the aircraft power plant is a gas turbine engine and the work developed from the auxiliary turbine is used to heat a combustor flow or to drive a shaft that couples a turbine and a compressor. 1. An apparatus comprising:an aircraft capable of operating over a range of Mach numbers;an aircraft power plant structured to provide power to the aircraft and having a combustor, the aircraft power plant characterized by a thermodynamic cycle; anda ram air turbine that receives a working fluid and that rotates to produce a power when the working fluid traverses therethrough, wherein the ram air turbine is structured to extract work from the working fluid and provide one of heat or work to the thermodynamic cycle of the aircraft power plant.2. The apparatus of which further includes a power device structured to receive power from the ram air turbine.3. The apparatus of claim 2 , wherein the power device is an electric generator claim 2 , and which further includes a heat source located to heat an airflow downstream of an aircraft power plant compressor or turbine claim 2 , the heat source powered by the electric generator.4. The apparatus of claim 1 , wherein the aircraft power plant is a gas turbine engine having a compressor and a turbine claim 1 , and wherein the working fluid discharged from the ram air turbine is used to cool a compressor airflow ...

Aircraft system

An aircraft capable of operating at a variety of speeds includes a power plant and an auxiliary turbine. The auxiliary turbine can be a ram air turbine used to expand and cool an airflow and provide work. The cooled airflow from the auxiliary turbine can be used in a heat exchange device such as, but not limited to, a fuel/air heat exchanger. In one embodiment the cooled airflow can be used to exchange heat with a compressor airflow being routed to cool a turbine. Work developed from the auxiliary turbine can be used to power a heating device and rotate a device to add work to a shaft of the aircraft power plant. In one form the aircraft power plant is a gas turbine engine and the work developed from the auxiliary turbine is used to heat a combustor flow or to drive a shaft that couples a turbine and a compressor.

RAM AIR THERMAL MANAGEMENT SYSTEM

An aircraft may have a heat generating component and an engine, at least one of which generates a heat load, and a thermal management system to cool the heat load. The engine may have a duct and an engine fan configured to draw an inlet air stream into an inlet portion of the duct, where at least a portion of the inlet air stream may be used as an engine air stream. The thermal management system may include a cooling circuit configured to circulate a fluid through the heat load such that at least a portion of it may be transferred to the fluid, a heat exchanger configured to enable heat transfer between the fluid and a cooling air stream, and a pumping device. The pumping device may be configured to draw the cooling air stream through the heat exchanger and into a portion of the engine air stream. 1. A thermal management system for an aircraft having an engine and a heat generating component , at least one of which generates a heat load , the engine having an engine fan configured to draw in an engine inlet air stream , at least a portion of which is to be used as an engine air stream downstream of the engine fan , the thermal management system comprising:a cooling circuit configured to circulate a fluid through the heat load such that at least a portion of the heat load is transferrable to the fluid;a heat exchanger in fluid communication with the cooling circuit, the heat exchanger being configured to enable heat transfer between the fluid and a cooling air stream; anda pumping device configured to draw the cooling air stream through the heat exchanger and into a portion of the engine air stream downstream of the engine fan.2. The thermal management system of wherein the engine includes a first dividing plate downstream of the engine fan claim 1 , the first dividing plate being configured to divide the engine air stream into a core stream and a bypass stream claim 1 , and wherein the pumping device is configured to deposit the cooling air stream from the heat ...

ADAPTIVE TRANS-CRITICAL CO2 COOLING SYSTEMS FOR AEROSPACE APPLICATIONS

A cooling system for an aircraft includes a first cooling circuit having a first evaporator and a second evaporator, and a second cooling circuit having a third evaporator and a fourth evaporator. One of the first and second cooling circuits includes a first set of valves arranged to direct refrigerant through a first cooling sub-circuit, a second cooling sub-circuit, or both the first and second cooling sub-circuits based on ambient conditions. Two of the evaporators are installed on a first side of the aircraft, and the other two of the four evaporators are installed on a second side of the aircraft opposite the first side, and the first and second cooling circuits reject heat, via a heat exchanger, from their respective cooling circuit to air passing into an engine of the aircraft. 1. A cooling system for an aircraft comprising:a first cooling circuit having a first evaporator and a second evaporator; anda second cooling circuit having a third evaporator and a fourth evaporator; one of the first and second cooling circuits includes a first set of valves arranged to direct refrigerant through a first cooling sub-circuit, a second cooling sub-circuit, or both the first and second cooling sub-circuits based on ambient conditions;', 'two of the evaporators are installed on a first side of the aircraft, and the other two of the four evaporators are installed on a second side of the aircraft opposite the first side; and', 'the first and second cooling circuits reject heat, via a heat exchanger, from their respective cooling circuit to air passing into an engine of the aircraft., 'wherein2. The cooling system of claim 1 , wherein the heat exchanger is a first heat exchanger that is configured to reject heat from the first cooling circuit to air passing into the engine of the aircraft; andfurther comprising a second heat exchanger that is configured to reject heat from the second cooling circuit to the air passing into the engine of the aircraft.3. The cooling system of ...

HEAT EXCHANGER INTEGRATED WITH A GAS TURBINE ENGINE AND ADAPTIVE FLOW CONTROL

One embodiment of an engine may include an enclosure surrounding an engine having an engine centerline, and the enclosure defining a passage for a cold-side airflow. The engine may also include one or more contiguous heat exchangers having a cold side inlet surface receiving a cold-side airflow. The heat exchanger may be disposed within the passage, such that a surface normal relative to the cold side inlet surface is offset by at least 30 degrees from the engine centerline. 1. A gas turbine engine , comprising:an engine including an engine centerline and a passage being one of a fan bypass duct, a third stream duct and a ram air duct; andat least one contiguous heat exchanger having a cold side inlet surface receiving a cold-side airflow;wherein the at least one contiguous heat exchanger is disposed within the passage such that a surface normal relative to the cold side inlet surface is offset by at least 30 degrees from the engine centerline.2. The engine of claim 1 , wherein the at least one contiguous heat exchanger includes a pair of contiguous heat exchangers carried by a module.3. The engine of claim 2 , wherein each one of the pair of contiguous heat exchangers is spaced apart from one another.4. The engine of claim 2 , further comprising a variable flow control mechanism located within the passage.5. The engine of claim 1 , wherein the heat exchanger includes at least one of a plurality of guide plates disposed within the heat exchanger claim 1 , a plurality of fins and a plurality of vanes disposed upstream and downstream of the heat exchanger.6. The engine of claim 5 , wherein the module includes an internal passage receiving the cold-side airflow from the pair of contiguous heat exchangers claim 5 , and the module has a plurality of vanes disposed within the internal passage for distributing the cold-side airflow within the internal passage.7. The engine of claim 5 , wherein the module includes a nose cone for directing the cold-side airflow ...

HIGH HEAT FLUX POWER ELECTRONICS COOLING DESIGN

A base plate for cooling a power electronics device is provided, the base plate comprising cooling fins, the base plate configured to receive the power electronics device directly above the cooling fins, the cooling fins integral to the base plate, the base plate configured to conduct a liquid coolant past the cooling fins 1. A method of manufacturing a base plate for a semiconductor , method comprising:forming a plurality of pieces of the base plate from a metal sheet by etching openings in the metal sheet, wherein after etching, a first set of the pieces comprises a plurality of cooling fins and a second set of the pieces comprises openings for passages;stacking the pieces such that each of the pieces comprising the cooling fins is adjacent to a corresponding one of the pieces comprising openings for the passages; andbonding the stacked pieces together to form the base plate or a portion thereof, the cooling fins and the passages between the cooling fins being integral to, and included within, the base plate.2. The method of wherein the bonding is diffusion bonding.3. A method of manufacturing a base plate for a semiconductor claim 1 , method comprising:forming a plurality of pieces of the base plate from a metal sheet by etching openings in the metal sheet, wherein after etching, the pieces comprise a plurality of cooling fins;etching a portion of each of the cooling fins to form a corresponding passage adjacent to the respective one of the cooling fins;stacking the pieces; andbonding the stacked pieces together to form the base plate or a portion thereof, the cooling fins and the passages between the cooling fins being integral to, and included within, the base plate.4. A system for cooling a power electronics device claim 1 , the system comprising: a base plate including cooling fins claim 1 , the base plate configured to receive the power electronics device directly above the cooling fins claim 1 , the cooling fins integral to the base plate claim 1 , the base ...

HEAT EXCHANGER ASSEMBLY FOR A GAS TURBINE ENGINE PROPULSION SYSTEM

A propulsion system including a gas turbine engine is disclosed herein. The propulsion system further includes a heat exchanger arranged outside the gas turbine engine and adapted to cool fluid from the gas turbine engine. 1. A propulsion system for an aircraft , the propulsion system comprisinga gas turbine engine including an engine core defining a central axis and a fan coupled to the engine core, the fan configured to discharge pressurized bypass air that is passed around the engine core through a fan duct that extends along the central axis coaxially with the engine core, anda nacelle mounted to the gas turbine engine, the nacelle including an outer shroud that surrounds at least a portion of the engine core defining a portion of the fan duct and a strut that extends away from the engine core through the fan duct to the outer shroud, anda heat exchanger assembly fluidly coupled to the gas turbine engine to cool fluid or gas from the gas turbine engine and return the cooled fluid or gas to the gas turbine engine, the heat exchanger assembly including an inlet duct having at least a portion positioned in the strut, a heat exchanger housing coupled fluidly to the inlet duct and positioned radially inward of the strut relative to the central axis, heat exchangers housed by the heat exchanger housing, and a valve system that is movable in the heat exchanger housing from a first position arranged to direct pressurized bypass air received from the inlet duct into contact with the heat exchangers to a second position arranged to divert pressurized bypass air received from the inlet duct around the heat exchangers without contacting the heat exchangers.2. The propulsion system of claim 1 , wherein the heat exchanger assembly further includes an outlet duct fluidly coupled to the heat exchangers that extends aft of the strut along the central axis.3. The propulsion system of claim 2 , wherein the engine core is configured to discharge core air that is passed through the ...

Inlet guide vane draw heat exchanger system

A system includes a housing for a gas turbine engine, and a fan disposed in the housing to rotate coaxially with a gas turbine included in the housing. The system also includes an inlet guide vane disposed in the housing in axial alignment with the fan and configured to have an open position where a first flow of air is received by the fan through the inlet guide vane, and a closed position where airflow through the inlet guide vane is obstructed. The system further includes a heat exchanger disposed in a supply passage in fluid communication with a second flow of air received by the fan. The second flow of air is received by the fan via the supply passage with the inlet guide vane in the open position or in the closed position.

HEAT EXCHANGER INTEGRATED WITH FUEL NOZZLE

Fuel-air heat exchange system and methods thereof for use in a gas turbine engine. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine. 1. A gas turbine engine comprisinga power train including a compressor, combustor, and a turbine, the compressor arranged to compress air, the combustor including a combustion chamber arranged to receive compressed air and fuel for combustion to form exhaust products, the turbine including a rotor and blade extending from the rotor to receive exhaust products expanding across the blades to drive rotation of the rotor,an outer casing defining a high pressure cavity adapted to receive compressed air from the compressor, the combustion chamber arranged at least partly within the high pressure cavity, and a nozzle configured to discharge fuel into the combustion chamber,', 'a stem secured to and penetrating through the outer casing, the stem fluidly coupled with the nozzle and configured to conduct fuel to the nozzle, and', 'a microchannel fuel-air heat exchanger integral with the stem and including a body having a fuel passageway fluidly coupled with the nozzle to pass fuel and a cooling air passageway arranged in thermal communication with the fuel passageway to transmit heat to fuel within the fuel passageway,, 'a fuel injector including'}wherein the microchannel heat exchanger is arranged outside of the high pressure cavity and secured with the outer casing.2. The gas turbine engine of claim 1 , wherein the outer casing includes a port defined therethrough and the stem is arranged to extend through the port.3. The gas turbine engine of claim 2 , wherein the outer casing includes a transfer passage arranged in fluid communication with each of the high pressure cavity and the cooling air passage to direct compressed air from the high pressure cavity into the cooling air passage.4. The gas turbine engine of claim 3 , wherein the ...

ADAPTIVE TRANS-CRITICAL CO2 COOLING SYSTEMS FOR AEROSPACE APPLICATIONS

A cooling system includes a heat exchanger through which a refrigerant flows, and which rejects heat to a fluid, an evaporator, a first circuit having an expansion device, a second circuit having an expansion machine coupled to a compressor, and a set of valves arranged to direct the refrigerant through the first circuit, the second circuit, or both the first and second circuits based on ambient conditions. 1. A cooling system comprising:a heat exchanger through which a refrigerant flows, and which rejects heat to a fluid;an evaporator;a first circuit having an expansion device;a second circuit having an expansion machine coupled to a first compressor; anda set of valves arranged to direct the refrigerant through the first circuit, the second circuit, or both the first and second circuits based on ambient conditions.2. The cooling system of claim 1 , wherein the cooling system is for an aircraft claim 1 , and the ambient conditions are defined by an operating condition of the aircraft.3. The cooling system as claimed in claim 1 , further comprising:the first compressor configured to compress the refrigerant to a first pressure; andan evaporator configured to evaporate the refrigerant;wherein the expansion machine is a first turbine that is rotationally coupled to the compressor through a shaft, and the expansion device is an expansion valve.4. The cooling system as claimed in claim 3 , further comprising:a second compressor configured to, prior to entering the first compressor, compress the refrigerant to a second pressure that is less than the first pressure; anda second heat exchanger configured to cool the refrigerant prior to entering the first compressor but after exiting the second compressor.5. The cooling system as claimed in claim 4 , further comprising at least one of: pass the refrigerant therethrough in a first direction and prior to entering one of the first and second compressors; and', 'pass the refrigerant therethrough in a second direction that is ...

THERMAL MANAGEMENT SYSTEM INCLUDING TWO-PHASED PUMP LOOP AND THERMAL ENERGY STORAGE

A thermal management system for regulating dissipation of multiple thermal loads during operation of an apparatus includes a two-phase pump loop (TPPL), a vapor cycle system (VCS), and a liquid thermal energy storage (TES) system integrated together to maintain the apparatus at a constant temperature. The TPPL is configured to remove heat from the apparatus; the TES system is configured to provide thermal energy storage and temperature regulation; and the VCS is configured to transfer heat to the environment. The multiple thermal loads include a primary thermal load in the form of heat from the apparatus and a secondary thermal load in the form of at least one of a housekeeping thermal load or a power electronics thermal load. The primary and secondary loads are at different temperatures with each being independently selected to be transient or steady state. 1. A thermal management system for regulating dissipation of multiple thermal loads during operation of an apparatus , the thermal management system comprising a two-phase pump loop (TPPL) , a vapor cycle system (VCS) , and a thermal energy storage (TES) system;wherein the TPPL, the VCS, and the TES system are integrated together to maintain the apparatus at a constant temperature, the TPPL is configured to remove heat from the apparatus, the TES system is configured to provide thermal energy storage for some or all of the multiple thermal loads and temperature regulation to at least one of the multiple thermal loads, and the VCS is configured to transfer heat to the environment;wherein the multiple thermal loads comprise a primary thermal load in the form of heat from the apparatus and a secondary thermal load in the form of at least one of a housekeeping thermal load or one or more thermal loads associated with conditioning, distributing, or converting energy;wherein the primary and secondary loads are at different temperatures with each being independently selected to be transient or steady state.2. The ...

THERMAL ENERGY STORAGE AND HEAT REJECTION SYSTEM

A cooling system is provided including a two-phase pump loop and a vapor compression system. The two-phase pump loop cools a thermal load with a first coolant. The vapor compression system is configured to circulate a second coolant. The vapor compression system includes a liquid vapor separator which separates the second coolant into a liquid portion and a gaseous portion. The liquid vapor separator is a thermal energy storage for the two-phase pump loop. A condenser of the two-phase pump loop transfers heat from the first coolant to the liquid portion of the second coolant in the liquid-vapor separator. 1. A cooling system comprising:a two-phase pump loop comprising a pump, an evaporator configured to evaporate a first coolant supplied by the pump, a condenser configured to condense the first coolant evaporated by the evaporator, and an accumulator configured to deliver the first coolant condensed in the condenser to the pump;a thermal energy storage configured to deliver a liquid portion of a second coolant to the condenser of the two-phase pump loop, wherein the thermal energy storage is configured to separate the second coolant into the liquid portion and a gaseous portion; anda vapor compression system configured to circulate the second coolant, the vapor compression system comprising a compressor and a cooler, wherein the thermal energy storage is configured as a liquid-vapor separator of the vapor compression system, the compressor is configured to compress the gaseous portion of the second coolant from the thermal energy storage, and the cooler is configured to cool the gaseous portion compressed by the compressor.2. The cooling system of claim 1 , wherein the vapor compression system is a trans-critical vapor compression system.3. The cooling system of claim 1 , wherein the vapor compression system is detachable from the thermal energy storage claim 1 , and wherein the thermal energy storage is still operable while vapor compression system is detached.4. ...

Tight temperature control at a thermal load with a two phase pumped loop, optionally augmented with a vapor compression cycle

A two-phase pump loop (TPPL) for dissipating a thermal load during operation of an apparatus includes a coolant, a vapor/liquid receiver, a pump, an evaporator, a condenser, a valve (V1) configured to regulate a pressure at an outlet of the condenser; a valve (V2) having a control set point set equivalent to a low pressure (PL) measured in the vapor/liquid receiver; and a controller configured to control the set points of V1 and V2. The TPPL is configured to cool the thermal load with tight control of the temperature of the coolant that is cooling the apparatus. The TPPL may be combined with a vapor cycle system (VCS) to provide a thermal management system with the VCS being configured to use the same or different coolant than the TPPL.

Aircraft vehicle thermal management system and method

An air vehicle power and thermal management system includes an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system and an electrically powered load. The controller is configured to direct the power to create a first duration cooling power to the cooling system to cool the engine fuel cooling medium over a first power time period. The controller is configured to shift the power to reduce the first duration cooling power to create a load electrical power to drive the electrically powered load over a second power time period. By operation of the controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period.

BLEED EXPANDER COOLING WITH TURBINE

An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air. 1. A thermal management system , comprising:a first heat exchanger comprising a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air from the first heat exchanger;an air starter turbine comprising a turbine inlet fluidically coupled to the bleed air outlet of the first heat exchanger and a turbine outlet, wherein the air starter turbine is configured to drive a shaft mechanically coupled to the air starter turbine in response to expansion of the cooled bleed air through the air turbine starter; anda second heat exchanger comprising an expanded bleed air inlet fluidically coupled to the turbine outlet, wherein the second heat exchanger is configured to extract heat from at least one heat source using the expanded bleed air.2. The thermal management system of claim 1 , wherein the at least one heat source comprises at least one of an electric machine claim 1 , an electrical power distribution system claim 1 , a power conversion system claim 1 , power electronics claim 1 , digital electronics claim 1 , or an environmental control system.3. The thermal management system of claim 1 , further comprising a cooling fluid circuit claim 1 , wherein the second heat exchanger is configured to ...

Systems and Methods for Cooling Electronics and Electrical Machinery in a Hybrid Electric Aircraft

Systems and methods for cooling electrical components disposed in a jet engine. An example system includes an evaporation chamber configured to contain the electrical components in contact with a coolant liquid. The coolant vapor formed during the heat transfer from the electrical components to the coolant liquid flows to a condenser assembly having a fuel-cooled condenser and an air-cooled condenser. The air-cooled condenser cools the coolant vapor to condensation using either fan stream air or engine bleed air from the intermediate pressure compressor or the high pressure compressor. An air cycle machine cools the engine bleed air. A controller may be used to select a coolant source for condensing the coolant vapor based on operating conditions of the aircraft. Spent air from the air-cooled condenser may be recycled back to the engine for engine cooling, added thrust, oil sump buffering, oil or fuel cooling, or blade tip clearance control. 1. A system for cooling an electrical component in association with a gas turbine engine , the system comprising:an evaporation chamber configured to contain the electrical component and a cooling liquid in contact with the electrical component, the evaporation chamber comprising a liquid input port and a vapor output port, where the cooling liquid evaporating while cooling the electrical component emits a coolant vapor via the vapor output port and receives condensed coolant liquid via the liquid input port;a condenser assembly configured to receive the coolant vapor from the vapor output port of the evaporation chamber and to effect condensation of the coolant vapor using a cooling air flow; andan air cycle machine configured to receive an engine bleed air to use as the cooling air flow, the air cycle machine comprising a heat exchanger and a turbine to cool the engine bleed air before the engine bleed air flows to the condenser assembly.2. The system of where the evaporation chamber includes a liquid output port and a pumped ...

X-ray tube anode cooling device and systems incorporating same

An anode target for use within an x-ray generating device including a target frame having an inner surface and an outer surface and a thermal energy transfer device. The thermal energy transfer device including a heat exchanger having an inner surface and an outer surface, at least a portion of the outer surface of the heat exchanger positioned adjacent to at least a portion of the inner surface of the target frame; a cooling medium circulating through the heat exchanger for convectively cooling the anode target; and a thermal coupling medium disposed between the inner surface of the target frame and the outer surface of the heat exchanger, the thermal coupling medium thermally coupling the target frame with the heat exchanger while permitting relative motion between the target frame and the heat exchanger.

X-RAY TUBES AND X-RAY SYSTEMS COMPRISING A THERMAL GRADIENT DEVICE

Un dispositif de transfert d'énergie thermique utilisé avec un dispositif de génération de rayons X comprend un ensemble d'anode (40) comportant une cible (48), un ensemble de cathode à une certaine distance de l'ensemble d'anode (40), configuré pour émettre des électrons frappant la cible (48), pour produire des rayons X et de la chaleur, et un arbre rotatif (58) supporté par un ensemble de palier (50). Le dispositif comprend un dispositif à gradient thermique (72) au voisinage d'une extrémité de l'arbre (58) et en communication thermique avec celle-ci, le dispositif à gradient thermique (72) transférant de la chaleur pour l'éloigner de cette extrémité de l'arbre (58), et une structure à ailettes (80) au voisinage du dispositif à gradient thermique (72) et en communication thermique avec celui-ci, la structure à ailettes (80) refroidissant par convexion le dispositif à gradient thermique (72). A thermal energy transfer device used with an X-ray generating device includes an anode assembly (40) having a target (48), a cathode assembly at a distance from the anode assembly (40 ), configured to emit electrons striking the target (48), to produce X-rays and heat, and a rotating shaft (58) supported by a bearing assembly (50). The device includes a thermal gradient device (72) adjacent to and in thermal communication with one end of the shaft (58), the thermal gradient device (72) transferring heat away from it this end of the shaft (58), and a fin structure (80) in the vicinity of the thermal gradient device (72) and in thermal communication therewith, the fin structure (80) convectionally cooling the device thermal gradient (72).

Gas turbine engine and heat exchange system

A gas turbine engine and with a heat exchange system is disclosed wherein the heat exchange system provides adaptive cooling to the object of cooling in the gas turbine. The gas turbine engine comprises an engine core, a first fan operable to generate a first air stream at a first pressure, a second fan operable to receive a portion of the first air stream and generate a second air stream at a second pressure, a fan bypass duct to receive a portion of the second air stream, and another fan bypass duct to receive a portion of the first air stream. The heat exchange system is configurable to cool the object of cooling using one of or both of the first and the second air stream to provide adaptive cooling.

Heat exchanger integrated with a gas turbine engine and adaptive flow control

One embodiment of an engine may include an enclosure surrounding an engine having an engine centerline, and the enclosure defining a passage for a cold-side airflow. The engine may also include one or more contiguous heat exchangers having a cold side inlet surface receiving a cold-side airflow. The heat exchanger may be disposed within the passage, such that a surface normal relative to the cold side inlet surface is offset by at least 30 degrees from the engine centerline.

X-RAY TUBES AND X-RAY SYSTEMS HAVING A THERMAL GRADIENT DEVICE

Systems and methods for cooling electronics and electrical machinery in a hybrid electric aircraft

Systems and methods for cooling electrical components disposed in a jet engine. An example system includes an evaporation chamber configured to contain the electrical components in contact with a coolant liquid. The coolant vapor formed during the heat transfer from the electrical components to the coolant liquid flows to a condenser assembly having a fuel-cooled condenser and an air-cooled condenser. The air-cooled condenser cools the coolant vapor to condensation using either fan stream air or engine bleed air from the intermediate pressure compressor or the high pressure compressor. An air cycle machine cools the engine bleed air. A controller may be used to select a coolant source for condensing the coolant vapor based on operating conditions of the aircraft. Spent air from the air-cooled condenser may be recycled back to the engine for engine cooling, added thrust, oil sump buffering, oil or fuel cooling, or blade tip clearance control.

Adaptive trans-critical co2 cooling systems for aerospace applications

Liquid metal cooled anode for an X-ray tube

A system and method are proposed for cooling the anode of an X-ray tube. A bearing shaft associated with the anode has an associated single rotating seal there around, and contains a liquid metal. A primary liquid metal flow path is used to transfer heat from the anode, and a secondary liquid metal flow path is provided to seal the single rotating seal. Accordingly, the present invention provides an effective means for containing liquid metal in the bearing shaft of an anode assembly, and using the liquid metal to cool the anode of the X-ray tube.

Thermal management system including two-phased pump loop and thermal energy storage

A thermal management system for regulating dissipation of multiple thermal loads during operation of an apparatus includes a two-phase pump loop (TPPL), a vapor cycle system (VCS), and a liquid thermal energy storage (TES) system integrated together to maintain the apparatus at a constant temperature. The TPPL is configured to remove heat from the apparatus; the TES system is configured to provide thermal energy storage and temperature regulation; and the VCS is configured to transfer heat to the environment. The multiple thermal loads include a primary thermal load in the form of heat from the apparatus and a secondary thermal load in the form of at least one of a housekeeping thermal load or a power electronics thermal load. The primary and secondary loads are at different temperatures with each being independently selected to be transient or steady state.

Enhanced heat sink availability on gas turbine engines through the use of coolers

A cooling assembly for a gas turbine engine including a heat source at a first temperature, a heat sink at a second temperature, and a heat pump coupled to the first heat source and the first heat sink. The heat pump is configured to convey a quantity of heat from the heat source through the heat pump and to the heat sink.

Gas turbine engine and heat exchange system

Adaptive trans-critical CO2 cooling systems for aerospace applications

A cooling system for an aircraft includes a first cooling circuit having a first evaporator and a second evaporator, and a second cooling circuit having a third evaporator and a fourth evaporator. One of the first and second cooling circuits includes a first set of valves arranged to direct refrigerant through a first cooling sub-circuit, a second cooling sub-circuit, or both the first and second cooling sub-circuits based on ambient conditions. Two of the evaporators are installed on a first side of the aircraft, and the other two of the four evaporators are installed on a second side of the aircraft opposite the first side, and the first and second cooling circuits reject heat, via a heat exchanger, from their respective cooling circuit to air passing into an engine of the aircraft.

Rotating anode X-ray tube for mammography, has heat pipe that transfers thermal energy away from anode assembly target, through heat conducting liquid filled near its condenser end

Cathode and anode assemblies (42,40) separated by preset distances, are arranged inside a vacuum vessel (44). A heat pipe (70) supported with respect to the assembly (40), has heat conducting liquid filled near condenser end through which thermal energy is transferred away from target (60) of assembly (40). Independent claims are also included for the following: (a) Method for dissipating heat from an anode; (b) X-ray tube assembling method

Aircraft vehicle thermal management system and method

An air vehicle power and thermal management system includes an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system and an electrically powered load. The controller is configured to direct the power to create a first duration cooling power to the cooling system to cool the engine fuel cooling medium over a first power time period. The controller is configured to shift the power to reduce the first duration cooling power to create a load electrical power to drive the electrically powered load over a second power time period. By operation of the controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period.

Thermoelectrically controlled X-ray detector array statement regarding federally sponsored research

Disclosed is an X-ray detector assembly for use in a computed tomography system. The X-ray detector assembly comprises an array of detector cells coupled between two rails. A thermoelectric cooler is coupled to an end of each of the rails, and is controlled to alternatively heat or cool the detector array to maintain the array in a substantially isothermal and thermally stable condition. The detector assembly preferably includes both passive and active cooling devices and insulation materials for controlling the temperature of the detector assembly. An electrical heater coupled at the center of the detector array can be used in conjunction with the TEC's to control the temperature profile of the detector array, and to minimize changes in the temperature gradients.

Heat exchanger for a gas turbine engine propulsion system

A propulsion system including a gas turbine engine is disclosed herein. The propulsion system further includes a heat exchanger arranged outside the gas turbine engine and adapted to cool fluid from the gas turbine engine.

Thermal energy storage and heat rejection system

A cooling system is provided including a two-phase pump loop and a vapor compression system. The two-phase pump loop cools a thermal load with a first coolant. The vapor compression system is configured to circulate a second coolant. The vapor compression system includes a liquid vapor separator which separates the second coolant into a liquid portion and a gaseous portion. The liquid vapor separator is a thermal energy storage for the two-phase pump loop. A condenser of the two-phase pump loop transfers heat from the first coolant to the liquid portion of the second coolant in the liquid-vapor separator.

Thermal filter for an x-ray tube window

A thermal energy storage and transfer assembly is disclosed for use in electron beam generating devices that generate residual energy. The residual energy comprises radiant thermal energy and kinetic energy of back scattered electrons. The thermal energy storage and transfer assembly absorbs and stores an amount of the residual energy to reduce the heat load on other components in the electron beam generating device. The thermal energy storage and transfer device comprises a body portion of a sufficient thermal capacity to permit the rate of transfer of the amount of the residual energy absorbed into the assembly to substantially exceed the rate of transfer of the amount of the residual energy out of the assembly. The assembly also comprises a heat exchange chamber filled with a circulating fluid that transfers the thermal energy out of the assembly. Additionally, in an x-ray generating device, an x-ray transmissive filter suitable for absorbing residual energy is positioned between the anode and an x-ray transmissive window. The filter reduces the exposure of the window to the residual energy. The filter may additionally comprise a coating layer that further reduces the exposure of the window to the residual energy.

Tight temperature control at a thermal load with a two phase pumped loop, optionally augmented with a vapor compression cycle

A two-phase pump loop (TPPL) for dissipating a thermal load during operation of an apparatus includes a coolant, a vapor/liquid receiver, a pump, an evaporator, a condenser, a valve (V 1 ) configured to regulate a pressure at an outlet of the condenser; a valve (V 2 ) having a control set point set equivalent to a low pressure (P L ) measured in the vapor/liquid receiver; and a controller configured to control the set points of V 1 and V 2 . The TPPL is configured to cool the thermal load with tight control of the temperature of the coolant that is cooling the apparatus. The TPPL may be combined with a vapor cycle system (VCS) to provide a thermal management system with the VCS being configured to use the same or different coolant than the TPPL.

Aircraft vehicle thermal management system and method

An air vehicle power and thermal management system includes an aircraft controller structured to distribute power provided by a gas turbine engine between a cooling system and an electrically powered load. The controller is configured to direct the power to create a first duration cooling power to the cooling system to cool the engine fuel cooling medium over a first power time period. The controller is configured to shift the power to reduce the first duration cooling power to create a load electrical power to drive the electrically powered load over a second power time period. By operation of the controller to shift the power from the first duration cooling power to the load electrical power a second duration cooling power is provided to the cooling system to cool the electrically powered load using the engine fuel cooling medium that was cooled during the first power time period.

Solid-state CT system and method

Es ist eine Festkörper-Röntgenquelle (14) für ein Computertomographie-(CT-)Abbildungssystem (10) dargestellt. Die Röntgenquelle (14) weist eine Kathode (58) auf, die vorzugsweise aus einer Vielzahl adressierbarer Elemente gebildet ist. Die Kathode ist in einer Vakuumkammer (74) positioniert, so dass von ihr emittierte Elektronen auf eine Anode (68) treffen, die von der Kathode (58) räumlich entfernt ist. Ein Elektronenstrahl (82) wird gebildet und entlang der Länge der Kathode (58) bewegt. Die Anode (68) ist in einem Kühlblockabschnitt (58) angeordnet und liegt funktionsfähig an einem Röntgentransmissionsfenster (66). Die Anode (68) und das Röntgentransmissionsfenster (66) sind in einem verlängerten Kanal (64) des Kühlblockabschnitts (56) angeordnet.

Gas turbine engine with microchannel cooled electric device

A gas turbine engine includes an electrical device and a microchannel cooling system in communication with the electrical device to remove heat.

Aircraft with thermal energy storage system

A thermal energy system for use with an aircraft includes a cooling loop and a cooler. The cooling loop includes a fluid conduit and a pump configured to move fluid through the fluid conduit to transfer heat from a heat source to the fluid in the fluid conduit to cool the heat source. The cooler includes an air-stream heat exchanger located in a duct and is in thermal communication with the fluid conduit to transfer heat between the fluid in the cooling loop and the air passing through the duct.

Aircraft with thermal energy storage system with bypass control

A thermal energy system for use with an aircraft includes a cooling loop and a cooler. The cooling loop includes a fluid conduit and a pump configured to move fluid through the fluid conduit to transfer heat from a heat source to the fluid in the fluid conduit to cool the heat source. The cooler includes an air-stream heat exchanger located in a duct and is in thermal communication with the fluid conduit to transfer heat between the fluid in the cooling loop and the air passing through the duct.

Bimodal cooling system

Cooling systems and methods of operation are provided. The cooling system may include a two-phase pumped loop (TPPL). The two-phase pumped loop may include, a receiver, a pump downstream from the receiver, a heat load downstream from the pump, a TPPL tee downstream from the heat load, a TPPL check valve downstream from the TPPL tee, and a heat exchanger downstream from the TPPL check valve and upstream from the receiver. The cooling system may further include a vapor cycle system (VCS) loop. The vapor cycle system loop may include the receiver, a compressor downstream from a vapor outlet of the receiver, a compressor check valve downstream from the compressor and upstream of the heat exchanger, the heat exchanger, and the heat load downstream from a liquid outlet of the receiver.

傾斜した回転軸を有するｘ線管とその方法

(57)【要約】 【課題】 機械荷重を均衡させた改良型のＸ線管及び均 衡方法を提供すること。 【解決手段】 コンピュータ断層システム（１０）はガ ントリ（１２）及びＸ線管（１４）を備える。ガントリ （１２）はガントリ回転軸（２４）の周りを回転する。 Ｘ線管（１４）は、ガントリ（１２）に取り付けられて いると共に、Ｘ線管回転軸（８８）を有する回転可能な アセンブリ（９０）を備えている。Ｘ線管回転軸（８ ８）はガントリ回転軸（２４）から傾斜角（θ）だけ角 度変位させている。Ｘ線管（１４）のガントリ回転軸 （２４）の周りでの回転により第１のモーメントが生成 され、また回転可能なアセンブリ（９０）のＸ線管回転 軸（８８）の周りでの回転により第１のモーメントと反 対向きに第２のモーメントが生成される。

Système de Gestion Thermique pour Refroidir des Charges Calorifiques Transitoires à Faible Consommation de Puissance

Des systèmes de gestion thermique pour refroidir des charges thermiques de puissance élevée et à faible cycle de service qui incluent au moins trois sous-systèmes fermés couplés thermiquement sont fournis. Les systèmes de gestion thermique fournis ici incluent une boucle de stockage d’énergie thermique principale incluant un réservoir à température froide et un réservoir à température chaude. Les systèmes de gestion thermique fournis ici incluent également un système de compression à plusieurs étages ou une architecture en cascade d’un système de compression de vapeur à basse température et d’un système de compression de vapeur à haute température. Des procédés de transfert de chaleur d’une ou de plusieurs charge(s) thermique(s) à un environnement ambiant sont également fournis. Figure pour l’abrégé : [Fig. 1]

Système de Gestion Thermique avec Stockage d’Énergie Thermique en Série à Double Usage pour la Réduction de la Dimension du Système

L’invention se rapporte à des systèmes de gestion thermique pour refroidir des charges thermiques à haute puissance et à faible cycle d’utilisation en rejetant de la chaleur des charges thermiques vers l’environnement ambiant. Les systèmes de gestion thermique comprennent une boucle de pompe à deux phases en communication fluidique avec une boucle de système de compression de vapeur, des évaporateurs disposés en parallèle entre la boucle de pompe à deux phases et la boucle de système de compression de vapeur, et une boucle de stockage d’énergie thermique comprenant un réservoir à température froide et un réservoir à température chaude couplés thermiquement à la boucle de pompe à deux phases et à la boucle de système de compression de vapeur. L’invention se rapporte également à des procédés de transfert de chaleur d’une ou de plusieurs charge(s) thermique(s) à un environnement ambiant.

Method for performing design trade-off

A method for performing design trade-off. A plurality of critical to quality parameters corresponding to features of the design are obtained. A plurality of design specifications, each design specification corresponding to one of the critical to quality parameters, are also obtained. A plurality of designs are obtained where each design includes a plurality of design values and each design value corresponds to one of the critical to quality parameters. The design values are compared to the design specifications for each of the plurality of designs. A total score is generated for each design in response to the comparison.

等温状態に保持されるｘ線検出器アレイ

(57)【要約】 （修正有） 【課題】 回転自在なガントリを有するＣＴイメージン グ・システムで使用するＸ線検出器アレイを等温状態に 保持するための装置を提供する。 【解決手段】 複数の検出器セル４０の弓形アレイ１８ を保持し、かつこれらのセルをガントリ１２上に取り付 けるための２つの湾曲させたレール４２ａ、４２ｂを備 える。複数の導管セグメント４６がレールに沿って分布 して配置され、その各々は対応する一群のＸ線検出器セ ルの直近に位置し、その内部には一定量の選択した動作 流体及び多孔質芯構造が封入されている。流体が気体状 態では対流により導管セグメント４６に沿って移動し、 液体状態では毛管現象によって芯構造を通って反対方向 に移動することにより、熱が導管セグメントに沿って伝 達されて、その直近にある検出器セルの間で実質的に等 温状態が保持される。

Système de refroidissement bimodal

Un système de refroidissement peut comprendre une boucle à pompe (106) à deux phases (TPPL). La boucle à pompe (106) à deux phases peut comprendre un récepteur (104), une pompe (106) en aval du récepteur (104), une charge calorifique (102) en aval de la pompe (106), un raccord en T TPPL (112) en aval de la charge calorifique (102), une soupape anti-retour TPPL (114) en aval du raccord en T TPPL (112)et un échangeur de chaleur (122) en aval de la soupape anti-retour TPPL (114) et en amont du récepteur (104). Le système de refroidissement peut en outre comprendre une boucle de système de cycle de vapeur (VCS). La boucle de système de cycle de vapeur peut comprendre le récepteur (104), un compresseur (122) en aval d’une sortie de vapeur du récepteur (104), une soupape anti-retour de compresseur (122) en aval du compresseur (122) et en amont de l’échangeur de chaleur (122), l’échangeur de chaleur (122) et la charge calorifique en aval d’une sortie de liquide du récepteur (104).

Système de Gestion Thermique avec Stockage d’Énergie Thermique en Série à Double Usage pour la Réduction de la Dimension du Système

L’invention se rapporte à des systèmes de gestion thermique pour refroidir des charges thermiques à haute puissance et à faible cycle d’utilisation en rejetant de la chaleur des charges thermiques vers l’environnement ambiant. Les systèmes de gestion thermique comprennent une boucle de pompe à deux phases en communication fluidique avec une boucle de système de compression de vapeur, des évaporateurs disposés en parallèle entre la boucle de pompe à deux phases et la boucle de système de compression de vapeur, et une boucle de stockage d’énergie thermique comprenant un réservoir à température froide et un réservoir à température chaude couplés thermiquement à la boucle de pompe à deux phases et à la boucle de système de compression de vapeur. L’invention se rapporte également à des procédés de transfert de chaleur d’une ou de plusieurs charge(s) thermique(s) à un environnement ambiant.

Aircraft with thermal energy storage system with bypass control

A thermal energy system for use with an aircraft includes a cooling loop and a cooler. The cooling loop includes a fluid conduit and a pump configured to move fluid through the fluid conduit to transfer heat from a heat source to the fluid in the fluid conduit to cool the heat source. The cooler includes an air-stream heat exchanger located in a duct and is in thermal communication with the fluid conduit to transfer heat between the fluid in the cooling loop and the air passing through the duct.

Aeronef avec systeme de stockage d'energie thermique

Système d’énergie thermique (12) pour une utilisation avec un aéronef (10) comprenant une source de chaleur (16), une boucle de refroidissement (20) et un refroidisseur (22). La boucle de refroidissement (20) comprend un conduit de fluide (30) et une pompe (32) configurée pour déplacer un fluide à travers le conduit de fluide (30) pour transférer de la chaleur de la source de chaleur (16) au fluide pour refroidir la source de chaleur (16). Le refroidisseur (22) comprend un échangeur de chaleur à flux d’air (36) situé dans la conduite (34) et en communication thermique avec le fluide dans la boucle de refroidissement (20) pour transférer de la chaleur entre la boucle de refroidissement (20) et l’air conduit à travers la conduite (34). Figure pour l’abrégé : [Fig. 2]

Vaste stof CT systeem en werkwijze.

Vaste stof CT systeem en werkwijze.

High performance x-ray target

A brazed X-ray target (60) includes a metallic cap (64) and a graphite back (62) including a nonlinear record groove attached thereto along a stepped surface. An upper corner joint of the stepped surface is distanced from a cap outer edge (68) and a focal track (82) where the maximum heat is generated during use of the target. The graphite back is extended outward toward the cap outer edge to increase a thermal storage of the graphite, and a recess (74) is formed into the cap to maintain a selected moment of inertia of the target and thereby maintain the rotordynamics of a given X-ray tube.

Inlets for gas turbine engine bypass duct heat exchangers

A bypass duct assembly for a gas turbine engine includes a bypass duct, a heat exchanger assembly, and an inlet cowl. The bypass duct is configured to direct bypass air around an engine core of the gas turbine engine. The heat exchanger assembly includes a heat exchanger located in the bypass duct and configured to transfer heat to the bypass air. The inlet cowl is coupled with the bypass duct and the heat exchanger assembly. The inlet cowl includes a cowl duct configured to collect a first portion of the bypass air and conduct the first portion of the bypass air into the heat exchanger and a flow diverter that guides a second portion of the bypass air around the heat exchanger.

Bypass duct heat exchanger with access panel for gas turbine engines

A heat exchanger assembly for a gas turbine engine includes an access panel configured to be removably coupled with an outer wall of a bypass duct arranged circumferentially around a central axis of the gas turbine engine. A heat exchanger is adapted to receive cooling fluid therein and is coupled with the access panel. A shroud extends between the access panel and the heat exchanger to collect a first portion of a flow of air through the bypass duct and direct the first portion through the heat exchanger.