AIRCRAFT HEAT EXCHANGE SYSTEM INCLUDING A THERMOELECTRIC DEVICE

A heat exchange system for an aircraft includes an aircraft controller for controlling an operation of an aircraft, a thermoelectric device having a low temperature side and a high temperature side. Heat is transferred away from the inlet line to the outlet line through the thermoelectric device when a predetermined condition is met.

AIRCRAFT HEAT EXCHANGE SYSTEM INCLUDING A THERMOELECTRIC DEVICE

A heat exchange system for an aircraft includes an aircraft controller for controlling an operation of an aircraft, a thermoelectric device having a low temperature side and a high temperature side, an inlet line that carries fluid through the low temperature side of the thermoelectric device and to the aircraft controller; and an outlet line that carrier the fluid away from the and aircraft controller through the high temperature side of the thermoelectric device. Heat is transferred away from the inlet line to the outlet line through the thermoelectric device when a predetermined condition is met.

Aircraft heat exchange system including a thermoelectric device

A heat exchange system for an aircraft includes an aircraft controller for controlling an operation of an aircraft, a thermoelectric device having a low temperature side and a high temperature side. Heat is transferred away from the inlet line to the outlet line through the thermoelectric device when a predetermined condition is met.

INTEGRATED FUEL CELL AIRCRAFT PRESSURIZATION AND COOLING SYSTEM

According to one embodiment of this disclosure an integrated fuel cell and environmental control system includes a turbo-compressor. The turbo-compressor includes a rotatable shaft, a compressor rotatable with the shaft to generate a flow of compressed air, a motor connected to the shaft, and a turbine connected to the shaft. The system further includes a fuel cell connected to the compressor by a first compressed air supply line that supplies a first portion of the flow of compressed air to the fuel cell. The fuel cell is connected to the turbine by a fuel cell exhaust line that supplies a flow of fuel cell exhaust to the turbine and causes the turbine to rotate. The system further includes an environmental control system connected to the compressor by a second compressed air supply line that supplies a second portion of the flow of compressed air to the environmental control system. 1. An integrated fuel cell and environmental control system comprising: a rotatable shaft;', 'a compressor rotatable with the shaft to generate a flow of compressed air;', 'a motor connected to the shaft; and', 'a turbine connected to the shaft;, 'a turbo-compressor comprisinga fuel cell connected to the compressor by a first compressed air supply line that supplies a first portion of the flow of compressed air to the fuel cell, wherein the fuel cell is connected to the turbine by a fuel cell exhaust line that supplies a flow of fuel cell exhaust to the turbine and causes the turbine to rotate; andan environmental control system connected to the compressor by a second compressed air supply line that supplies a second portion of the flow of compressed air to the environmental control system.2. The system of claim 1 , wherein the environmental control system comprises at least one of an air cycle system and a vapor cycle system.3. The system of claim 1 , wherein the motor of the turbo-compressor is connected to the fuel cell by a cable and draws a load of electrical power from the fuel ...

AIRCRAFT INCLUDING PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, and a turbine section with a first turbine and a second turbine. The first compressor is connected to the first turbine via a first shaft and the second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. An electric energy storage component is electrically connected to the electric motor, and electrically connected to at least one aircraft taxiing system. The gas turbine engine is configured such that the gas turbine engine requires supplemental power from the electric motor during at least one mode of operations. 1. A gas turbine engine comprising:a core including a compressor section having a first compressor and a second compressor, a turbine section having a first turbine and a second turbine;the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;an electric motor connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft;an electric energy storage component electrically connected to said electric motor, and electrically connected to at least one aircraft taxiing system; andwherein the gas turbine engine is configured such that the gas turbine engine requires supplemental power from the electric motor during at least one mode of operations.2. The gas turbine engine of claim 1 , wherein the aircraft taxiing system is a traction drive.3. The gas turbine engine of claim 2 , wherein the traction drive is drivably connected to at least one of an aircraft landing gear wheels and transmission.4. The gas turbine engine of claim 1 , wherein the aircraft taxiing system is a fan connected to said first shaft via at least one gearing system.5. The gas ...

HIGH EFFICIENCY AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor, and a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine. The first compressor is connected to the first turbine via a first shaft. The second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. A fan is connected to the first shaft via a gear system. The gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations. The gas turbine engine is sized to operate at peak efficiency in the maximum cruise mode of operations. 1. A gas turbine engine comprising:a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor;a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine,the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;an electric motor connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft;a fan connected to the first shaft via a gear system; andwherein the gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations, the gas turbine engine being sized to operate at peak efficiency in said maximum cruise mode of operations.2. The gas turbine engine of claim 1 , wherein operating said engine at peak efficiency comprises operating said engine at a maximum turbine inlet temperature of said ...

DESCENT OPERATION FOR AN AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, a turbine section with a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section. The first compressor is connected to the first turbine via a first shaft, the second compressor is connected to the second turbine via a second shaft, and a motor is connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. The gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation. The gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation, and a controller is configured to control the mode of operation of the gas turbine engine. 1. A gas turbine engine comprising:a core including a compressor section having a first compressor and a second compressor, a turbine section having a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section;the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;a motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft;wherein the gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation, and wherein the gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation; anda controller configured to control the mode of operation of the gas turbine engine.2. The gas turbine engine of claim 1 , wherein the at least one additional mode of operation includes a ...

POWER SYSTEM FOR AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a compressor section having a first compressor and a second compressor and a turbine section having a first turbine and a second turbine. The first compressor is connected to the first turbine via a first shaft and the second compressor is connected to the second turbine via a second shaft. A motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. A power distribution system connects the motor to a stored power system including at least one of an energy storage unit and a supplementary power unit. The power distribution system is configured to provide power from the stored power system to the motor. 1. A gas turbine engine comprising:a compressor section having a first compressor and a second compressor;a turbine section having a first turbine and a second turbine,the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;a motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft; anda power distribution system connecting the motor to a stored power system including at least one of an energy storage unit and a supplementary power unit, wherein the power distribution system is configured to provide power from the stored power system to the motor.2. The gas turbine engine of claim 1 , wherein the power distribution system is isolated within the gas turbine engine.3. The gas turbine engine of claim 1 , wherein the power distribution system is integrated with an aircraft power distribution system.4. The gas turbine engine of claim 3 , wherein the power distribution system is connected to at least one aircraft electric system claim 3 , and is configured to provide operational power to the at least one aircraft electric system.5. The gas turbine engine of claim 1 , wherein the stored power system includes an energy storage unit ...

AIRCRAFT HAVING HYBRID-ELECTRIC PROPULSION SYSTEM WITH ELECTRIC STORAGE LOCATED IN FUSELAGE

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. The aircraft includes an electrical system having an electric storage. The electric storage is positioned within the fuselage. 1. An aircraft comprising:a fuselage defining a longitudinal axis between a forward end and an aft end; andan electrical system having an electric storage and an electric-motor controller electrically connected to the electric storage, wherein the electric-motor controller is positioned on a top portion of the aircraft.2. The aircraft as recited in claim 1 , wherein the electric-motor controller is positioned in a wing space above the fuselage.3. The aircraft as recited in claim 1 , wherein the electric storage is positioned on a bottom side of a cabin opposite from the electric-motor controller claim 1 , wherein the electrical system includes at least one conductor extending from the electric storage claim 1 , up a first side of a cabin wall to the electric-motor controller.4. The aircraft as recited in claim 3 , wherein the at least one conductor follows a circumferential curvature of the fuselage.5. The aircraft as recited in claim 3 , wherein the at least one conductor is part of a DC circuit.6. The aircraft as recited in claim 3 , further comprising an airfoil extending laterally from the fuselage in a direction opposite from the first side of the cabin wall claim 3 , and a nacelle mounted to the airfoil claim 3 , wherein an electric motor is positioned in at least one of the airfoil or the nacelle.7. The aircraft as recited in claim 6 , wherein the electrical system is electrically coupled to the electric motor by way of a power bus extending from the electric-motor controller through the airfoil.8. The aircraft as recited in claim 7 , wherein the power bus include five conductors.9. The aircraft as recited in claim 7 , wherein the power bus is part of an AC circuit.10. The aircraft as recited in claim 6 , further comprising a hybrid ...

ENERGY USAGE SUB-METERING SYSTEM UTILIZING INFRARED THERMOGRAPHY

An energy usage sub-metering system of a building management system may include a focal plane array and a computer-based processor coupled to the focal plane array. The focal plane array includes a plurality of radiant energy sensors configured to monitor a predefined space and measure energy radiating from a plurality of objects in the predefined space. The computer-based processor is configured to profile temperatures of the plurality of objects throughout the predefined space in concert with executing heat dissipation algorithms to provide an estimate of energy load. 1. An energy usage sub-metering system comprising:a focal plane array including a plurality of radiant energy sensors configured to monitor a predefined space and measure energy radiating from a plurality of objects in the predefined space; anda computer-based processor coupled to the focal plane array and configured to profile temperature throughout the predefined space in concert with executing heat dissipation algorithms to provide an estimate of energy load.2. The energy usage sub-metering system set forth in claim 1 , wherein the energy load comprises a HVAC load.3. The energy usage sub-metering system set forth in claim 1 , wherein the plurality of radiant energy sensors comprise a plurality of infrared sensors.4. The energy usage sub-metering system set forth in claim 1 , wherein the computer-based processor is configured to estimate surface area of each one of the plurality of objects based at least in-part upon a number of energy sensors of the plurality of radiant energy sensors mapped upon the surface area.5. The energy usage sub-metering system set forth in claim 4 , wherein the computer-based processor is configured to apply a temperature value of supply air flowing into the predefined space.6. The energy usage sub-metering system set forth in further comprising:a wireless communication module configured to transmit and receive signals; anda substrate platform wherein the focal plane ...

ON-BOARD AIRCRAFT DRIED INERT GAS SYSTEM

An on-board aircraft dried inert gas system includes a source inert gas containing water, an air cycle or vapor cycle cooling system, and a heat exchanger condenser. The heat exchanger condenser has a heat absorption side in thermal communication with the air cycle or vapor cycle cooling system. The heat exchanger condenser has a heat rejection side that receives the inert gas containing water and outputs dried inert gas. 1. An on-board aircraft dried inert gas , comprising:a source of an inert gas comprising water;an air cycle or vapor cycle cooling system; anda heat exchanger condenser comprising a heat absorption side in thermal communication with an air cycle or vapor cycle cooling system, and a heat rejection side that receives the inert gas comprising water and outputs dried inert gas.2. The system of claim 1 , wherein the source of inert gas comprising water comprises a proton exchange membrane electrochemical cell that reacts oxygen in air with hydrogen to produce the oxygen-depleted gas.3. The system of claim 2 , further comprising a water recycle of condensate from the heat rejection side of the heat exchanger condenser to the proton exchange membrane electrochemical cell4. The system of claim 1 , wherein the source of inert gas comprising water comprises a catalytic reactor that reacts oxygen with hydrocarbon in fuel tank vapor to produce the oxygen-depleted gas.5. The system of claim 1 , wherein the heat absorption side of the heat exchanger condenser is in thermal communication with an air cycle cooling system.6. The system of claim 5 , wherein the heat absorption side of the heat exchanger condenser receives claim 5 , or is in thermal communication through a heat transfer fluid with claim 5 , the conditioned air output from the air cycle cooling system.7. The system of claim 5 , wherein the heat absorption side of the heat exchanger condenser receives claim 5 , or is in thermal communication through a heat transfer fluid with claim 5 , an intermediate ...

Fin-diffuser heat sink with high conductivity heat spreader

A method and apparatus for cooling a heat source is disclosed. The apparatus includes a fin-diffuser including a blower integrated with fins of a diffuser. A heat spreader is coupled to the fin-diffuser. The heat spreader is configured to spread heat from a location proximate the blower to location of the fins. The apparatus spreads heat from a heat source proximate a blower of the fin-diffuser to a location away from the blower to cool the heat source.

SUPPLEMENTAL POWER FOR REDUCTION OF PRIME MOVER

Embodiments are directed to selecting, by a computing device comprising a processor, the size of at least one prime mover associated with an aircraft to satisfy a baseline power requirement for operation of the aircraft during a steady state load condition, selecting at least one power source configuration to supplement power provided by the at least one prime mover during a transient load condition associated with the operation of the aircraft, and selecting, by the computing device, a parameter of the at least one power source configuration to provide a total power in an amount that is greater than a threshold during the transient condition. 1. A method comprising:selecting, by a computing device comprising a processor, the size of at least one prime mover associated with an aircraft to satisfy a baseline power requirement for operation of the aircraft during a steady state load condition;selecting at least one power source configuration to supplement power provided by the at least one prime mover during a transient load condition associated with the operation of the aircraft; andselecting, by the computing device, a parameter of the at least one power source configuration to provide a total power in an amount that is greater than a threshold during the transient condition.2. The method of claim 1 , wherein the at least one power source configuration comprises an electrical power system claim 1 , and wherein the electrical power system comprises a battery that provides power to a motor-generator during the transient load condition claim 1 , and wherein the at least one prime mover charges the battery during the steady state condition.3. The method of claim 1 , wherein the at least one power source configuration comprises a heat engine that operates on the same fuel source as the at least one prime mover.4. The method of claim 1 , wherein the at least one prime mover comprises a spark ignition (SI) engine claim 1 , and wherein the at least one power source ...

BOOSTER COMPRESSOR WITH SPEED CHANGE SYSTEM

A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed. 1. A gas turbine engine comprising;a main engine compressor section;a booster compressor for changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system, wherein the booster compressor operates at airflow conditions greater than a demand by the pneumatic system; anda variable speed transmission in driving engagement with the booster compressor, the variable speed transmission configured to drive the booster compressor at different speeds with a constant speed driving input.2. The gas turbine engine as recited in claim 1 , including an accessory gearbox driven by a shaft of the gas turbine engine claim 1 , the accessory gearbox coupled to drive the variable speed transmission.3. The gas turbine engine as recited in claim 2 , wherein the shaft of the gas turbine engine is coupled to one of a high spool or a low spool of the gas turbine engine.4. The gas turbine engine as recited in claim 3 , wherein the low spool of the gas turbine engine includes a low pressure compressor coupled to a low pressure turbine through an inner shaft.5. The gas turbine engine as recited in claim 4 , wherein the high spool of the gas turbine engine includes a high pressure compressor coupled to a high pressure turbine through an outer shaft.6. The gas turbine engine as recited in claim 1 , including an electric motor coupled to drive the variable transmission.7. The gas turbine engine as recited in ...

ENVIRONMENTAL CONTROL SYSTEM AIR CIRCUIT

An aircraft has a gas turbine engine including a compressor section that includes at least one compressor bleed. An environmental control system has an air input configured to receive pressurized cabin air. An intercooler has an input and an output. A selection valve is configured to selectively connect the bleeds to an intercooler input. At least one auxiliary compressor is connected to the intercooler output. An output of at least one auxiliary compressors is connected to an ECS air input. A controller is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor. A method for supplying engine air to an environmental control system and a system for use on a turbine engine powered aircraft are also disclosed. 1. An aircraft comprising:a gas turbine engine including a compressor section, the compressor section including at least one compressor bleed;an environment control system (ECS) having an air input configured to receive pressurized cabin air;an intercooler having an input and an output;a selection valve configured to selectively connect said bleed to said intercooler input; andat least one auxiliary compressor connected to said intercooler output, wherein an output of the at least one auxiliary compressor is connected to said ECS air input; anda controller configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least said selection valve to achieve a desired pressure and temperature for the air downstream of said auxiliary compressor.2. The system of claim 1 , wherein the at least one auxiliary compressor comprises a plurality of auxiliary compressors.3. The system of claim 1 ...

ELECTRO-PNEUMATIC ENVIRONMENTAL CONTROL SYSTEM AIR CIRCUIT

An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input. 1. An engine driven environmental control system (ECS) air circuit comprising:a gas turbine engine including a compressor section, the compressor section including a plurality of compressor bleeds;a selection valve selectively connecting each of said bleeds to an input of an intercooler; anda second valve configured to selectively connect an output of said intercooler to at least one auxiliary compressor, the output of each of the at least one auxiliary compressors being connected to an ECS air input.2. The engine driven ECS air circuit of claim 1 , wherein the at least one auxiliary compressor comprises a plurality of auxiliary compressors.3. The engine driven ECS air circuit of claim 1 , wherein at least one of said compressor bleeds is a compressor bleed positioned between a low pressure compressor and a high pressure compressor.4. The engine driven ECS air circuit of claim 1 , wherein the intercooler is an air to air heat exchanger.5. The engine driven ECS air circuit of claim 4 , wherein a heat sink of the air to air heat exchanger is fan air.6. The engine driven ECS air circuit of claim 1 , further comprising an aircraft controller controllably connected to the selection valve and to the second valve such that the aircraft controller controls a state of the selection valve and a state of the second valve.7. The engine driven ECS air circuit of claim 6 , wherein the aircraft controller includes a memory storing instructions configured to cause the controller to connect a bleed ...

COMPLEX AIR SUPPLY SYSTEM FOR GAS TURBINE ENGINE AND ASSOCIATED AIRCRAFT

A lower pressure tap is connected to a first heat exchanger to be cooled by cooling air, and then to a selection valve. The selection valve selectively delivers the lower pressure tap air to a boost compressor. The lower pressure tap air downstream of the boost compressor is connected to cool the at least one turbine. The selection valve also selectively delivers a portion of the lower pressure tap air across a first cooling turbine, and to a line associated with an air delivery system for a cabin on an associated aircraft. A portion of the air downstream of the first cooling turbine is connected to a second cooling turbine, and air downstream of the second cooling turbine is connected for use in a cold loop A method of operating an air supply system is also disclosed. 1. An aircraft air supply system comprising:a gas turbine engine including a fan and at least one compressor, the at least one compressor driven by at least one turbine section, and a combustor for combusting air from the at least one compressor and delivering it across at least one turbine;a lower pressure tap within the at least one compressor, and the lower pressure tap air being connected to a first heat exchanger to be cooled by cooling air, and then to a selection valve, the selection valve for selectively delivering the lower pressure tap air to a boost compressor, and lower pressure tap air downstream of the boost compressor being connected to cool the at least one turbine; andthe selection valve for also selectively delivering a portion of the lower pressure tap air across a first cooling turbine, and to a line associated with an air delivery system for a cabin on an associated aircraft, and a portion of the air downstream of the first cooling turbine being connected to a second cooling turbine, and air downstream of the second cooling turbine being connected for use in a cold loop such that the lower pressure tap air downstream of the second cooling turbine is at a lower temperature than air ...

FIN-DIFFUSER HEAT SINK WITH HIGH CONDUCTIVITY HEAT SPREADER

A method and apparatus for cooling a heat source is disclosed. The apparatus includes a fin-diffuser including a blower integrated with fins of a diffuser. A heat spreader is coupled to the fin-diffuser. The heat spreader is configured to spread heat from a location proximate the blower to location of the fins. The apparatus spreads heat from a heat source proximate a blower of the fin-diffuser to a location away from the blower to cool the heat source. 1. A method of cooling a heat source , comprising:coupling an integrated fin-diffuser to a heat spreader to form a cooling assembly;coupling the cooling assembly to the heat source; andspreading heat from the heat source generated proximate a blower of the fin-diffuser to a location away from the blower to cool the heat source.2. The method of claim 1 , wherein the heat spreader further comprises a vapor chamber for spreading the heat using a motion of working fluid in the vapor chamber.3. The method of claim 2 , wherein the working fluid transfers heat via an evaporation-condensation cycle.4. The method of claim 1 , wherein the heat spreader further comprises one o£ a capillary wick heat pipe; and an oscillating heat pipe.5. The method of claim 4 , wherein the oscillating heat pipe is one of: attached to a surface of the heat spreader claim 4 , and embedded in the heat spreader.6. The method of claim 4 , wherein the oscillating heat pipe transfers heat away from the heat source along a radial direction.7. The method of claim 4 , wherein heat source further comprises a plurality of heat sources claim 4 , further comprising coupling providing a plurality of oscillating heat pipes claim 4 , with one of the plurality of oscillating heat pipes centered at one of the plurality heat sources.8. An apparatus for cooling a heat source claim 4 , comprising:a fin-diffuser comprising a blower integrated with fins of a diffuser; anda heat spreader coupled to the fin-diffuser, wherein the heat spreader is configured to spread heat ...

ENHANCED THERMAL MANAGEMENT FOR DIRECTED ENERGY WEAPON

Described herein is a thermal management system and methodology for a directed energy weapon on an aircraft. The thermal management system includes an evaporator in thermal communication with the directed energy weapon and operatively configured to cool the directed energy weapon by evaporating a refrigerant therein. The thermal management system also includes a refrigerant storage tank in fluid communication with the evaporator and a pump in fluid communication with the refrigerant storage tank and the evaporator configured to pump substantially liquid refrigerant to the evaporator. 1. A thermal management system for a directed energy weapon on an aircraft the thermal management system comprising:an evaporator in thermal communication with the directed energy weapon and operatively configured to cool the directed energy weapon by evaporating a refrigerant therein;a refrigerant storage tank in fluid communication with the evaporator, the refrigerant storage tank configured to separate liquid refrigerant and vapor refrigerant; anda pump in fluid communication with the refrigerant storage tank and the evaporator and configured to pump substantially liquid refrigerant to the evaporator.2. The thermal management system of claim 1 , further including a check valve in fluid communication with the pump and the evaporator operable to ensure that the substantially liquid refrigerant flows to the evaporator.3. The thermal management system according to claim 1 , further including a bypass valve operably connected in parallel to the evaporator.4. The thermal management system of claim 3 , wherein the evaporator is a heat exchanger.5. The thermal management system of claim 1 , wherein the refrigerant storage tank includes a separator section.6. The thermal management system of claim 5 , wherein the separator section includes a coolant coil to condense vapor refrigerant.7. The thermal management system of claim 5 , wherein the separator section includes a centrifugal separator.8 ...

Electrocaloric heat transfer method

A heat transfer system cycles between a first mode where a heat transfer fluid is directed to a first electrocaloric module and from the first electrocaloric module to a heat exchanger to a second electrocaloric module while one of the first and second electrocaloric modules is energized, and a second mode where the heat transfer fluid is directed to the second electrocaloric module and from the second electrocaloric module to the heat exchanger to the first electrocaloric module, while the other of the first and second electrocaloric modules is energized. The modes are repeatedly cycled in alternating order directing the heat transfer fluid to cause a temperature gradient in each of the first and second electrocaloric modules, and heat is rejected to the fluid from the heat exchanger or is absorbed by the heat exchanger from the fluid.

Electro-pneumatic environmental control system air circuit

An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

AIRCRAFT HAVING HYBRID-ELECTRIC PROPULSION SYSTEM WITH ELECTRIC STORAGE LOCATED IN FUSELAGE

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. The aircraft includes an electrical system having an electric storage. The electric storage is positioned within the fuselage. 1. An aircraft comprising:a fuselage defining a longitudinal axis between a forward end and an aft end;an electrical system having an electric storage, wherein the electric storage is positioned within the fuselage.2. The aircraft as recited in claim 1 , further comprising a hybrid electric propulsion system claim 1 , wherein the electrical system is part of the hybrid electric propulsion system claim 1 , wherein the hybrid electric propulsion system includes a heat engine.3. The aircraft as recited in claim 2 , wherein the hybrid electric propulsion system includes an electric motor claim 2 , wherein the electrical system is electrically coupled to the electric motor by way of a 1000-volt power bus.4. The aircraft as recited in claim 2 , wherein the hybrid electric propulsion system includes an electric motor claim 2 , wherein the electrical system is electrically coupled to the electric motor by way of a high voltage power bus.5. The aircraft as recited in claim 2 , wherein the hybrid electric propulsion system includes an electric motor claim 2 , wherein the electrical system and the electric storage are operatively connected to the electric motor for receiving power therefrom or for supplying power thereto.6. The aircraft as recited in claim 1 , wherein the fuselage defines an interior cabin space claim 1 , wherein the interior cabin space includes a cabin floor claim 1 , wherein the electrical system includes a plurality of batteries claim 1 , wherein the plurality of batteries are mounted to the cabin floor.7. The aircraft as recited in claim 6 , wherein the cabin floor defines a lower surface claim 6 , wherein the plurality of batteries are mounted to the lower surface of the cabin floor.8. The aircraft as recited in claim 7 , wherein ...

AIRCRAFT HAVING HYBRID-ELECTRIC PROPULSION SYSTEM WITH ELECTRIC STORAGE LOCATED IN WINGS

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. At least one airfoil is laterally extending from the fuselage defining an airfoil axis. An electrical system has an electric storage. The electric storage is positioned within the airfoil. 1. An aircraft comprising:a fuselage defining a longitudinal axis between a forward end and an aft end;at least one airfoil laterally extending from the fuselage defining an airfoil axis;an electrical system having an electric storage, wherein the electric storage is positioned within the airfoil.2. The aircraft as recited in claim 1 , further comprising a hybrid-electric propulsion system claim 1 , wherein the electrical system is part of the hybrid-electric propulsion system claim 1 , wherein the hybrid-electric propulsion system includes a heat engine.3. The aircraft as recited in claim 2 , wherein the hybrid-electric propulsion system includes an electric-motor claim 2 , wherein the electrical system is electrically coupled to the electric-motor by way of a 1000-volt power bus.4. The aircraft as recited in claim 2 , wherein the hybrid-electric propulsion system includes an electric-motor claim 2 , wherein the electrical system is electrically coupled to the electric-motor by way of a high voltage power bus.5. The aircraft as recited in claim 2 , wherein the hybrid-electric propulsion system includes an electric-motor claim 2 , wherein the electric system and electric storage are operatively connected to the electric-motor for receiving power therefrom or for supplying power thereto.6. The aircraft as recited in claim 1 , wherein the aircraft includes a nacelle mounted to the airfoil.7. The aircraft as recited in claim 6 , wherein the electric storage is positioned at least one of inboard of or outboard of the nacelle.8. The aircraft as recited in claim 6 , further comprising a hybrid-electric propulsion system claim 6 , wherein the electrical system is part of the hybrid-electric ...

CRYOGENIC COOLING SYSTEM FOR AN AIRCRAFT

A cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for collecting liquid air. The first air cycle machine is operable to output a cooling air stream based on a first air source. The second air cycle machine is operable to output a chilled air stream at a cryogenic temperature based on a second air source cooled by the cooling air stream of the first air cycle machine. An output of the second air cycle machine is provided to the means for collecting liquid air. 1. A cryogenic cooling system for an aircraft , the cryogenic cooling system comprising:a first air cycle machine operable to output a cooling air stream based on a first air source;a second air cycle machine operable to output a chilled air stream at a cryogenic temperature based on a second air source cooled by the cooling air stream of the first air cycle machine; anda means for collecting liquid air from an output of the second air cycle machine.2. The cryogenic cooling system of claim 1 , wherein the first air cycle machine comprises a first compressor section and a first turbine section claim 1 , and the second air cycle machine comprises a second compressor section and a second turbine section.3. The cryogenic cooling system of claim 2 , wherein the first compressor section comprises a first compressor wheel operably coupled to a first turbine wheel of the first turbine section and a first fan claim 2 , and the second compressor section comprises a second compressor wheel operably coupled to a second turbine wheel of the second turbine section and a second fan.4. The cryogenic cooling system of claim 3 , further comprising a first heat exchanger system operable to pre-cool a first air flow from the first air source prior to entry into the first compressor wheel and cool the first air flow after exiting the first compressor wheel claim 3 , and a second heat exchanger system operable to pre-cool a second air flow from the second air ...

THERMAL REGULATION OF BATTERIES

A battery thermal management system for an air vehicle includes a first heat exchange circuit, a battery in thermal communication with the first heat exchange circuit, and a heat exchanger positioned on the first heat exchange circuit. The heat exchanger is operatively connected to a second heat exchange circuit. The system includes a controller operatively connected to the second heat exchange circuit. The controller is configured to variably select whether heat will be rejected to the second heat exchange circuit. A method for controlling a thermal management system for an air vehicle includes determining an expected fluid temperature of fluid in a fluid heat exchange circuit. The method includes commanding a flow restrictor at least partially closed or commanding the flow restrictor at least partially open. 1. A battery thermal management system for an air vehicle comprising:a first heat exchange circuit;a battery in thermal communication with the first heat exchange circuit;a heat exchanger positioned on the first heat exchange circuit, wherein the heat exchanger is operatively connected to a second heat exchange circuit; anda controller operatively connected to the second heat exchange circuit, wherein the controller is configured and adapted to variably select whether heat will be rejected to the second heat exchange circuit.2. The system as recited in claim 1 , further comprising at least one of (i) a first bypass circuit extending from the first heat exchange circuit upstream of the heat exchanger and reconnecting to the first heat exchange circuit downstream from the heat exchanger claim 1 , and/or (ii) a second bypass circuit branching from the second heat exchange circuit upstream from the heat exchanger and reconnecting to an outlet side of the second heat exchange circuit downstream from the heat exchanger.3. The system as recited in claim 1 , wherein the second heat exchange circuit is in fluid communication with a ram air source.4. The system as ...

THERMAL REGULATION OF BATTERIES

A battery thermal management system for an air vehicle includes a liquid heat exchange circuit. The system includes at least one battery in thermal communication with the liquid heat exchange circuit. The system includes a controller operatively connected to the liquid heat exchange circuit. The controller is configured and adapted to variably select whether heat will be rejected to the liquid heat exchange circuit. The system includes at least one heat exchanger positioned on the liquid heat exchange circuit. The at least one heat exchanger is a liquid-air heat exchanger or a liquid-liquid heat exchanger. A method for controlling a thermal management system for an air vehicle includes determining an expected temperature of at least one battery, and cooling the at least one battery with a cooling device if the expected temperature exceeds a pre-determined expected temperature threshold. 1. A battery thermal management system for an air vehicle comprising:a liquid heat exchange circuit;at least one battery in thermal communication with the liquid heat exchange circuit;a controller operatively connected to the liquid heat exchange circuit, wherein the controller is configured and adapted to variably select whether heat will be rejected to the liquid heat exchange circuit; andat least one heat exchanger positioned on the liquid heat exchange circuit, wherein the at least one heat exchanger is a liquid-air heat exchanger or a liquid-liquid heat exchanger.2. The system as recited in claim 1 , wherein the liquid-liquid heat exchanger is configured and adapted to be operatively connected to a movable second liquid heat exchange circuit.3. The system as recited in claim 1 , wherein the liquid-air heat exchanger is positioned on an air heat exchange circuit in fluid communication with an air scoop.4. The system as recited in claim 3 , further comprising a temperature sensor upstream from the liquid-air heat exchanger positioned to measure a total air temperature of air ...

Electrocaloric heat transfer system

A heat transfer system cycles between a first mode where a heat transfer fluid is directed to a first electrocaloric module and from the first electrocaloric module to a heat exchanger to a second electrocaloric module while one of the first and second electrocaloric modules is energized, and a second mode where the heat transfer fluid is directed to the second electrocaloric module and from the second electrocaloric module to the heat exchanger to the first electrocaloric module, while the other of the first and second electrocaloric modules is energized. The modes are repeatedly cycled in alternating order directing the heat transfer fluid to cause a temperature gradient in each of the first and second electrocaloric modules, and heat is rejected to the fluid from the heat exchanger or is absorbed by the heat exchanger from the fluid.

ENGINE BLLED AIR WITH COMPRESSOR SURGE MANAGEMENT

A gas turbine engine includes a main compressor section. A booster compressor includes an inlet and an outlet. The inlet receives airflow from the main compressor section and the outlet provides airflow to a pneumatic system. A recirculation passage is between the inlet and the outlet. A flow splitter valve controls airflow between the outlet and the inlet through the recirculation passage for controlling airflow to the pneumatic system based on airflow output from the booster compressor. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed. 1. A gas turbine engine comprising;a main compressor section;a booster compressor including an inlet and an outlet, the inlet receiving airflow from the main compressor section and the outlet providing airflow to a pneumatic system;a recirculation passage between the inlet and the outlet; anda flow splitter valve controlling airflow between the outlet and the inlet through the recirculation passage for controlling airflow to the pneumatic system based on airflow output from the booster compressor.2. The gas turbine engine as recited in claim 1 , including a heat exchanger for cooling airflow through the recirculation passage prior to the inlet of the booster compressor.3. The gas turbine engine as recited in claim 2 , wherein the main compressor section includes a low pressure compressor section suppling airflow through a first passage to the booster compressor inlet and a high pressure compressor section supplying airflow through a second passage directly to the pneumatic system separate from the booster compressor.4. The gas turbine engine as recited in claim 3 , wherein the second passage may extend through the heat exchanger separate from airflow through the recirculation passage for cooling airflow supplied directly to the pneumatic system from the high pressure compressor section.5. The gas turbine engine as recited in claim 3 , including a first control valve ...

BOOSTER COMPRESSOR WITH SPEED CHANGE SYSTEM

A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed. 1. A gas turbine engine comprising;a main engine compressor section;a booster compressor for changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system, wherein the booster compressor operates at airflow conditions greater than a demand by the pneumatic system; anda speed change system for driving the booster compressor at speeds corresponding to a demand of the pneumatic system.2. The gas turbine engine as recited in claim 1 , wherein the speed change system comprises a variable speed transmission coupled to the booster compressor.3. The gas turbine engine as recited in claim 2 , including an accessory gearbox driven by a shaft of the gas turbine engine claim 2 , the accessory gearbox coupled to drive the variable speed transmission.4. The gas turbine engine as recited in claim 1 , including a turbine driven by airflow from the main engine compressor claim 1 , wherein the speed change system comprises a gearbox coupled between the turbine and the booster compressor such that the turbine and the booster compressor rotate at different speeds.5. The gas turbine engine as recited in claim 4 , including a control valve controlling airflow through the turbine for controlling a speed of the turbine.6. The gas turbine engine as recited in claim 1 , including an exhaust valve for exhausting airflow in excess of an airflow communicated to the pneumatic system.7. The gas turbine ...

ENGINE BLEED AIR SYSTEM WITH WASTE GATE VALVE FOR COMPRESSOR SURGE MANAGEMENT

A gas turbine engine includes a main engine compressor section. A booster compressor changes a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor is configured to operate at airflow conditions greater than a demand of the pneumatic system. An exhaust valve controls airflow between an exhaust outlet and an outlet passage to the pneumatic system. The exhaust valve is operable to exhaust airflow from the booster compressor in excess of the demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed. 1. A gas turbine engine comprising;a main engine compressor section;a booster compressor for changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system, wherein the booster compressor is configured to operate at airflow conditions greater than a demand of the pneumatic system; andan exhaust valve controlling airflow between an exhaust outlet and an outlet passage to the pneumatic system, the exhaust valve operable to exhaust airflow from the booster compressor in excess of the demand of the pneumatic system.2. The gas turbine engine as recited in claim 1 , wherein the booster compressor is driven by one of an output shaft of a gearbox and an electric motor.3. The gas turbine engine as recited in claim 1 , wherein the main engine compressor includes a low pressure air source that communicates airflow to the booster compressor at a pressure less than the demand of the pneumatic system.4. The gas turbine engine as recited in claim 1 , including a turbine coupled to drive the booster compressor claim 1 , the turbine driven by airflow from the main engine compressor section.5. The gas turbine engine as recited in claim 4 , wherein the turbine is driven by a high pressure air source separate from a source of air pressure in communication with the ...

TURBINE ENGINES HAVING HYDROGEN FUEL SYSTEMS

An aircraft propulsion systems and aircraft having the same are described. The aircraft propulsion systems have one or more aircraft systems including at least one hydrogen tank and a first heat exchanger and one or more engine systems including at least a main engine core, a second heat exchanger, and a third heat exchanger. The main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is configured to be supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the first heat exchanger of the aircraft systems, the second heat exchanger of the engine systems, and the third heat exchanger of the engine systems, and then supplied into the burner for combustion.

HYDROGEN POWERED ENGINE WITH EXHAUST HEAT EXCHANGER

A turbine engine system includes at least one hydrogen fuel tank, a core flow path heat exchanger in a core flow path; and engine systems located in the core flow path. The engine system including at least a compressor section, a combustor section having a burner, and a turbine section. The core flow path heat exchanger is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank through a hydrogen fuel supply line, passing through the core flow path heat exchanger and then supplied into the burner for combustion.

ASSEMBLY OF A SUBSTRATE FOR ELECTRONIC COMPONENTS AND A COOLING STRUCTURE

L'INVENTION CONCERNE UN ENSEMBLE CONSTITUE D'UN SUBSTRAT PORTANT UN COMPOSANT ELECTRONIQUE ET D'UNE STRUCTURE DE DISSIPATION DE LA CHALEUR. LA STRUCTURE 10 DE DISSIPATION DE LA CHALEUR DEFINIT UNE CHAMBRE INTERIEURE FERMEE 30 DELIMITEE PAR LES SURFACES INTERIEURES DE PAROIS DE CETTE STRUCTURE. DES RAINURES 40, 42 SONT MENAGEES DANS LES SURFACES INTERIEURES AFIN QU'UN FLUIDE DE REFROIDISSEMENT PUISSE CIRCULER PAR CAPILLARITE. LE SUBSTRAT 14 PORTANT LE COMPOSANT ELECTRONIQUE 16 EST MONTE A PLAT SUR LA PAROI SUPERIEURE DE LA STRUCTURE 10. UN RADIATEUR THERMIQUE EST MONTE EN RELATION D'ECHANGE DE CHALEUR AVEC UNE EXTREMITE DE LA CHAMBRE 30 AFIN QUE CETTE DERNIERE SOIT DIVISEE EN UNE PARTIE FORMANT EVAPORATEUR, AU-DESSUS DE LAQUELLE LE SUBSTRAT 14 EST MONTE, ET UNE PARTIE FORMANT CONDENSEUR. DOMAINE D'APPLICATION : REFROIDISSEMENT DES BLOCS ELECTRONIQUES, ETC. THE INVENTION CONCERNS AN ASSEMBLY CONSISTING OF A SUBSTRATE CARRYING AN ELECTRONIC COMPONENT AND A HEAT DISSIPATING STRUCTURE. THE HEAT DISSIPATING STRUCTURE 10 DEFINES A CLOSED INTERIOR CHAMBER 30 DELIMITED BY THE INTERIOR WALL SURFACES OF THIS STRUCTURE. GROOVES 40, 42 ARE PROVIDED IN THE INTERIOR SURFACES SO THAT A COOLING FLUID CAN CIRCULATE THROUGH CAPILLARITY. THE SUBSTRATE 14 CARRYING THE ELECTRONIC COMPONENT 16 IS MOUNTED FLAT ON THE UPPER WALL OF THE STRUCTURE 10. A THERMAL RADIATOR IS MOUNTED IN A HEAT EXCHANGE RELATIONSHIP WITH AN END OF CHAMBER 30 SO THAT THIS LATTER IS DIVIDED INTO A FORMING PART EVAPORATOR, OVER WHICH THE SUBSTRATE 14 IS MOUNTED, AND A CONDENSER PART. FIELD OF APPLICATION: COOLING OF ELECTRONIC BLOCKS, ETC.

ASSEMBLY OF A SUBSTRATE FOR ELECTRONIC COMPONENTS AND A COOLING STRUCTURE

On-board aircraft dried inert gas system

An on-board aircraft dried inert gas system includes a source inert gas (38) containing water, an air cycle or vapor cycle cooling system (140), and a heat exchanger condenser (118). The heat exchanger condenser has a heat absorption side in thermal communication with the air cycle or vapor cycle cooling system. The heat exchanger condenser has a heat rejection side that receives the inert gas containing water and outputs dried inert gas.

Complex air supply system for gas turbine engine and associated aircraft

A lower pressure tap (46) is connected to a first heat exchanger (48) to be cooled by cooling air, and then to a selection valve (52). The selection valve (52) selectively delivers the lower pressure tap air to a boost compressor (44). The lower pressure tap air downstream of the boost compressor (44) is connected to cool the at least one turbine (32). The selection valve (52) also selectively delivers a portion of the lower pressure tap air across a first cooling turbine (78), and to a line (79) associated with an air delivery system for a cabin (80) on an associated aircraft. A portion of the air downstream of the first cooling turbine (78) is connected to a second cooling turbine (84), and air downstream of the second cooling turbine (84) is connected for use in a cold loop (88).

Cooling technique for compact electronics inverter

A substrate with a heat producing electronic component is mounted directly on a heat dissipating structure. The heat dissipating structure has a closed chamber with adjacent condenser and evaporator sections and contains a supply of liquid cooling fluid with a predetermined vaporization temperature. The evaporator section is in heat transfer relation with the substrate and the electronic component thereon. A plurality of channels are strategically arranged in the surfaces bounding the chamber to move the cooling fluid by capillary action in a predetermined path between the condenser and evaporator sections in heat exchange relationship over the chamber walls to maintain the substrate at an acceptable temperature. A heat sink, in heat exchange relationship with the condenser section of the heat dissipating structure, maintains a temperature differential between the evaporator and condenser sections and assures that the condenser section is cooled sufficiently to condense the cooling fluid.

Vapor compression cycle environmental control system

The present invention relates to a vapor compression cycle environmental control system for use on aircraft. The system includes an environmental control subsystem which supplies pressurized ram air at a desired temperature to the aircraft's flight deck and/or cabin. The environmental control subsystem uses a vapor compression cycle subsystem to provide the ram air at the desired temperature. The system further includes an air turbine driven by engine bleed air to provide power to an aircraft mounted accessory drive and to provide heated air to an anti-ice system. The aircraft mounted accessory drive provides power to an air compressor which forms part of the environmental control subsystem and to a working fluid compressor which forms part of the vapor compression cycle subsystem.

Adaptive heat sink for aircraft environmental control system

An avionics cooling system includes a first heat exchange system, a second heat exchange system and a vapor cycle system. The second heat exchange system has a heat sink capacity that is generally out of phase with a heat sink capacity of the first heat exchange system. The vapor cycle includes a fluid loop in communication with the first and second heat exchange systems and the fluid loop transfers heat to the first and second heat exchange systems. A method for cooling aircraft components includes selectively directing a fluid having an elevated temperature in a fluid loop of a vapor cycle system to a first heat exchanger to transfer heat from the fluid to a fuel based on a heat sink capacity of the fuel, selectively directing the fluid to a second heat exchanger to transfer heat from the fluid to air based on a heat sink capacity of the air, and cooling aircraft components with the fluid.

Aircraft having hybrid-electric propulsion system with electric storage located in fuselage

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. The aircraft includes an electrical system having an electric storage. The electric storage is positioned within the fuselage.

Supplemental power for reduction of prime mover

Aircraft power and thermal management system with electric co-generation

A power and thermal management system (10) includes an integrated power package (14) which receives bleed air (52), communicates conditioned air to an environmental control air distribution system (12) and selectively communicates electrical power with an electrical distribution system (18).

On-board aircraft dried inert gas system

An on-board aircraft dried inert gas system includes a source inert gas containing water, an air cycle or vapor cycle cooling system, and a heat exchanger condenser. The heat exchanger condenser has a heat absorption side in thermal communication with the air cycle or vapor cycle cooling system. The heat exchanger condenser has a heat rejection side that receives the inert gas containing water and outputs dried inert gas.

Combination engines for aircraft

An engine combination (10) for generating forces with a gas turbine engine (14) generating force, and an internal combustion engine (34) provided in the combination (10) as an intermittent combustion engine (34) generating force having an air intake (32), there being an air transfer duct (30) connected from a compressor (21,23) in the gas turbine engine (14) to the air intake (32) to transfer compressed air thereto.

Integrated system for providing aircraft environmental control

The present invention relates to an integrated enviromental control system for an aircraft. The system uses bleed air from a propulsion engine to drive an air turbine which provides power to at least one aircraft component such as an aircraft mounted accessory drive. The system uses the bleed air exiting the air turbine as an air source for the cabin and/or flight deck enviromental control system.

Hydrogen powered geared turbofan engine with reduced size core engine

A turbine engine system includes aircraft systems including at least one hydrogen fuel tank, engine systems comprising a compressor section, a combustor section having a burner, and a turbine section, and a hydrogen fuel flow supply line configured to supply hydrogen fuel from the at least one hydrogen fuel tank into the burner for combustion. The turbine engine system has a bypass ratio between 5 to 20.

Environmental control system air circuit

An aircraft has a gas turbine engine (20,120) including a compressor section (122) that includes at least one compressor bleed (102,104,106,108). An environmental control system (ECS) has an air input configured to receive pressurized cabin air. An intercooler (130) has an input and an output. A selection valve (140) is configured to selectively connect the bleeds to an intercooler input. At least one auxiliary compressor (160,162) is connected to the intercooler output. An output of at least one auxiliary compressors is connected to an ECS air input. A controller (101) is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor.

Integrated fuel cell aircraft pressurization and cooling system

Improved aircraft architecture with a reduced bleed aircraft secondary power system

The present invention relates to an improved architecture for an aircraft. The aircraft has a first engine (10), a first gearbox (12) associated with the first engine, a first starter/generator (18) associated with the gearbox, and a first motor drive (56) connected to the first starter/generator for providing the starter/generator with electric power to start the first engine and to receive electric power from the starter/generator after the engine has been started to operate electrically driven systems onboard the aircraft. The aircraft preferably further has at least one other engine (10') which has a gearbox (12) and a starter/generator (18) associated with it and at least a second motor drive (56) connected to the starter/generator. The electrically driven systems operated by the motor drive(s) include an environmental control system (38), a wing anti-icing system (58), an aircraft control system, and the aircraft fuel system. The aircraft also includes an auxiliary power unit (20) for supplying electrical power to at least the first motor drive for initiating operation of the first engine.

Aircraft and method for supplying engine air to an environmental control system

Aircraft power and thermal management system with electric co-generation

A power and thermal management system includes an integrated power package which receives bleed air, communicates conditioned air to an environmental control air distribution system and selectively communicates electrical power with an electrical distribution system.

Heat exchanger assembly for an aircraft control

A heat exchanger assembly for an aircraft control has an aircraft control for controlling an operation of an aircraft. The aircraft control is in thermal communication with a first fluid. A first thermoelectric device is configured to transfer heat between the first fluid and the second fluid against a temperature gradient of the first fluid and the second fluid. A temperature sensor is provided for sensing a temperature of the first fluid. A temperature control is also configured to control the first thermoelectric device based on an input from the temperature sensor.

Aircraft having hybrid-electric propulsion system with electric storage located in wings

An aircraft includes a fuselage defining a longitudinal axis between a forward end and a aft end. At least one airfoil is laterally extending from the fuselage defining an airfoil axis. An electrical system has an electric storage. The electric storage is positioned within the airfoil.

Descent operation for an aircraft parallel hybrid gas turbine engine propulsion system

A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, a turbine section with a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section. The first compressor is connected to the first turbine via a first shaft, the second compressor is connected to the second turbine via a second shaft, and a motor is connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. The gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation. The gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation, and a controller is configured to control the mode of operation of the gas turbine engine.

Hydrogen steam injected and inter-cooled turbine engine

A propulsion system (20) includes a core engine generating a high energy gas flow (55), a condenser (80) configured to extract water (35) from the high energy gas flow and an evaporator (72) transforming the extracted water into a steam flow (106). The steam flow is injected into a core flow path (C) of the core engine to increase mass flow.

Turbo expanders for turbine engines having hydrogen fuel systems

An aircraft propulsion system includes aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander (800). The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander (800), prior to being injected into the burner for combustion. The turbo expander (800) includes a rotor (804) separated into a first expander portion (812) and a second expander portion (814) arranged about an output shaft (806) and the output shaft (806) is operably connected to a generator configured to generate electrical power.

Turbine engines having hydrogen fuel systems

Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Hybrid electric power for turbine engines having hydrogen fuel systems

An aircraft propulsion systems (400) has an aircraft systems (404) including at least one hydrogen tank (432) and an aircraft-systems heat exchanger (458) and engine systems (402) includes at least a main engine core, a high pressure pump (416), a hydrogen-air heat exchanger (424), and a turbo expander (420). The main engine core includes a compressor section, a combustor section having a burner (410), and a turbine section arranged along an engine shaft. Hydrogen is supplied from the at least one hydrogen tank (432) through a hydrogen flow path (444), passing through the aircraft-systems heat exchanger (458), the high pressure pump (416), the hydrogen-air heat exchanger (424), and the turbo expander (420), prior to being injected into the burner (410) for combustion. The turbo expander (420) is configured to impart power to the engine shaft.

Hybrid electric power for turbine engines having hydrogen fuel systems

Aircraft propulsion systems have aircraft systems including at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems includes at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section arranged along an engine shaft. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander, prior to being injected into the burner for combustion. The turbo expander is configured to impart power to the engine shaft.

Turbo expanders for turbine engines having hydrogen fuel systems

Aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander, prior to being injected into the burner for combustion. The turbo expander includes a rotor separated into a first expander portion and a second expander portion arranged about an output shaft and the output shaft is operably connected to a generator configured to generate electrical power.

Hydrogen steam injected and inter-cooled turbine engine

A propulsion system includes a core engine generating a high energy gas flow, a condenser configured to extract water from the high energy gas flow and an evaporator transforming the extracted water into a steam flow. The steam flow is injected into a core flow path of the core engine increase mass flow.

Aircraft having hybrid-electric propulsion system with electric storage located in fuselage

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. The aircraft includes an electrical system having an electric storage. The electric storage is positioned within the fuselage.

Aircraft having hybrid-electric propulsion system with electric storage located in wings

An aircraft includes a fuselage defining a longitudinal axis between a forward end and a aft end. At least one airfoil is laterally extending from the fuselage defining an airfoil axis. An electrical system has an electric storage. The electric storage is positioned within the airfoil.

Hydrogen powered geared turbofan engine with reduced size core engine

A turbine engine system includes an aircraft systems including at least one hydrogen fuel tank, engine systems comprising a compressor section, a combustor section having a burner, and a turbine section, and a hydrogen fuel flow supply line configured to supply hydrogen fuel from the at least one hydrogen fuel tank into the burner for combustion. The turbine engine system has a bypass ratio between 5 to 20.

System for superconducting electronics in aerospace applications

A powertrain system of an aircraft includes one or more electrical components to provide electrical power to one or more electrical loads of the aircraft. The system further includes a rechargeable cryogenic heat sink containing a volume of cryogenic cooling material. The cryogenic heat sink is configured to cool the one or more electrical components. A method of operating a powertrain system of an aircraft includes generating thermal energy at one or more electrical components of the powertrain system, fluidly connecting a cryogenic heat sink to the one or more electrical components, and cooling the one or more electrical components via a volume of cryogenic cooling material of the cryogenic heat sink.

Turbine engines having hydrogen fuel systems

Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Turbine engines having hydrogen fuel systems

Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Turbo expanders for turbine engines having hydrogen fuel systems

Aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander, prior to being injected into the burner for combustion. The turbo expander includes a rotor separated into a first expander portion and a second expander portion arranged about an output shaft and the output shaft is operably connected to a generator configured to generate electrical power.

Aircraft having hybrid-electric propulsion system with electric storage located in fuselage

An aircraft comprising a fuselage (20) and an electrical system (101) having an electric storage (103) and an electric-motor controller (121) electrically connected to the electric storage (103), wherein the electric-motor controller (121) is positioned in a wing above the fuselage, wherein the electric storage (103) is positioned on a bottom side of a cabin opposite from the electric-motor controller (121), wherein the electrical system includes at least one conductor extending from the electric storage (103), up a first side of a cabin wall to the electric-motor controller (121), the aircraft further comprising a hybrid electric propulsion system, wherein the electrical system is part of the hybrid electric propulsion system, wherein the hybrid electric propulsion system includes a heat engine (104).

System for superconducting electronics in aerospace applications

A powertrain system (10) of an aircraft includes one or more electrical components (16, 24, 28, 32) to provide electrical power to one or more electrical loads (20) of the aircraft. The system (10) further includes a rechargeable cryogenic heat sink (42) containing a volume of cryogenic cooling material. The cryogenic heat sink (42) is configured to cool the one or more electrical components (16...32). A method of operating a powertrain system (10) of an aircraft includes generating thermal energy at one or more electrical components (16... 32) of the powertrain system (10), fluidly connecting a cryogenic heat sink (42) to the one or more electrical components (16...32), and cooling the one or more electrical components (16...32) via a volume of cryogenic cooling material of the cryogenic heat sink (42).

Hydrogen powered engine with exhaust heat exchanger

A turbine engine system (200) includes at least one hydrogen fuel tank (224), a core flow path heat exchanger (300) in a core flow path; and engine systems located in the core flow path. The engine system (200) including at least a compressor section, a combustor (210) section having a burner, and a turbine section. The core flow path heat exchanger (300) is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank (224) through a hydrogen fuel supply line (226), passing through the core flow path heat exchanger (300) and then supplied into the burner for combustion.

Turbine engines having hydrogen fuel systems

Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Turbine engines having hydrogen fuel systems

Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Aircraft having hybrid-electric propulsion system with electric storage located in wings

An aircraft includes a fuselage defining a longitudinal axis between a forward end and a aft end. At least one airfoil is laterally extending from the fuselage defining an airfoil axis. An electrical system has an electric storage. The electric storage is positioned within the airfoil.

Hydrogen powered engine with exhaust heat exchanger

A turbine engine system includes at least one hydrogen fuel tank, a core flow path heat exchanger in a core flow path; and engine systems located in the core flow path. The engine system including at least a compressor section, a combustor section having a burner, and a turbine section. The core flow path heat exchanger is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank through a hydrogen fuel supply line, passing through the core flow path heat exchanger and then supplied into the burner for combustion.

Thermal regulation of batteries

A battery thermal management system for an air vehicle includes a first heat exchange circuit, a battery in thermal communication with the first heat exchange circuit, and a heat exchanger positioned on the first heat exchange circuit. The heat exchanger is operatively connected to a second heat exchange circuit. The system includes a controller operatively connected to the second heat exchange circuit. The controller is configured to variably select whether heat will be rejected to the second heat exchange circuit. A method for controlling a thermal management system for an air vehicle includes determining an expected fluid temperature of fluid in a fluid heat exchange circuit. The method includes commanding a flow restrictor at least partially closed or commanding the flow restrictor at least partially open.

Hydrogen steam injected and inter-cooled turbine engine

A propulsion system includes a core engine generating a high energy gas flow, a condenser configured to extract water from the high energy gas flow and an evaporator transforming the extracted water into a steam flow. The steam flow is injected into a core flow path of the core engine increase mass flow.

Cryogenic cooling system for an aircraft

A gas turbine engine includes a compressor section and a turbine section operably coupled to the compressor section. The gas turbine engine further includes a means for selectively releasing a cooling fluid flow produced at a cryogenic temperature and a plumbing system in fluid communication with the means for selectively releasing the cooling fluid flow. The plumbing system is configured to route the cooling fluid flow to one or more of the compressor section and the turbine section.

Engine bleed air and compressor surge management and corresponding method

Hybrid electric power for turbine engines having hydrogen fuel systems

Aircraft propulsion systems have aircraft systems including at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems includes at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section arranged along an engine shaft. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander, prior to being injected into the burner for combustion. The turbo expander is configured to impart power to the engine shaft.