465results about "Conditioning fuel arrangements" patented technology

An aircraft thermal management system comprising a high temperature loop and a low temperature loop. The high temperature loop is configured to manage a temperature of a fuel cell stack in a nacelle using a nacelle heat exchanger. The low temperature loop is configured to cool a number of heat loads in a nacelle using a heat capacity of liquid hydrogen.

An aircraft comprises a fuselage; wings; engines connected to the wings; a set of liquid hydrogen tanks; a vent located on the tail on the fuselage; a conduit system connecting the liquid hydrogen tanks to the vent; and a controller. The controller is configured to control the conduit system to remove gaseous hydrogen in the set of liquid hydrogen tanks to travel through the conduit system and exit at the vent in response to pressure in the set of liquid hydrogen tanks being greater than a specified tolerance.

An aircraft conditioning system including: a fuel inlet (2) to a fuel conduit (3), an engine (4), a first heat exchanger (6) connected to a power unit outlet (5) of the engine, a first portion (3.1) of the fuel conduit (3) connecting the fuel inlet (2) to a fuel inlet (6.3) of the first heat exchanger (6) and a second portion (3.2) connecting a fuel outlet (6.4) of the first heat exchanger (6) to an inlet (4.1) of the engine (4), a heater (7) converting liquid fuel in the fuel conduit (3) to a low temperature gas form, a bypass conduit (8) connecting the first and second portions (3.1, 3.2) wherein the first heat exchanger (6) transfers heat from exhaust gases of the power unit outlet (5) to fuel of the fuel conduit (3).

An aircraft cooling system comprises an evaporator, condenser, an accumulator, and a pump system. The evaporator is configured to cool a set of heat loads in an aircraft using a liquid. The liquid forms a vapor in response to cooling the set of heat loads. The condenser is configured to receive the vapor from the evaporator and cool the vapor in which cooling the vapor forms the liquid. The accumulator is configured to receive the liquid from the condenser and store the liquid. The pump system is configured to pump the liquid stored in the accumulator to the evaporator.

A fuel cell system configured to provide electrical power to an electrical motor producing propulsive thrust for an aircraft includes a plurality of fuel cell stacks radially arranged about a central cavity. Fuel cell plates within each of the fuel cell stacks are oriented such that major surfaces of the fuel cell plates are parallel to an airflow created by the aircraft as it moves through the air. The fuel cell system further includes a plurality of cooling devices disposed in radial cavities located between at least two fuel cell stacks in the plurality of fuel cell stacks.

The invention relates to the technical field of hydrogen energy aviation, and discloses a hydrogen energy aviation fuel storage-supply system and method based on solid hydrogen storage, and the hydrogen energy aviation fuel storage-supply system and method based on solid hydrogen storage are characterized in that the hydrogen release amount adjusting range of a hydrogen storage material is matched with the hydrogen fuel demand adjusting range of a power device by adjusting the hydrogen release catalytic reaction environment temperature of the solid hydrogen storage material; by adopting the pressure stabilizing device, the problem of flow and pressure fluctuation caused by hysteresis of hydrogen release adjustment is solved, and hydrogen release flow adjustment and engine throttling control can be carried out through the thrust rod.

A fuel system for a gas turbine engine. The fuel system comprises a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit and at least first and second heat exchangers. The first heat exchanger is configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and the second heat exchanger is provided upstream in hydrogen flow of the first heat exchanger and is configured to transfer heat from a further heat exchange fluid to hydrogen fuel in the main fuel conduit. In an embodiment, the further heat exchange fluid is compressor bleed air bled from a core compressor of the gas turbine engine.

A fuel system for a hydrogen fueled gas turbine engine comprises a main hydrogen fuel storage unit configured to store liquid hydrogen. The main fuel storage unit comprises an ullage space configured to store ullage fluid comprising gaseous or supercritical hydrogen. The system further comprises a liquid hydrogen drain line configured to provide liquid hydrogen from the main hydrogen fuel storage unit to a fuel conduit, the fuel conduit being configured to supply hydrogen fuel to a combustor of the gas turbine engine. A priming line is provided, which is configured to provide ullage fluid to the fuel conduit, and a priming valve is configured to selectively control flow through the priming line.

An electrical power (EP) generation system (100) for an aircraft comprising a fuel cell (3), a cryogenic tank (1), a conditioning circuit (C1) configured to heat a liquid fuel stream (LH2) to a tank pressure (P2) to supply a gaseous fuel stream (GH2) to the fuel cell (3), a pressurization circuit (C2) to increase the gas pressure (P1) of the gaseous phase (PHL) in said cryogenic tank (1), the pressurization circuit (C2) comprising at least one flow control device (4), a gas pressure (P1) sensor (10) for measuring the gas pressure (P1) of the gaseous phase (PHG), and a computer (5) configured to control the flow control device (4) based on at least said gas pressure (P1) to control the tank pressure (P2). Abstract figure: Figure 1

An aircraft, such as a rotorcraft, may have at least one area designated to house at least one fuel cell and at least one aircraft structure that may translate, during a drop impact of the aircraft, into the area designated to house the fuel cell. At least one shaped fuel cell may be provided and deployed therein, in accordance with the present systems and methods. Each respective shaped fuel cell may define at least one respective through-voids defined through the respective shaped fuel cell, and / or at least one respective edge cavity defined along an edge of the shaped fuel cell, wherein the respective through-void and / or the respective edge cavity correspond to the respective aircraft structure that may translate, during the drop impact of the aircraft, into the area of the aircraft designated to house the respective fuel cell to receive and accommodate the respective structure during the drop impact.

A valve for a hydraulic circuit for an aircraft, the valve being configured to unite or separate multiple circuit portions in at least two configurations. More particularly, the valve allows to create one hydraulic circuit or two independent hydraulic circuits in order to circulate a coolant in two distinct manners and thus to adapt the function of the hydraulic circuit to any of two distinct phases of the functioning of an engine of the aircraft.

A system is provided for an aircraft. This aircraft system includes an engine, a fuel system and an air system. The fuel system includes a fuel circuit, a fuel source, a heat exchanger and a turboexpander. The fuel circuit is configured to direct fuel from the fuel source, through the heat exchanger and the turboexpander, to the engine for combustion. The heat exchanger is configured to transfer heat energy from combustion products generated by the combustion of the fuel in the engine into the fuel in the fuel circuit. The air system is discrete from the engine. The air system includes an air compressor and the turboexpander. The air system is configured to expand the fuel directed across the turboexpander by the fuel circuit to power operation of the air compressor.

The invention belongs to the field of target drone fuel oil systems, and particularly relates to a self-adaptive supersonic speed target drone fuel oil system and a control method thereof. The system comprises a machine body oil tank, a first electromagnetic valve, an overflow valve, a pressurization oil tank, a fuel pump, an engine pump, a monitoring module and a control module. The body fuel tank comprises a main fuel tank and an auxiliary fuel tank which are arranged in a fuel tank in the middle of the target drone body and located in front of and behind the theoretical mass center of the target drone respectively. The pressurization oil tank is additionally arranged in the fuel oil cabin, fuel oil in the engine body oil tank enters the pressurization oil tank for pressure maintaining after being pressurized by the fuel oil pump, and then oil is supplied to the engine from the pressurization oil tank, so that stable oil supply to the engine is achieved by establishing a local pressure environment instead of a large pressure environment of the whole fuel oil cabin, and therefore when a fuel oil system works, the fuel oil can be stably supplied to the engine. The oil supply flow and pressure are controlled in a closed-loop mode in real time through the control system according to the engine state.

A fuel system for a hydrogen fueled aircraft propulsion system comprises first and second hydrogen fuel tanks configured to store liquid hydrogen, first fuel line and second fuel lines configured to supply hydrogen from the first hydrogen fuel tank to a combustor of a first gas turbine engine and from the second hydrogen fuel tank to a combustor of a second gas turbine engine respectively. First and second fuel pumps are provided, each being configured to pump fuel in a respective first and second fuel line. First and second fuel heaters are provided, each being configured to heat fuel in a respective first and second fuel line. A fuel cross-feed fuel line is provided, which is configured to transfer fuel between the first and second fuel lines. The fuel cross-feed line is provided downstream in fuel flow of the first and second fuel heaters.

An aircraft thermal management system includes a high temperature circuit and a low temperature circuit. The high temperature loop is configured to manage a temperature of a fuel cell stack in the engine compartment using the engine compartment heat exchanger. The cryogenic circuit is configured to cool a plurality of thermal loads in the engine compartment using the thermal capacity of the liquid hydrogen.

An aircraft comprises a fuselage; wings; engines connected to the wings; a set of liquid hydrogen tanks; a vent located on the tail on the fuselage; a conduit system connecting the liquid hydrogen tanks to the vent; and a controller. The controller is configured to control the conduit system to remove gaseous hydrogen in the set of liquid hydrogen tanks to travel through the conduit system and exit at the vent in response to pressure in the set of liquid hydrogen tanks being greater than a specified tolerance.

A fuel system for a hydrogen fuelled aircraft propulsion system comprises a fuel tank configured to store hydrogen fuel, a main fuel conduit configured to provide fuel to a combustor of a gas turbine engine, a fuel heater comprising a catalytic combustor configured to catalytically combust a portion of the hydrogen fuel prior to delivery to the combustor and a heat exchanger configured to exchange heat between exhaust gases from the fuel heater and hydrogen fuel in the fuel conduit.

Apparatus, systems, and articles of manufacture are disclosed to power thermal management systems with heat of a working fluid therein. Example thermal management system include: a thermal transport bus loop fluidly coupled to at least one heat source exchanger and at least one heat sink exchanger; a turbomachine including a turbine and a compressor, the turbine and compressor rotatably interlocked via a shaft, the compressor coupled to the thermal transport bus loop, the turbine including an inlet and an outlet, the inlet connected to a first point of the thermal transport bus loop via a first flowline, the outlet connected to a second point of the thermal transport bus loop via a second flowline; and a control valve coupled to the first flowline, the control valve to adjust a mass flowrate of the heat exchange fluid in the first flowline based on a speed of the shaft.

A fuel conditioning system for an aerobic hydrogen engine, including: at least one hydrogen pump configured to increase the pressure of the liquid hydrogen delivered at the outlet of a tank, one or several heat exchangers configured to increase the temperature of the pressurized hydrogen, an air supply circuit, at least one combustion device configured to ensure a partial combustion of the hydrogen with air coming from the air supply circuit in order to produce a fuel including a gas mixture including gaseous hydrogen and devoid of oxygen.

An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

A fuel system for a hydrogen fuelled gas turbine engine comprises a tank configured to store hydrogen, a pump configured to provide pressurised hydrogen at an outlet thereof, and a pre-heater configured to heat at least a portion of pressurised hydrogen fuel downstream of the pump. A preheater return offtake is configured to return at least a portion of hydrogen fuel heated by the preheater to the tank.

A method of operating a gas turbine engine including: a core including a turbine, compressor, combustor to combust fuel, and core shaft connecting the turbine and compressor; a fan upstream of the core; a fan shaft; a bearing supporting the fan shaft; an oil loop system to supply oil to the bearing; and a heat exchange system including: an air-oil heat exchanger through which oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow to transfer heat between the oil and the fuel; a bypass pipe allowing a proportion of the oil to flow past one of the air-oil or the fuel-oil heat exchanger; and a bypass valve allowing the proportion of the oil sent through the bypass pipe to be varied. Under idle conditions, a heat transfer ratio of:rate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ air⁢ (kJkg-1)rate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ fuel⁢ (kJkg-1)is from 0 to 1.5.

A gas turbine engine includes a staged combustion system having pilot fuel injectors and main fuel injectors. A fuel delivery regulator controls delivery of fuel to the pilot and main fuel injectors, receives fuel from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic. In a transition range of operation between the pilot-only and the pilot-and-main ranges of operation, fuel is delivered to both the pilot and main fuel injectors at a transition staging ratio different from the pilot-and-main staging ratio. The fuel delivery regulator delivers fuel to one or both the pilot and main fuel injectors during the transition range of operation having a different fuel characteristic from fuel delivered to one or both the pilot and main fuel injectors during at least part of the pilot-and-main range of operation.

A fuel delivery device and an auxiliary power plant fuel system equipped with the same, the fuel delivery device being connected between a main fuel pipe and an auxiliary fuel pipe, the fuel delivery device comprising: a housing formed as a closed housing and having a peripheral wall portion, a first bottom plate close to an upstream side of a fuel flow path, and a second bottom plate close to a downstream side of the fuel flow path; the main fuel oil through pipe is arranged between the peripheral wall part of the shell and the main fuel oil pipe; the auxiliary fuel oil through pipe is arranged between the peripheral wall part of the shell and the auxiliary fuel oil pipe; the one-way valve is arranged in the auxiliary fuel oil through pipe and allows fuel oil in the auxiliary fuel oil through pipe to flow to the shell; and a fuel recharging mechanism which is provided in the main fuel pipe and which, when the pressure of the main fuel pipe decreases due to the opening of the auxiliary fuel pipe, recharges the fuel supplied from the auxiliary fuel pipe into the housing via the one-way valve into the main fuel pipe. Fuel pressure fluctuation of a main fuel pipe in the starting process of the auxiliary power device can be improved, and the starting reliability is improved.

A fuel system (205) for an aircraft propulsion system (103) comprising a gas turbine engine. The fuel system includes a fuel tank (104) configured to store compressed gaseous hydrogen pressures greater than 100 Bar at a temperature below 200 Kelvin and a tank heating system (224, 232, 234, 242, 302, 406) configured to raise the temperature of hydrogen fuel within the hydrogen tank (104). The tank heating system includes a primary tank heater comprising an oil to fuel heat exchanger (242) and an auxiliary tank heater comprising one or more of a recuperator (224), an electric heater (232), a gas turbine engine core compressor bleed-air powered heater (302, 402), and an auxiliary combustor (306).

A sub-cooler for a sub-cooling cryogenic refueling system is disclosed herein. The sub-cooler includes a first valve to separate flowing cryogenic fuel into a primary flowline and an auxiliary flowline, a second valve to reduce the saturated pressure and temperature of the cryogenic fuel in the auxiliary flowline, a cryogenic heat exchanger to transfer heat from the primary flowline to the auxiliary flowline, a temperature sensor to measure the temperature of the sub-cooled cryogenic fuel in the primary flowline, and a sub-cooler controller to control the effective areas of the primary flowline inlet and the auxiliary flowline inlet at the first valve.

A turbine engine for an aircraft includes a fuel delivery assembly for a hydrocarbon fuel to flow therethrough, a combustor combusting the fuel to generate combustion gases, and a core air exhaust nozzle exhausting the combustion gases from the turbine engine. The turbine engine also includes a contrail mitigation system having a heater and a fuel precipitate separator. The heater is selectively operable to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel, and the fuel precipitate separator separates the fuel precipitates generated by the heater from the fuel. A controller is coupled to the heater to operate the heater to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel in response to a contrail mitigation input.

A fuel conditioning system configured to supply an aircraft turbine engine with fuel from a cryogenic tank, the fuel circuit comprising a buffer tank (R2) for supplying the turbine engine and a plurality of compression modules (1A-1D) configured to supply the buffer tank (R2), each compression module (1A-1D) comprising an elementary tank (3A-3D) of fixed volume, an elementary heat source (2A-2D) configured to increase the temperature of the fuel in the elementary tank (3A-3D) in an isochoric manner, an inlet valve (V1A-V1D) connecting the elementary tank (3A-3D) to an upstream part of the fuel circuit (CQ), an outlet valve (V3A-V3D) connecting the elementary tank (3A-3D) to the buffer tank (R2) and a venting valve (V3A-V3D) connecting the elementary tank (3A-3D) to the cryogenic tank via a return circuit (CR) in which a gaseous stream (G) circulates.

A fuel system for a gas turbine engine including a compressor section, a combustion section, and a turbine section includes a fuel tank for storing a fuel, a plurality of heat exchangers downstream of the fuel tank, and a valve downstream of the plurality of heat exchangers. The valve includes a fuel inlet in fluid communication with the plurality of heat exchangers, a first fuel outlet in fluid communication with a first fluid pathway, and a second fuel outlet in fluid communication with a second fluid pathway. The fuel system includes a fuel cell including an anode inlet and an anode outlet. The anode inlet is in fluid communication with the first fluid pathway. The fuel system also includes a combustion chamber of the combustion section in fluid communication with the second fluid pathway and the anode outlet of the fuel cell and a controller communicatively coupled to the valve.

To stabilize an engine operation which uses hydrogen fuel of a supercritical phase.SOLUTION: A hydrogen fuel feed system 10 for an aircraft equipped with an engine comprises: a fuel tank 21 that stores hydrogen fuel in a liquid phase; a fuel feed path 22 through which the fuel tank 21 is connected to the engine; a pressure pump 26 provided on the fuel feed path 22 to pressure the hydrogen fuel equal to or higher than critical pressure; and a sensor 29 provided downstream of the pressure pump 26 on the fuel feed path 22 to measure the pressure of the hydrogen fuel passed through the pressure pump 26.SELECTED DRAWING: Figure 3