206results about "Conditioning fuel arrangements" patented technology

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.

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.

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.

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 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.

Methods, apparatus, systems, and articles of manufacture are disclosed for a hydrogen fuel distribution system (100) with integrated thermal management. An example fuel distribution system (100) with integrated thermal management comprises a hydrogen fuel distributor (402) including a first tank (406) and a second tank (408). The hydrogen fuel distributor (402) circulates hydrogen fuel to a first fuel cell (606) and a second fuel cell (608). A coolant distributor (702) includes a heat exchanger (416). The coolant distributor (702) regulates a temperature of the fuel distribution system (100). A compressed air distributor (706) provides compressed air from a propulsor (730) to the first fuel cell (606) and the second fuel cell (608). An electrical distributor (602) transports electricity from the first fuel cell (606) and the second fuel cell (608) to an electric motor (726). The electric motor (726) drives the propulsor (730).

A fuel system having a hydrogen fuel tank for fueling a fuel feed and propulsion system. The fuel system having a jet pump within the hydrogen fuel tank, at least one primary pump receiving fuel from the jet pump and pressurizing fuel to an outlet path and a first fuel recirculation path configured to selectively recirculate fuel to the hydrogen fuel tank. The first fuel recirculation path is fluidly connected to the outlet path. The first fuel recirculation path includes a heat exchanger positioned within the hydrogen fuel tank to cool fuel received from at least one of the fuel feed system or a tank pressurization system. The heat exchanger and a Joule-Thomson expansion at the jet pump minimizes recirculated fuel temperature, which in combination with a selective flow recirculation rate improve the life of the primary pump downstream.

Hybrid propulsion systems that utilize liquid natural gas solid oxide fuel cells in a manner practical for use in aircraft that avoid the use of heavy batteries, provide transient response times suitable for use in aircraft, and / or simplify reactant pre-conditioning systems using a compressor and turbine pair operatively coupled to the solid oxide fuel cell. Such hybrid propulsion systems for an aircraft may include a liquid natural gas turboshaft engine, a turbo generator operatively coupled to the turboshaft engine, a liquid natural gas solid oxide fuel cell, and an electric fan. The electric fan may be configured to generate thrust for the aircraft and may be powered by the turbo generator and / or the solid oxide fuel cell. Fuel cell reactants may be pre-conditioned using turboshaft bleed air, and the liquid natural gas may be pre-heated by fuel exhaust from the solid oxide fuel cell.

A heating device comprises a thermally-conductive outer tube, and a combustor arranged to combust air and gaseous hydrogen input to the device and introduce resulting combustion products into a first end of the outer tube. The device includes an igniter arranged to ignite the air and gaseous hydrogen. Combustion products produced by the combustor flow from the combustor to an output of the device along a flow path such that they are in contact with the internal surface of the outer tube over at least a portion of the flow path. Apparatus for heating a flow of hydrogen comprises a conduit for conducting the flow of hydrogen from a principal input port to a principal output port, the apparatus further comprising the heating device located at least partially within the conduit between the principal input port and the principal output port.

An aircraft fuel system includes a first integrated boost stage having a boost pump configured to receive a fuel flow from a fuel tank via a boost supply line, at least one generator collocated with and operatively connected to the boost pump for driving the boost pump, and a tap fluidly connecting the boost pump and the at least one generator for selectively delivering an amount of the fuel flow to the at least one generator. The fuel system further includes a gearbox operatively connected to the at least one generator, and a main stage downstream of the boost stage and operatively connected to the gearbox. The main stage includes a main pump configured to receive the fuel flow from the boost pump along a main supply line, and a filter disposed along the main supply line upstream of the main pump for filtering the fuel flow.

Systems and methods of operating a system are provided. For example, a system includes a fuel cooling loop including a cold fuel flow path having fuel flowing therethrough, a fuel cooler heat exchanger for cooling the fuel in fluid communication with the cold fuel flow path, and a cold fuel tank disposed along the cold fuel flow path for accumulating at least a portion of the cooled fuel. The system further includes a fuel heating loop including a hot fuel flow path for flow of the fuel, a fuel heater heat exchanger for heating the fuel in fluid communication with the hot fuel flow path, and a hot fuel tank disposed along the hot fuel flow path for accumulating at least a portion of the heated fuel. The fuel cooling loop is coupled to the fuel heating loop such that the fuel circulates through both the fuel cooling loop and the fuel heating loop.

Hydrogen supply device comprising at least two tanks configured to store hydrogen in liquid and supercritical states, respectively. The invention relates to a hydrogen supply device (18) comprising: at least one first tank (20) configured to store hydrogen in a liquid state at a pressure below 13 bar; at least one second tank (22) configured to store hydrogen in a supercritical state at a pressure greater than or equal to 13 bar; at least one upstream portion (26) comprising at least one first hydrogen flow control system (32, 32', 32'') configured to regulate the amount of hydrogen flowing from the first tank (20) to the second tank (22); at least one downstream portion (28) connecting the second tank (22) and a hydrogen-powered system (16). The invention also relates to an aircraft comprising at least one such supply device. Figure 2

A system (22) is disclosed having an inerting device (44) that receives an intake air stream and separates it into nitrogen-enriched air (NEA) and oxygen-enriched air (OEA). The inerting system also has an oxygen concentration sensor (46) and a controller (48) connected to the oxygen concentration sensor (46). The controller (48) adjusts operation of the inerting system (22) during use based at least in part on feedback received from the oxygen concentration sensor (46).

The invention relates to a fuel conditioning system (SC) configured to supply an aircraft propulsion system with fuel (Q) from a cryogenic tank (R), the conditioning system (SC) comprising a plurality of compression modules (2A-2C) configured to increase the temperature of the fuel (Q) in an elementary tank (3A-3C) in an isochoric manner, and a metering tank (5) having a fixed calibrated volume (VOLc); wherein each compression module (2A-2C) comprises an inlet valve (V1A-V1C) for supplying the elementary tank (3A-3C) with a volume of fuel (Q) corresponding to the calibrated volume (VOLc) of the metering tank (5), an outlet valve (V2A-V2C) and a degassing valve (V3A-V3C) connected to the elementary tank (3A-3C) and configured to circulate a gaseous fuel flow (G) through a return circuit (CR).

The invention relates to a system (1) for controlling the temperature of a heat transfer fluid (F) configured to transfer heat to a fluid to be heated (Q) originating from a cryogenic tank (R), the control system (1) comprising: a loop (2) for circulating the heat transfer fluid (F), comprising an engine branch (21) and a tank branch (22); a first engine heat exchanger (41), configured to heat the heat transfer fluid (F) to a second temperature (T2) above a maximum operating temperature (Tmax); a mechanical pump (3) configured to circulate the heat transfer fluid (F) in the circulation loop (2), such that, in a second engine heat exchanger (42), a first part of the heat is transferred from the heat transfer fluid (F) to the fluid to be heated (Q) and the heat transfer fluid (F) is cooled to a third temperature (T3) below the maximum operating temperature (Tmax), before it leaves the engine enclosure (EN-M).

An aircraft fuel system includes a first fuel storage tank arranged to store a first fuel, a fuel oxygen reduction unit arranged to generate a deoxygenated fuel from the first fuel, and a second fuel storage tank arranged to store the deoxygenated fuel generated by the fuel oxygen reduction unit, and being arranged to supply the deoxygenated fuel to an engine.

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.

The invention relates to an aircraft comprising: - a fuselage (32), - a wing (34), - at least one hydrogen-powered turbomachine (40) generating thrust at a propulsion unit (36) located away from the fuselage (32), - at least one fuel tank (46) positioned in the fuselage (32) and configured to store hydrogen in a cryogenic state, - at least one hydrogen supply system connecting the fuel tank (46) and the turbomachine (40) comprising: o at least one pump (50) positioned in the fuselage (32) near the fuel tank (46), o at least one hydrogen heating system (52) positioned in the fuselage (32) near the pump (50). This solution makes it possible to reduce the length of the complex double-walled conduits configured to channel the hydrogen in a cryogenic state between the fuel tank and the hydrogen heating system.

A fuel transfer device and an auxiliary power unit fuel system equipped with the fuel transfer device are disclosed. The fuel transfer device is connected between a main fuel line and an auxiliary fuel line, and includes: a housing, which is formed as a sealed housing, having a peripheral wall portion, a first base plate near the upstream side of the fuel flow path, and a second base plate near the downstream side of the fuel flow path; a main fuel line, which is disposed between the peripheral wall portion of the housing and the main fuel line; an auxiliary fuel line, which is disposed between the peripheral wall portion of the housing and the auxiliary fuel line; a one-way valve, which is disposed in the auxiliary fuel line, allowing fuel in the auxiliary fuel line to flow to the housing; and a fuel replenishment mechanism, which is disposed in the main fuel line, and replenishes the main fuel line with fuel supplied from the auxiliary fuel line to the housing via the one-way valve when the pressure in the main fuel line drops due to the opening of the auxiliary fuel line. This improves fuel pressure fluctuations in the main fuel line during the starting process of the auxiliary power unit, thereby increasing starting reliability.

A fuel system (200) for a hydrogen fuelled aircraft propulsion system (103) comprises a tank (104) configured to store hydrogen, a first inductor (436a-c) provided within the tank (104) and configured to heat hydrogen fuel within the tank (104), and a second inductor (432a-c) provided externally to the tank (104), and configured to induce a current in the first inductor (436a-c).

A hydrogen fuel supply system is a hydrogen fuel supply system of an aircraft including an engine and includes: a fuel tank storing hydrogen fuel in a liquid phase; a fuel supply passage connecting the fuel tank to the engine; a pressure pump that is located in the fuel supply passage and pressurizes the hydrogen fuel to critical pressure or more; and a sensor that is located downstream of the pressure pump in the fuel supply passage and measures pressure of the hydrogen fuel which has passed through the pressure pump.

The invention discloses an aircraft fuel oil thermal management system based on a variable consumable heat sink, which comprises a fuel oil storage tank, a first heat exchanger, a second heat exchanger, a third heat exchanger, a fourth heat exchanger, a distribution valve and a fifth heat exchanger which are sequentially connected through an oil path pipeline, and a heat sink flow path sequentially passes through a consumable heat sink storage tank, the fifth heat exchanger and a control valve. Fuel oil in the first heat exchanger exchanges heat with a heat source of an FAEDC system, fuel oil in the second heat exchanger exchanges heat with a heat source of a fuel oil pump, fuel oil in the third heat exchanger exchanges heat with a heat source of airborne equipment, and fuel oil in the fourth heat exchanger exchanges heat with a heat source of an engine. And the high-temperature fuel oil in the fifth heat exchanger exchanges heat with the consumable heat sink. Airborne fuel oil and consumable heat sinks are reasonably utilized, the fuel oil amount needed by an aircraft engine is guaranteed by adjusting the fuel oil flow and selecting the heat sink type and flow, the heat dissipation purpose is achieved, the situation that the combustion performance of the aircraft engine is affected by the too high fuel oil temperature is avoided, and the service life of the aircraft engine is prolonged.

The invention relates to a method for determining a first heat quantity (ThA) and a second heat quantity (ThM) to be provided in a first heating module (31) and a second heating module (31) of a fuel conditioning system (SC), the method comprising a step of determining an overall range of second enthalpy (HC2) of a fuel flow (Q) from the formulas γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073 and γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778; and a step of determining, in a database (BdD) of elementary ranges of second enthalpies (HC2), a pair of heat quantities (ThA; ThM), the elementary range of which is entirely within the overall range so as to determine the first heat quantity (ThA) and the second heat quantity (ThM).

A fuel supply system for an aircraft according to the present invention comprises: an intermediate tank which is fluidly connected to a fuel pump and an engine via a fuel supply path; a pressure sensor which detects the internal pressure of the intermediate tank or a tank pressure that is correlated with the internal pressure of the intermediate tank; and a processing circuit which executes intermittent control for controlling a pump actuator in accordance with the tank pressure. The intermittent control includes: when the tank pressure increases and reaches a first threshold value, controlling the pump actuator to stop the fuel pump in a state where supply of a fuel to the engine from the intermediate tank is allowed; and when the tank pressure decreases and reaches a second threshold value which is lower than the first threshold value, controlling the pump actuator to drive the fuel pump.

AIRCRAFT PROPULSION SYSTEM COMPRISING A PIPING CIRCUIT AT LEAST PARTIALLY VENTILATED. The invention relates to a propulsion system (100) comprising: a fuel tank (520); a turbojet engine (102) comprising an engine and a fan blowing an airflow (F); a piping circuit (20) comprising a single-wall duct (21) extending from the tank and a second piping portion (23) comprising two double-wall ducts (230a, 230b) fluidly connecting the single-wall duct and a connection interface with the engine. The second portion also comprises a second connector (24) disposed between the two double-wall ducts and an air supply duct (25) fluidly connecting the fan and the double-wall ducts so as to draw a portion (F1) of the airflow (F) from the fan into the double-wall ducts and cool the latter. Fig. 3

An aircraft turbine engine comprising:-a fuel supply system (12) for supplying fuel to a combustion chamber of the turbine engine; -a heat pump (15) comprising a closed circuit (20) in which a heat transfer fluid circulates, the circuit (20) comprising: an evaporator (21) configured to exchange heat with the oil of the turbine engine; a first condenser (22) configured to exchange heat with fuel of the supply system (12); a second condenser (23) mounted in parallel with the first condenser (22) and configured to exchange heat with a cold source (24), the cold source being different from the fuel; an expansion valve (25) configured to expand the heat transfer fluid before it enters the evaporator (21); a compressor (26) configured to compress the heat transfer fluid before it enters the condenser (22, 23); and a metering device (27) configured to regulate the flow of the heat transfer fluid entering the first condenser (22).

An aircraft includes a first tank and a second tank each storing low-temperature liquid fuel; a pump that receives supply of the low-temperature liquid fuel from the first tank and the second tank; and a controller. Each of the first tank and the second tank is switched between a fuel supply phase of supplying the low-temperature liquid fuel to the pump and a temperature restoring phase of cooling the low-temperature liquid fuel in each of the first tank and the second tank. The controller switches between a first mode where the first tank is in the fuel supply phase and the second tank is in the temperature restoring phase, and a second mode where the first tank is in the temperature restoring phase and the second tank is in the fuel supply phase.

An aircraft includes a first tank and a second tank each storing low-temperature liquid fuel; a pump that receives supply of the low-temperature liquid fuel from the first tank and the second tank; and a controller. Each of the first tank and the second tank is switched between a fuel supply phase of supplying the low-temperature liquid fuel to the pump and a temperature restoring phase of cooling the low-temperature liquid fuel in each of the first tank and the second tank. The controller switches between a first mode where the first tank is in the fuel supply phase and the second tank is in the temperature restoring phase, and a second mode where the first tank is in the temperature restoring phase and the second tank is in the fuel supply phase.