26results about "Conditioning fuel arrangements" patented technology

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.

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).

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.

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.

A fuel injector for a combustor of a gas turbine engine is described. The fuel injector includes a fuel injection device having a fuel injection port configured to inject fuel into the combustor, wherein the fuel injection port includes an elongate aperture. A combustor for a gas turbine engine, a reheat assembly for a gas turbine engine, a gas turbine engine, and an aircraft.

Improved cryogenic fluid supply device for an aeronautical turbojet engine. Cryogenic fluid supply device (10) for an aeronautical turbojet engine (1), a tank (20) storing the fluid in liquid form, the device (10) comprising a first heat exchanger (30) capable of first heating the fluid drawn from the tank (20) via a sampling line (11) connecting the tank (20) to the first heat exchanger (30), and a main line (12) connected in a loop to the first heat exchanger (30), the device (10) comprising a bypass channel (14) comprising a first branch (141) branching from the main line (12) at a bypass point (140), to a mixer (50) capable of mixing a liquid phase and a gaseous phase of the fluid, and a second branch (142) connecting the mixer (50) to the sampling line (11), a bypass valve (60) disposed downstream of the bypass point. (140),and movable between a first position allowing a deviation of the fluid flow via the bypass channel (14), and a second position limiting this deviation. Fig. 3.

There is provided a reheat assembly (300A, 300B, 300C) for a gas turbine engine (10). The reheat assembly (300A, 300B, 300C) comprises a jetpipe casing (310), a support duct (340), a plurality of fuel discharge ports (362), and a fuel distribution passageway (342). The jetpipe casing (310) comprises a reheat core section (320) and a reheat bypass section (330). The support duct (340) radially separates the reheat core section (320) and the reheat bypass section (330). Each fuel discharge port (362) is configured to discharge fuel received from the fuel distribution passageway (342). The fuel distribution passageway (342) is defined by an interior channel embedded within the support duct (340).

A non-propulsive turbomachine (2) for an aircraft fuel conditioning system (Q) comprising: a first compressor (21) configured to supply a cabin air conditioning device (6) with a first pressurized airflow (A1p) via a circulation channel (31); a second compressor (22); a combustion chamber (23) configured to be supplied by a supply airflow (A) circulating in a supply channel (32) connecting to the second compressor (22); a single gas turbine (24) connected to the first compressor (21) and the second compressor (22) by a single turbomachine shaft (25); a bypass channel (33) connected to the circulation channel (31) and the supply channel (32); and a relief valve (4) mounted on the bypass channel (33) and configured to move between: a closed position and an open position, in which part of the first pressurized airflow (A1p) flows in the bypass channel (4).Figure from the summary: Figure 3.

The application discloses a temperature self-adaptive control method of a thermal management system, belonging to the field of aviation. The dynamic scheduling of the heat sink flow is realized through the self-adaptive control method, the flow of each branch heat sink is controlled according to the temperature and heat dissipation data acquisition information of the equipment, and the refrigeration capacity of the heat sink is dynamically adjusted. The self-adaptive control method realizes the accurate control of the heat sink outlet temperature of the equipment / heat exchanger and the like through the dynamic scheduling of the heat sink, and prevents the violation of the temperature constraint. The self-adaptive control method controls and optimizes the heat flow process, guarantees the efficient operation of the system, and can reduce the size and power consumption of the thermal management system. According to the measurement results of the temperature sensors on the equipment and pipelines and the working condition of the equipment, the controller adjusts the rotating speed of the pump and the opening degree of the electric flow regulating valve, so that the system has the self-adaptive control capability.

The application provides a small negative feedback constant flow starting valve assembly, and belongs to the field of structural design of aerospace mechanical hydraulic control systems, which comprises a valve bushing, a starting valve, an adjusting spring and a throttle nozzle. The starting valve is arranged in the valve bushing, and the inside of the valve bushing is divided into a spring cavity and a control cavity by the starting valve. The adjusting spring is arranged in the spring cavity. The starting valve is further provided with a flow guide channel, and the throttle nozzle is arranged in the flow guide channel. An oil inlet bushing type hole and an oil outlet hole are arranged on the side wall of the valve bushing, and a first flow guide hole is arranged on the side wall of the starting valve, the oil inlet bushing type hole is connected with the control cavity through the first flow guide hole, and the oil outlet hole is connected with the spring cavity. The small negative feedback constant flow starting valve assembly can ensure that the pressure difference on both sides of the throttle nozzle is basically constant, and a stable flow supply effect is formed.

A heat transfer device for an aircraft providing heat to the fuel of a fuel conduct, a conditioning system having the heat transfer device, an aircraft including the heat transfer device of the fuel conditioning system, and a method for conditioning fuel of an aircraft. The fuel conduct includes a first end and a second end, in which the first end is connected in fluidic communication to the first inlet of the heat transfer device and the second end is connected in fluidic communication to the first outlet of the heat transfer device.

There is provided a reheat assembly for a gas turbine engine. The reheat assembly comprises a jetpipe casing, a support duct, a plurality of fuel discharge ports, and a fuel distribution passageway. The jetpipe casing comprises a reheat core section and a reheat bypass section. The support duct radially separates the reheat core section and the reheat bypass section. Each fuel discharge port is configured to discharge fuel received from the fuel distribution passageway. The fuel distribution passageway is defined by an interior channel embedded within the support duct.

The foregoing describes a reheat assembly for a gas turbine engine. The reheat assembly comprises: a support duct section comprising a core side for facing a core reheat region, and a bypass side for facing a bypass reheat region; the support duct comprising a plurality of circumferentially spaced inlets, each inlet being configured to communicate the bypass reheat region with the core reheat region; a fuel supply system comprising a plurality of bypass fuel injection ports, each bypass fuel injection port being associated with an inlet of the plurality of inlets; wherein each bypass fuel injection port is configured to discharge fuel into the bypass reheat region for transit through the associated inlet into the core reheat region.

This invention relates to a heat transfer device (1) for an aircraft (100), in particular for providing heat to the fuel of a fuel conduct (4). The present invention also related to a conditioning system comprising a heat transfer device (1) thereof, an aircraft (100) comprising a heat transfer device (1) of a fuel conditioning system thereof and a method for conditioning fuel of an aircraft (100).

The foregoing describes a reheat assembly (300) for a gas turbine engine (100). The reheat assembly (300) comprises: a support duct section (321') comprising a core side (322') for facing a core reheat region (232), and a bypass side for facing a bypass reheat region (231); a support duct (21') comprising a plurality of circumferentially spaced inlets (380), each inlet (380) being configured to communicate the bypass reheat region (231) with the core reheat region (232); a fuel supply system (360) comprising a plurality of bypass fuel injection ports (363), each bypass fuel injection port (363) being associated with an inlet (380) of the plurality of inlets (380); wherein each bypass fuel injection port (363) is configured to discharge fuel into the bypass reheat region (231) for transit through the associated inlet into the core reheat region (232).

A conditioning system for an aircraft including a fuel conduit which receives a liquid fuel, an evaporator and a heater which heat the liquid fuel to convert the fuel to a gas fuel supplied to an engine on the aircraft, the evaporator transfers heat from a hot working fluid to the liquid fuel and the heater converts energy into heat applied to the liquid fuel, and a heat recovery circuit which circulates the working fluid through the evaporator and a heat exchanger in a hot exhaust gas stream generated by the engine.

A fuel mixing assembly is disclosed including a fuel vessel, such as a pipe for conveying flow of a first fuel, and a second-fuel conduit for conveying flow of, for example an additive, into the fuel vessel. The fuel mixing assembly further includes a vortex generator, the vortex generator is configured to generate vortices to mix the first fuel and the additive. The additive may be released from an outlet in the vortex generator.

A fuel conditioning system (SC) supplying at least one aircraft turbomachine (T) with fuel (Q) from a cryogenic tank (RC), the conditioning system (SC) comprising a fuel circuit (CQ) connected inlet to the cryogenic tank (RC) and outlet to the turbomachine (T), a purge circuit (CP) comprising an inlet connected to a purge point (PP) and an outlet connected to an external environment (EXT) of the aircraft, the purge circuit (CP) comprising at least one purge tank (RP) configured to store the fuel flow present in the portion to be purged (PAP) during a purge and at least one supply branch (B1), connecting the purge tank (RP) to the fuel circuit (CQ) or to a fuel-consuming device. Abstract Figure: Figure 2

An aircraft fuel tank venting and inerting arrangement is provided. It comprises a fuel tank (2, 3, 4, figure 1) , a venting fluid flow path between the fuel tank and a venting outlet, an inerting flu

Improved cryogenic fluid supply device for an aeronautical turbojet engine. Cryogenic fluid supply device (10) for an aeronautical turbojet engine (1), a tank (20) storing the fluid in liquid form, the device (10) comprising a first exchanger (30) providing initial heating of the fluid drawn via a sampling duct (11), a main duct (12) connected in a loop to the first exchanger (30), a bypass channel (14) comprising a first branch (141) branching from the main duct (12) at a bypass point (140), up to a phase separator (50), a second branch (142) connecting the liquid phase of the phase separator (50) to the sampling duct (11), and a third branch (143) connecting the gaseous phase of the phase separator (50) to the main duct (12) downstream of the bypass point (140), a bypass valve (60) disposed downstream of the bypass point. (140),and mobile between a first position diverting the fluid flow via the bypass channel (14), and a second position limiting this deviation. Fig. 3.

A fuel injector for a combustor (15, 300) of a gas turbine engine (10) is described. The fuel injector comprises a fuel injection device (360) having a fuel injection port (362) configured to inject fuel into the combustor, wherein the fuel injection device is configured to promote cavitation of the fuel. Also disclosed is a combustor (15) for a gas turbine engine (10), a reheat assembly (300) for a gas turbine engine (10), a gas turbine engine (10), and an aircraft (200).

A fuel mixing assembly comprising, a fuel vessel, such as a pipe 15 for conveying flow of a first fuel; and a second-fuel conduit 17 for conveying flow of, for example an additive, into the fuel vessel 15. The fuel mixing assembly further comprises a vortex generator 21, the vortex generator 21 being configured to generate vortices to mix the first fuel and the additive. The additive may be released from an outlet in the vortex generator 21.

The present invention provides a heat transfer device for an aircraft, a fuel conditioning system and an aircraft. The present invention relates to a heat transfer device for an aircraft, in particular a heat transfer device for providing heat to fuel of a fuel conduit. The present invention further relates to a conditioning system comprising a heat transfer device thereof, an aircraft comprising a heat transfer device of a fuel conditioning system thereof and a method for conditioning fuel of an aircraft.

A hydrogen aircraft utilizes hydrogen as fuel for a power unit. The hydrogen aircraft includes target piping and a cooling device. The target piping is piping for transportation of liquid hydrogen, and has a stagnation period during which liquid hydrogen does not substantially flow while the hydrogen aircraft is in operation. The cooling device cools the target piping.