413results about "Power installations" patented technology

An integrated hydrogen-electric engine includes, an air compressor system; a hydrogen fuel source; a fuel cell; an elongated shaft connected to the air compressor system or a propulsor; and a motor assembly disposed in electrical communication with the fuel cell, wherein the air compressor system includes a plurality of electrically driven compressors configured to run in series or parallel.

The present invention relates to an aircraft propulsion system (100) comprising: a fuel cell arrangement comprising at least one fuel cell (110); an air source (130) for providing air to the fuel cell arrangement; a compressor arrangement comprising a first compressor (120) in fluid communication with the air source and a fuel cell of the fuel cell arrangement; and, a turbine arrangement comprising a first turbine (124) mechanically coupled to the first compressor, wherein the first turbine is in fluid communication with the at least one fuel cell (110), the system being arranged so that, in use, air from the air source (130) flows in turn to the first compressor (120), the fuel cell arrangement and the first turbine (124).

The invention provides an APU characteristic correction method and device, and belongs to the field of aviation electromechanics. The method provided by the invention comprises the following steps: establishing an APU system simulation model; determining a plurality of correction parameters and correction targets; calculating a calculation value corresponding to the correction target based on the APU system simulation model and the correction parameter; constructing a correction error function based on the calculated value of the correction target and a test value of the correction target obtained through a test; determining an updated component in the APU, and determining an optimization strategy based on the identity of the updated component; based on the optimization strategy, executing a particle swarm algorithm to obtain an optimal correction parameter corresponding to the correction error function; and updating the APU system simulation model based on the optimal correction parameter to obtain a corrected APU characteristic model. According to the APU characteristic correction method and device provided by the invention, the accuracy of the model correction result can be improved after the local parts of the APU are updated.

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 solid oxide fuel cell, a motor driven by electric power from the solid oxide fuel cell, a gearbox operatively coupled to the motor, and a turbofan engine configured to generate thrust for the aircraft. The turbofan engine may be configured to provide electric power and shaft power and may include a duct fan that is operatively coupled to the motor via the gearbox, with the duct fan being driven by mechanical power from the gearbox and by the motor.

An aircraft power plant, has: an air mover for propelling an aircraft; an electric motor drivingly engaged with the air mover; a gas turbine engine having a compressor for pressurizing air, a combustor in which the pressurised air is mixed with fuel and ignited for generating combustion gases, and a turbine for extracting energy from the combustion gases; and a hydrogen fuel cell operatively connected to the electric motor for powering the electric motor with electricity generated by the hydrogen fuel cell, the hydrogen fuel cell operable to generate electricity using the air from the compressor and hydrogen from a source of hydrogen.

A propulsion system for an aircraft as disclosed herein may include a nacelle, a shaft positioned centrally within a cylindrical passageway of the nacelle, a fan coupled to one end of the shaft, a turbine coupled to an opposite end of the shaft, an electric motor coupled to the shaft, a compressor positioned within the cylindrical passageway, and a solid oxide fuel cell positioned with a hollow ring-shaped interior of the nacelle. The hollow ring-shaped interior may surround and be isolated from the cylindrical passageway. The turbine may be configured to provide primary torque to the shaft while the electric motor may be configured to provide additional torque to the shaft. The electric motor may be powered an electric output of the solid oxide fuel cell while the turbine may be powered at least in part by output gases from the solid oxide fuel cell.

One embodiment is an aircraft having a least one of an aircraft motor or an aircraft motor drive, and a metal-air fuel cell. The aircraft has a waste heat transfer system including a diverter system. The diverter system selectively couples the metal-air fuel cell and the at least one of the aircraft motor or the aircraft motor drive. The aircraft has a control system to operate the waste heat transfer system to selectively transfer waste heat from the at least one of the aircraft motor or the aircraft motor drive to the metal-air fuel cell. The control system determines a power output status of the metal-air fuel cell, and to operate the diverter system to transfer the waste heat from the at least one of the aircraft motor or the aircraft motor drive to the metal-air fuel cell in response to determining the power output status.

A propulsion system for an aircraft as disclosed herein may include a nacelle, a shaft positioned centrally within a cylindrical passageway of the nacelle, a fan coupled to one end of the shaft, a turbine coupled to an opposite end of the shaft, an electric motor coupled to the shaft, a compressor positioned within the cylindrical passageway, and a solid oxide fuel cell positioned with a hollow ring-shaped interior of the nacelle. The hollow ring-shaped interior may surround and be isolated from the cylindrical passageway. The turbine may be configured to provide primary torque to the shaft while the electric motor may be configured to provide additional torque to the shaft. The electric motor may be powered an electric output of the solid oxide fuel cell while the turbine may be powered at least in part by output gases from the solid oxide fuel cell.

Disclosed is an aircraft structure comprising at least one aircraft structural element defining a volume; a fuel-handling component arranged within the volume and configured to handle a fuel; a purging system comprising an inlet and an outlet, the purging system configured to provide a flow of purge gas from the inlet, through the volume, to the outlet, to thereby purge leaked fuel from the volume; and a flow guide member formed of a reticulated foam, arranged within the volume and externally to the fuel-handling component, and configured to guide the flow of purge gas within the volume. An aircraft comprising the aircraft structure is also disclosed.

A powerplant for an aircraft includes a first turbine engine, a second turbine engine and a fuel cell system. The second turbine engine includes a second engine flowpath, a second engine compressor section, a second engine combustor section and a second engine turbine section. The second engine flowpath extends from a second engine flowpath inlet to a second engine flowpath outlet. The second engine flowpath inlet and the second engine flowpath outlet are each fluidly coupled with a flowpath of the first turbine engine. The fuel cell system includes a fuel cell, a fuel circuit and an air circuit. The fuel circuit extends through the fuel cell and is fluidly coupled with and upstream of a first fuel injector in the second engine combustor section. The air circuit extends through the fuel cell and is fluidly coupled with and downstream of a bleed from the second engine flowpath.

Electrical system (3) comprising: an electrical bus (30) suitable for being electrically connected to electrical loads, electrical sources (31,32,33,34) configured to supply the electrical bus (30) with respective powers, and local controllers (311,321,331,341) respectively associated with the electrical sources, each local controller (311,321,331,341) being configured to: receive a local voltage measurement acquired by a sensor associated with the local controller (311,321,331,341), generate a local estimate (Vdc) of a voltage (Vdc) of the electrical bus from the local measurement (Vdc) and a power information;obtain a reference estimate (Vest_ref) of the electrical bus voltage, the reference estimate (Vest_ref) being common to all local controllers (311,321,331,341), from the local measurement (Vimes), a reference setpoint (Vref) and a correction term (dVi) representing a difference between the local estimate () and the reference estimate (Vest_ref), generate a setpoint (Pci) to control the electrical source (31,32,33,34) associated with the local controller (311,321,331,341). Figure for the abbreviation: Fig. 5;

A fuel-cell-powered aircraft system has integrated air-cooled fuel cell stacks positioned within an interior space of at least one wing of an aircraft. An airflow path is positioned in contact with at least a heat exchanger of the fuel cell stack. The induced flow of air through the airflow path cools the heat exchanger. The efficiently-induced flow of air for cooling the fuel cell stack has a zero or minimal drag penalty.

A power management system for supplying power from an aircraft electrical power source to PEDs. The power management system includes a plurality of power control circuits connected to power outputs configured to supply power to PEDs, and a power management circuit. The power management circuit obtains passenger-specific travel information associated with a passenger corresponding to a respective PED, and determines, based at least in part on the passenger-specific travel information, an allowed subscription power level for the respective PED. The power management circuit controls the power supplied to the respective PED via the respective power control circuit based on the determined allowed subscription power level. The operation to control the power supplied to the respective PED may include to negotiate with the respective PED an agreement for an agreed subscription power level that the respective PED is authorized to be supplied based on the determined allowed subscription power level.

A thermal management system (TMS) 100 for an aircraft comprising an air inlet 310 which receives ambient air from an external environment, an air outlet (120, figure 1) providing fluid communication b

A propulsion system for an aircraft includes a fan section having a fan; a turbomachine including a compressor section, a combustion section, and a turbine section arranged in serial flow order; at least one electric machine; a fuel cell assembly configured to supply power to the electric machine; and a controller including a memory and one or more processors. One or both of the turbomachine and the at least one electric machine are configured to drive rotation of the fan of the fan section. The fuel cell assembly includes at least one fuel cell, a first fluid inlet for receiving a flow of pressurized air, and a second fluid inlet for receiving a flow of fuel.

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.

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.

The invention relates to a fuel system (1) for supplying an aircraft (2) with hydrogen (3). The fuel system (1) comprises a first inner tank (4) and a separate second inner tank (6), each for storing hydrogen (3), and an outer tank (5) which surrounds the inner tanks (4, 6). The first inner tank (4) and the second inner tank (6) are fluidically connected, separately from one another, to at least one first supply device (7) for supplying the aircraft (2) with hydrogen (3). The invention also relates to an aircraft (2) having a fuel system (1), and to a method.

An aircraft power generation system includes: a turbo fan engine including a variable stator vane which is located upstream of a fan and whose installation angle is variable and a vane actuator that drives the variable stator vane; a power generator mechanically connected to a low-pressure shaft of the engine; and processing circuitry configured to control the vane actuator. The processing circuitry is configured to, when it is determined that the power generator generates electric power during an operation of the turbo fan engine, execute an electric power generating mode of controlling the vane actuator to change the installation angle of the variable stator vane such that the flow rate of air flowing through the fan decreases.

Aircraft propulsion system (1) (2) comprising a propeller (3) driven in rotation by an electric machine (4), the electric machine (4) being powered by a fuel cell (5) comprising a core (6), the core (6) being cooled by a cooling system (7) comprising a first heat exchanger (8) located in a first ducted channel (9) through which air flows, characterized in that the cooling system (7) comprises a second heat exchanger (8) located in a second ducted channel (9) through which air flows, the heat exchangers (8) and the channels (9) being annular, the channels (9) surrounding a central air-supplied compartment (12) in which the electric machine (4) and the fuel cell core (6) (5) are located, the inlets (10) of the channels (9) preferably each having a variable inlet cross-section. Figure for the abstract: 1

The application belongs to the technical field of high Mach number aircraft power system design, and particularly relates to a high Mach number aircraft power system, on the basis of a Mach number two aircraft power system, an APU+EPU scheme is adopted for the auxiliary power system, and a ram turbine is additionally arranged, the auxiliary power system can be used as an emergency energy source when the flight speed exceeds 2 Mach, main engine in-flight accidental shutdown and other abnormal conditions occur, so as to ensure the flight safety of the aircraft, in addition, the auxiliary power system can provide reliable auxiliary energy for the aircraft during aircraft mode conversion and in-flight unpowered taxiing, so as to keep the energy supply stable and ensure the continuous and reliable operation of the aircraft.

Disclosed is an aircraft accessory drive system (100) comprising: a drive shaft (102) configured to provide a drive input; a critical accessory (106) directly coupled to the drive shaft so as to be directly driven by the drive shaft; a parasitic drive system (110) configured to transmit drive input from the drive shaft to one or more less critical accessories, wherein the parasitic drive system comprises a torque limiter (112) configured to decouple the parasitic drive system from the drive shaft and the critical accessory in response to a jam affecting the parasitic drive system or one or more of the less critical accessories. Also disclosed are a gas turbine engine comprising an aircraft accessory drive system, and an aircraft comprising an aircraft accessory drive system.

Electrical power supply system of an aircraft electric propulsion system. The electrical power supply system (10) of an aircraft electric propulsion system (4) (1) comprises: - at least two first fuel cell assemblies (FC11, FC12), intended to electrically power an electric motor driving a propulsion propeller; - a first compressor (C1) configured to supply compressed air to at least some of the at least two first fuel cell assemblies (FC11, FC12); and - a first electric motor (M1) intended to drive the first compressor. The first electric motor (M1) comprises at least two electrical windings (W1a, W1b), each electrical winding being electrically powered from a separate fuel cell assembly among the at least two first fuel cell assemblies (FC11, FC12). Figure for the abstract: Fig. 2

A power distribution system and method of operation for an aircraft includes determining, by a supplemental power system, an operation mode of the aircraft, selecting, by the supplemental power system, at least one supplemental power source of a set of at least two supplemental power sources, and at least one of receiving a first current from the main power bus by the selected at least one supplemental power source, or selectively providing a second current to the main power bus from the selected at least one supplemental power source, based on an operating mode of the aircraft.

The invention relates to a turboprop (10) comprising a propeller (12), a propeller shaft (13) carrying the propeller (12), the propeller being a variable-pitch propeller having a propeller pitch, a rotating electric machine (19) having at least a first configuration in which it is mechanically coupled to the propeller shaft (13) and at least one oil pump (21, 21B) configured to supply a hydraulic circuit for adjusting the pitch of the propeller (12). The oil pump is configured to be electrically operated. The invention further relates to an aircraft comprising such a turboprop (10) and to the methods for controlling such a turboprop (10) and such an aircraft.

A power generation system for an aircraft, the power generation system comprising: a hydrogen fuel cell; an air supply system configured to supply air to the aircraft, the air supply system comprising an ORC-air heat exchanger; and an organic Rankine cycle (ORC) system comprising an organic fluid circuit, wherein the organic fluid circuit is configured to receive excess fuel cell heat from the hydrogen fuel cell to heat organic fluid in the organic fluid circuit, wherein the organic fluid circuit passes through the ORC-air heat exchanger such that the supply air heats the organic fluid in the organic fluid circuit, wherein the ORC system is configured to extract power from the organic fluid.

A power unit of an unmanned aerial vehicle (UAV) comprises an internal combustion engine (1) with a propeller (3) arranged on its output shaft (2) and an electrical generator (10). The engine (1) is equipped with a muffler (4) and coupled to an engine bulkhead (6) via a vibration-isolating assembly (5). A battery (7) and a fuel tank (8) are arranged in a fuselage middle part, the fuel tank (8) is coupled to the engine (1) via a fuel line (9) and arranged in a close proximity to a UAV's center of gravity. The electrical generator (10) is arranged on the output shaft (2) between the propeller (3) and the internal combustion engine (1), and it is configured to operate in a starter mode. The vibration-isolating assembly (5) consists of an engine plate (11) that is secured to the engine bulkhead (6) and an engine mount (12) that is coupled to the engine (1). The engine plate (11) and the engine mount (12) are coupled between each other with three supports (13) that are equipped with vibration-dampening assemblies (14) that are coupled to the engine mount (12). An arc-shaped bracket (18) is arranged between the engine plate (11) and the engine mount (12), the bracket (18) is secured to two supports (13) and coupled to the engine mount (12) via an additional vibration-dampening assembly (19). The engine (1) is coupled to the engine mount (12) via spacers (20).

An aircraft, including a wing having an exterior with an exterior surface, a first power supply structure disposed in the wing, an electric motor electrically connected to the first power supply structure, a first heat exchanger disposed in the wing, where the first heat exchanger is disposed adjacent to, and in thermal contact with, at least a portion of the exterior of the wing, where the first heat exchanger is configured to transfer heat to the exterior surface, and a first coolant pump in fluid communication with the first heat exchanger and the first power supply structure.