820results 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.

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 method of determining the capability of a nonlinear aircraft power source includes accessing status values representing a current state of an energy storage system in an aircraft, accessing demand values related to expected power demands on the energy storage system, modeling an ongoing status of the energy storage system using the status values and the demand values to predict when one of the status values will reach a threshold value, and providing an output of a capability of the aircraft based on the status value reaching the threshold value.

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).

A multilevel power converter includes a plurality of switches, a first DC link capacitor, a second DC link capacitor, and one or more processors configured to: generate, for a duty cycle of the multilevel power converter, a pulse width modulated pulse pattern in accordance with a reduced common mode voltage scheme; modify the pulse width modulated pulse pattern to render a modified pulse pattern; and cause the plurality of switches to implement the duty cycle based at least in part on the modified pulse pattern to render a common mode voltage pulse to balance voltages at the first DC link capacitor and the second DC link capacitor.

The present disclosure provides systems and methods for storing, transporting, and using hydrogen. In some embodiments, the method may comprise (a) storing hydrogen fuel in one or more fuel storage modules; (b) transporting the one or more fuel storage modules to a vehicle fueling site, wherein one or more hydrogen fuel compatible vehicles are located at or near the vehicle fueling site; (c) loading the one or more fuel storage modules into the one or more hydrogen fuel compatible vehicles, wherein the one or more fuel storage modules are configured to be releasably coupled to the one or more hydrogen fuel compatible vehicles; and (d) decoupling the one or more fuel storage modules from the one or more hydrogen fuel compatible vehicles after the one or more fuel storage modules are depleted or partially depleted.

A multilevel power converter includes a plurality of switches, a first DC link capacitor, a second DC link capacitor, and one or more processors configured to: generate, for a duty cycle of the multilevel power converter, a pulse width modulated pulse pattern in accordance with a reduced common mode voltage scheme; modify the pulse width modulated pulse pattern to render a modified pulse pattern; and cause the plurality of switches to implement the duty cycle based at least in part on the modified pulse pattern to render a common mode voltage pulse to balance voltages at the first DC link capacitor and the second DC link capacitor.

A turbine wheel may include an inner hub having a central axis, an outer ring concentric with the inner hub, and an intermediate ring concentrically disposed between the inner hub and the outer ring, and a plurality of turbine blades extending between the intermediate ring and the outer ring. The turbine blades may be oriented to be driven by a fluid stream flowing through the turbine wheel to rotate the turbine wheel in a first direction about the central axis. In addition, the turbine wheel may further include a plurality of compressor blades extending between the inner hub and the intermediate ring. Also, the compressor blades may be oriented to propel a fluid in a downstream direction when the turbine wheel is rotated in the first direction.

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.

A propulsion system (100) for an aircraft as disclosed herein may include a nacelle (180), a shaft (110) positioned centrally within a cylindrical passageway (184) of the nacelle (110), a fan (120) coupled to one end (112) of the shaft (110), a turbine (130) coupled to an opposite end (114) of the shaft (110), an electric motor (140) coupled to the shaft (110), a compressor (160) positioned within the cylindrical passageway (184), and a solid oxide fuel cell (150) positioned with a hollow ring-shaped interior (182) of the nacelle (180). The hollow ring-shaped interior (182) may surround and be isolated from the cylindrical passageway (184). The turbine (130) may be configured to provide primary torque to the shaft (110) while the electric motor (140) may be configured to provide additional torque to the shaft (110). The electric motor (140) may be powered an electric output (152) of the solid oxide fuel cell (150) while the turbine (130) may be powered at least in part by output gases (154) from the solid oxide fuel cell (150).

A system for a blended wing body aircraft with a combustion engine is illustrated. The aircraft comprises a blended wing body, at least a fuel store located within the blended wing body and configured to store a fuel, wherein the fuel includes liquid hydrogen, and at least a propulsor 5 configured to propel the blended wing body aircraft. The at least a propulsor comprises a combustion engine configured to burn the fuel from the fuel store and produce mechanical work to power the at least a propulsor.

A method of voltage regulation of at least one bus bar (12) of an electrical system (10), the bar being powered by at least one source, the source being a battery system (14) comprising at least one fuel cell (15) and a compressor (16) associated with the fuel cell, the bar furthermore powering loads including at least one dissipative load, the method comprising the step of, if the compressor is switched into doped mode, controlling a modification of the electrical power supply of at least one of the loads, starting first with the dissipative load(s). FIGURE OF THE ABSTRACT: Fig. 1

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.

A hydrogen fuel system including a fuel delivery assembly, a purge gas source, and a vent. The fuel delivery assembly is configured to receive hydrogen fuel from a hydrogen fuel source and to provide the hydrogen fuel from the hydrogen fuel source to a power generator. The purge gas source is fluidly coupled to the fuel delivery assembly and configured to provide a purge gas to the fuel delivery assembly. The vent is fluidly coupled to the fuel delivery assembly and configured to vent hydrogen fuel from the fuel delivery assembly when the purge gas is provided to the fuel delivery assembly.

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.

The application discloses a system for controlling an auxiliary power unit of an aircraft to perform an automatic start operation in the air, characterized by comprising: an auxiliary power unit (APU) for providing an air source and a power source for the aircraft; an integrated modular avionic system (IMA) for receiving an emergency state signal and an operating state signal associated with the aircraft, and generating an APU auto-start signal according to a preset decision logic based at least in part on the emergency state signal and the operating state signal; and the APU controller is in communication connection with the IMA and is used for controlling the APU to execute an automatic starting operation after receiving the APU automatic starting signal. A method for controlling an auxiliary power plant (APU) of an aircraft to perform an autostart operation in the air, as well as numerous other aspects, is also disclosed.

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.

An example carriage is configured for mounting to an unmanned aerial vehicle. The carriage generally includes a housing assembly configured for mounting to the unmanned aerial vehicle, a movable grip mounted to the housing assembly for movement between a capturing position and a releasing position, a latch device, and a driver. The latch device has a latching state and an unlatching state, is configured to retain the movable grip in the capturing position when the latch device is in the latching state, and is configured to permit movement of the movable grip from the capturing position to the releasing position when in the unlatching state. The driver is operable to transition the latch device from the latching state to the unlatching state.

A method and apparatus for generating electrical power, including providing a generator configured to generate a first voltage, providing a voltage converter configured to operably down-convert the first voltage output to a second voltage output to provide a first amount of power to a set of electrical loads, and changing a power demand for at least a subset of the set of electrical loads, wherein the changed power demand increases the power demanded. To meet the increased power demand, the voltage converter is controlled to operably down-convert the first voltage output to the second voltage output to provide a second amount of power to the set of electrical loads. The second amount of power is greater than the first amount of power. The power-generating capabilities of the generator are not modified between providing the first amount of power and providing the second amount of power.

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.

A system that provides a cooling liquid to a component of an aircraft, the system having: a cooling circuit that includes a fuel cell that receives a first flow and transfers first waste heat to the first flow; an air cycle machine (ACM) that transfers second waste heat to a second flow; a first heat exchanger, fluidly coupled to the cooling circuit downstream of the fuel cell, that thermally couples the first and second flows to superheat the first flow; a turbine, fluidly coupled to the cooling circuit downstream of the first heat exchanger, that extracts energy from the first flow; and a condenser, fluidly coupled to the cooling circuit downstream of the turbine, that condenses the first flow into the cooling liquid, wherein the component is fluidly coupled to the circuit downstream of the condenser.

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.

An aircraft system includes a turbine engine (32, 34) having a compressor (44), a combustor (48) having an inlet (105) and an outlet (107), and a turbine (46) having an inlet portion (110) and an outlet portion (112). An auxiliary power unit (37), APU, is operatively connected to the turbine engine (32, 24). The APU (77) includes a compressor portion (70), a generator (74), and a turbine portion (72). The compressor portion (70) is operatively connected to the turbine portion (72) through the generator (74). A source of cryogenic fluid (62) is operatively connected to the turbine engine (32, 34) and the APU (37). A heat exchange member (84) includes an inlet section (116) operatively connected to the source of cryogenic fluid (62), a first outlet section (120) operatively connected to the turbine engine (32, 34) and a second outlet section (122) operatively connected to the compressor portion (70).

An aircraft (1) is described comprising a motor member (2) which can be operated to power a first load with mechanical power; a second auxiliary load (9); a distribution system (10) electrically connected with the auxiliary load (9) so as to power it with a first electrical power value; a first electric machine (11) which can be operated as an electric generator to convert mechanical power into electrical power or as an electric motor to convert electrical power into mechanical power, and interposed between the motor member (2) and the distribution system (10); the aircraft (1) further comprises a first electrically rechargeable accumulator (20) electrically connected to the electric machine (11) acting as an electric motor to power it with a first 15 electrical power sufficient to cause the motor member (2) to power up, when the motor member (2) is, in use, deactivated.

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;

An energy conversion arrangement (10) for an aircraft (1), an energy system (2) and an aircraft (1) comprising an energy conversion arrangement (10) and / or an energy system (2) are provided, wherein the energy conversion arrangement (10) comprises a fuel conversion device (11), in particular a fuel cell system, for converting at least one fuel to electrical and / or mechanical energy; an expansion device (17) arranged in a flow path (15) of exhausts (E) produced in the fuel conversion device (11) and configured to decompress the exhausts (E); and at least one flow path (14) of air (A) leading to the expansion device (17) for enabling a decompression of the exhausts (E) along with the air (A).