Abstract: A gas turbine engine system includes a gas turbine engine including a compressor, combustor including a plurality of late lean fuel injectors supplied with secondary fuel; gas turbine, and wash system configured to be attached and in fluid communication with the late lean fuel injectors. The wash system includes a water source including water; first fluid source including a first fluid providing vanadium ash and vanadium deposit mitigation and removal from internal gas turbine components; a mixing chamber in communication with the water source and first fluid source; a water pump to pump the water to the mixing chamber; a first fluid pump the first fluid to the mixing chamber; a fluid line in fluid communication with the mixing chamber and late lean fuel injectors so fluid from the mixing chamber is injected into the combustor at the late lean fuel injectors while the gas turbine engine is on-line.

Abstract: The present disclosure provides methods, assemblies, and systems for power production that can allow for increased efficiency and lower cost components arising from the control, reduction, or elimination of turbine blade mechanical erosion by particulates or chemical erosion by gases in a combustion product flow. The methods, assemblies, and systems can include the use of turbine blades that operate with a blade velocity that is significantly reduced in relation to conventional turbines used in typical power production systems. The methods and systems also can make use of a recycled circulating fluid for transpiration protection of the turbine and/or other components. Further, recycled circulating fluid may be employed to provide cleaning materials to the turbine.

Abstract: A method of operating a multi-engine system of an aircraft having first and second engines includes accumulating compressed air in a pressure vessel external to the engines, and operating the first and second engines asymmetrically, by controlling the first engine to operate in an active operating condition providing sufficient power and/or rotor speed for demands of the aircraft, and controlling the second engine to operate in a standby operating condition wherein the second engine produces less power output than the first engine. In response to a power demand request, the second engine is accelerated out of the standby operating condition by introducing therein compressed air from the pressure vessel at a location upstream of a combustor of the second engine.

Abstract: A power generation installation in which exhaust gas from a gas turbine is fed to a thermal energy accumulator, wherein energy in the thermal energy accumulator can be employed for various purposes, a method for operating such an installation, and a method for the modification of existing installations. The thermal energy accumulator has sufficient capacity to permit the operation of a steam turbine in isolation for the storage of thermal energy from exhaust gas in the thermal energy accumulator.

Abstract: The present disclosure relates to systems and methods that are useful for controlling one or more aspects of a power production plant. More particularly, the disclosure relates to power production plants and methods of carrying out a power production method utilizing different fuel chemistries. Combustion of the different fuel mixtures can be controlled so that a defined set of combustion characteristics remains substantially constant across a range of different fuel chemistries.

Abstract: An aircraft power plant has a hybrid propulsion system having an electric motor, an output shaft drivingly connectable to a thrust generator, a combustion engine, a compressor, and a planetary gear train having an Auxiliary Power Unit (APU) mode in which the electric motor is in driving engagement with the compressor via the planetary gear train while the combustion engine is disengaged from the output shaft, and a propulsion mode in which the combustion engine and the electric motor are in driving engagement with the output shaft via the planetary gear train.

Abstract: A bleed system including control circuitry and a variable jet pump. The control circuitry is configured to receive a signal indicative of a fluid parameter in the bleed system and cause the jet pump to alter a mixing ratio of a higher pressure gas and a lower pressure gas based on the signal. The jet pump is configured to combine the lower pressure gas and the higher pressure gas in the mixing ratio to generate a mixed gas. The jet pump is configured to supply the mixed gas to one or more gas loads in the bleed system. In examples, the control circuitry is configured to establish a system setpoint for the fluid parameter based on an operating status of the one or more gas loads.

Abstract: A propulsion system for an aircraft having a nacelle and a fuselage is provided. The nacelle has a gas flow path and a nacelle interior region. The system includes a compressor section, an intermittent IC engine, a turbine section, a fan, and an IC cooling system. A first electric motor powers the compressor section. The compressor section produces a flow of elevated pressure compressor air. The intermittent IC engine selectively intakes the compressor air flow and produces an exhaust gas flow. The turbine section powered by exhaust gas in turn powers a first electric generator. The fan is driven by a second electric motor. The IC engine cooling system has a heat exchanger disposed within the gas flow path, coolant, coolant piping, and a pump. The heat exchanger is disposed in the nacelle.

Abstract: An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

Abstract: The gas turbine intake can have a swirl housing assembly with a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending around the central axis and between the tangential inlet and the annular outlet, the swirl housing assembly having a proximal portion defining a first portion of the swirl path, a distal portion defining a second portion of the swirl path, vanes located in the swirl housing assembly, the vanes circumferentially interspaced from one another relative the central axis and extending between the proximal portion and the distal portion, the proximal portion fastened to the distal portion via a plurality of fasteners, a gasket sandwiched between the proximal portion and the distal portion by the plurality of fasteners, the gasket extending in a radial plane relative the central axis.

Abstract: The present disclosure provides an ejector nozzle and an ejector including the same. The ejector nozzle includes a first tube having a first flow path into which a fluid is introduced, and a second tube provided outside the first tube and having an inner diameter larger than an inner diameter of the first tube, the second tube defining a second flow path between the first tube and the second tube, in which the first tube further includes a communication port that penetrates the first tube to allow the first flow path to communicate with the second flow path and is openably and closably provided, and in which when the communication port is opened, a part of the fluid flowing in the first flow path is allowed to flow along the second flow path.

Abstract: A fuel delivery system for a combustor of a gas turbine engine including a primary fuel tank configured to store a primary fuel, a secondary fuel tank configured to store a secondary fuel, a swirler configured to produce a main flame within a combustion chamber of the combustor, and a fuel nozzle configured to produce a pilot flame within the combustion chamber of the combustor. The fuel nozzle includes a nozzle outlet that is located proximate to an end of the swirler or at the end of the swirler, the end of the swirler being located at an inlet of the combustor. The fuel delivery system also includes a primary fuel line fluidly connecting the primary fuel tank to the fuel nozzle and a secondary fuel line fluidly connecting the secondary fuel tank to the swirler.

Abstract: An air management system with a set of compressed air sources for selectively supplying pressurized air to air consumer equipment according to an aircraft operation condition. Low pressure air, high pressure air, or a combination thereof may perform such supplying of compressed air depending on the aircraft operation condition.

Abstract: Power and cooling systems including a drive system, a power generation unit, and a cooled fluid generation unit. A primary working fluid that is expanded within a turbine of the drive system and compressed within compressors in a closed-loop cycle. The power generation unit includes a generator and a heat source configured to heat the primary working fluid prior to injection into the turbine. T cooled fluid generation unit includes an ejector downstream of the compressors and a separator arranged downstream of the ejector and configured to separate liquid and gaseous portions of the primary working fluid. The gaseous portion is directed to the compressors and the liquid portion is directed to an evaporator heat exchanger to generate cooled fluid.

Abstract: A hybrid electric propulsor equipped with a hydraulic coupling is provided. In one aspect, a propulsor includes a gas turbine engine and an electric machine operatively coupled thereto. The gas turbine engine includes a fan spool and a core spool. The electric machine is operatively coupled with the fan spool. A hydraulic coupling defines a sealed volume in which hydraulic transmission fluid is provided. The hydraulic coupling encases at least a portion of the fan spool and at least a portion of the core spool. The hydraulic coupling hydraulically couples the fan spool and the core spool and allows for power transmission between the fan spool and the core spool.

Abstract: The invention relates to a firewall (112) for a turbomachine (100) auxiliary arm (106). This firewall (112), which has one or more through openings (180) in a thickness direction (Z) of the firewall (112), comprises a fixed portion (120) and a first removable portion (130) extending, in a width direction (Y), orthogonal to said thickness direction (Z), from at least one adjacent opening among the through openings (180), to a first outer edge (190) of the firewall (112). To open this firewall (112) in order to allow access to one or more ducts and/or electrical cables (111) received in the through openings (180), the first removable portion (130) can be dismantled.

Abstract: A box-shaped air intake silencer with vertical baffles for a gas turbine system. The box-shaped silencer has a silencer housing having opposing ends with a first end closed and a second end opened to permit passage of air. Opposing sidewalls extend between the first end and the second end with each of the sidewalls operative to receive a perpendicular flow of intake air. A silencer having vertically, spaced baffles are disposed within the housing. A first end of the baffles face the first end of the housing and a second end of the baffles faces the second end of the housing. Each of the baffles extend horizontally between the first end and the second end of the housing. The baffles receive the perpendicular flow of intake air proximate the first end of the housing and direct the intake air towards the second end of the housing for passage therethrough.

Abstract: A gas turbine engine has a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction, and includes a lubricant pump to supply an air/oil mixture to an inlet of a deaerator. The deaerator includes a separator for separating oil and air, delivering separated air to an air outlet, and delivering separated oil back into an oil tank. The separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank. Air within the oil tank moves outwardly through an air exit into the deaerator. A method of designing a gas turbine engine is also disclosed.

Abstract: An aircraft engine, has: a low-pressure compressor and a high-pressure compressor located downstream of the low-pressure compressor; a gaspath valve upstream of the high-pressure compressor, the gaspath valve movable between an open configuration and a closed configuration; and a bypass flow path having in flow series a bypass inlet, a bypass valve, and a bypass outlet, the bypass inlet fluidly communicating with the gaspath upstream of at least one stage of the low-pressure compressor, the bypass valve having an open configuration in which the bypass valve allows a bypass flow and a closed configuration in which the bypass valve blocks the bypass flow, the bypass outlet fluidly communicating with the bypass inlet via the bypass valve and with the gaspath at a location in the gaspath fluidly downstream of the gaspath valve, downstream of the low-pressure compressor, and upstream of the high-pressure compressor.

Abstract: Turbines for an aircraft include a supply device of compressed air configured to supply compressed air to the aircraft, a cooling system of the compressed air supplied to the aircraft, having a scoop configured to collect cooling air in a flow duct of a secondary flow, and a management system configured to be supplied with the cooling air collected from the flow duct and configured to control radial clearances between a turbine casing and turbine rotor vanes tips. The management system is supplied with cooling air by the scoop of the cooling system.