Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner including a dilution slot in fluid communication with the combustion chamber; and one or more guides configured to redirect cooling medium into the dilution slot.

Abstract: A turbomachine including a low-pressure shaft; a stator supporting the low-pressure shaft by bearings including a rear bearing; an electrical machine located at a rear portion of the turbomachine and including a rotor disconnectably driven by the low-pressure shaft and a stator surrounding the rotor and attached to the stator of the turbomachine; the turbomachine includes a journal for supporting the rotor of the electrical machine, the journal including a shaft segment releasably attached to a rear end of the low-pressure shaft by extending the low-pressure shaft, a frame to which the rotor of the electrical machine is attached and which surrounds the shaft segment, and at least one bearing mounted between the frame and the shaft segment; and a sleeve that is axially movable about the low-pressure shaft and/or the shaft segment that is rotationally connected to the low-pressure shaft and the shaft segment.

Abstract: A method and system control the thrust of an aircraft turbomachine having a high bypass ratio by direct action on a variable-pitch system. The variable-pitch system varies the pitch of the vanes of a stator of a low-pressure compressor for the open-loop control of the thrust of the turbomachine. The method also provides closed-loop control of the pitch of the blades of a propeller based on a rotational speed of the propeller.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner further defining a gap; and an intermediate member disposed at least partially within the gap of the liner and in fluid communication with the combustion chamber, wherein the intermediate member is retained at a relatively fixed position with respect to the liner by a support extending between the intermediate member and a case of the gas turbine engine.

Abstract: When the output power of a gas turbine is going to be reduced and when the relationship between the target output power and the rotational frequency of the gas turbine satisfies predetermined conditions, a control device of a power unit system reduces the output power of the gas turbine to the target output power and thereafter reduces the rotational frequency of the gas turbine to the target rotational frequency, whereas when the relationship between the target output power and the rotational frequency of the gas turbine does not satisfy the predetermined conditions, the control device reduces the rotational frequency of the gas turbine to the target rotational frequency and thereafter reduces the output power of the gas turbine to the target output power.

Abstract: In accordance with at least one aspect of this disclosure, a nozzle for a fuel injector includes, a nozzle body defining a central axis from a nozzle inlet to a nozzle outlet. A first fuel circuit is defined in the nozzle body configured to issue a first fuel flow from a first outlet orifice into a combustor. A second fuel circuit is defined in the nozzle body radially outward from the first fuel circuit configured to issue a second fuel flow from a second outlet orifice at a prefilmer surface of the nozzle body. A third fuel circuit is defined in the nozzle body radially outward from the second fuel circuit configured to issue a third fuel flow from a third outlet orifice at the prefilmer surface of the nozzle body. In embodiments, the first fuel circuit, the second fuel circuit, and the third fuel circuit can all be fluidly isolated from one another within the nozzle body.

Abstract: The invention relates generally to methods and apparatus for start-up and control of liquid salt energy storage combined cycle systems.

Abstract: A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. The combustion section can include a combustor liner, a dome assembly coupled to the combustor liner, a fuel nozzle fluidly coupled to the dome assembly, a combustion chamber fluidly coupled to the fuel nozzle, and at least one set of dilution openings located in the dome assembly or combustor liner that fluidly couple to the combustion chamber. A swirler can define at least one passage extending between at least one annular entrance and at least one annular exit, wherein the at least one annular entrance is fluidly coupled to the compressor section. A variable area device is movable relative to the at least one set of dilution openings or at least a portion of the swirler.

Abstract: An engine assembly includes a combustor, a fuel cell stack integrated with the combustor, and a pre-burner system fluidly connected to the fuel cell stack. The fuel cell stack is configured to direct fuel and air exhaust from the fuel cell stack into the combustor. The pre-burner system is configured to control a temperature of an air flow directed into the fuel cell stack. The combustor is configured to combust the fuel and air exhaust from the fuel cell stack into one or more gaseous combustion products that drive a downstream turbine. The engine assembly can further include a catalytic partial oxidation convertor that is fluidly connected to the fuel cell stack. The catalytic partial oxidation convertor is configured to develop a hydrogen rich fuel stream to be directed into the fuel cell stack.

Abstract: An aircraft system includes, among other things, an aircraft and a gas turbine engine coupled to the aircraft. The gas turbine engine includes a propulsor section including a propulsor, a compressor section, a turbine section including a first turbine and a second turbine, and a gear reduction between the propulsor and the second turbine. The second turbine includes a number of turbine blades in each of a plurality of rows of the second turbine. The second turbine blades operating at least some of the time at a rotational speed. The number of blades and the rotational speed being such that the following formula holds true for a majority of the blade rows of the second turbine: 5500 Hz?(number of blades×speed)/60 sec?10000 Hz. The gas turbine engine is rated to produce 15,000 pounds of thrust or more.

Abstract: A three-stream engine is provided. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly encasing the core engine and at least partially defining a core duct. A fan cowl is disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl. The fan cowl at least partially defining an inlet duct and a fan duct. The fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl. A heat exchanger disposed within the fan duct. The heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger.

Abstract: Electric aircraft power plants and associated methods are provided. One power plant includes an emergency power unit (EPU) for providing electric power in the event of a malfunction of a battery pack of an electric aircraft to permit the electric aircraft to make an emergency maneuver. The EPU includes a rocket engine for generating a stream of exhaust fluid using a rocket propellant, a turbine operatively connected to extract energy from the stream of exhaust fluid generated by the rocket engine, and an electric generator operatively connected to be driven by the turbine and to supply electric power to an electric motor propelling the electric aircraft.

Abstract: A mid-frame section of a gas turbine engine having a radial clearance adjustability and a method for adjusting a radial clearance to a rotor in a mid-frame section of a gas turbine engine are presented. The mid-frame section includes a radial clearance adjusting assembly arranged at a compressor exit diffusor. The radial clearance adjusting assembly is configured to adjust the radial clearance to the rotor such that the rotor is concentric with respect to components of the mid-frame section, such as the compressor exit diffusor and shaft cover. The radial clearance adjusting assembly provides radial clearance adjustability in the mid-frame section at multiple locations, such as at forward end and aft end of the compressor exit diffusor. The radial clearance adjusting assembly improves efficiency of the gas turbine and reduces service and operating cost of the gas turbine engine.

Abstract: A gas turbine engine with a closed loop working fluid cycle.

Abstract: A recovery system includes a special organic Rankine cycle (SORC) system, an internal combustion system, a nitrogen generator and a gas-liquid separator. The SORC system includes a first stage heat exchanger (first HE), a plurality of second stage heat exchangers (second HE), a third stage heat exchanger (third HE), a turbo-expander, and at least one condenser. A recovery method, conducted in the recovery system, includes vaporizing a refrigerant, introducing the vaporized refrigerant to the turbo-expander to generate electricity, condensing the vaporized refrigerant in the at least one condenser, collecting nitrogen from the flue gas feed in the nitrogen generator, and collecting water from the flue gas feed in the gas-liquid separator. The refrigerant is vaporized by introducing the refrigerant through the first HE, second HE and third HE sequentially, and the flue gas feed through the third HE, second HE and first HE sequentially.

Abstract: The invention relates to a turbine engine separate flow mixer centered on a longitudinal axis, comprising an exhaust housing, a shroud directly connected to the exhaust housing and intended to mix the flows originating in the turbine engine, said shroud comprising a metal sheet formed by a succession of first and second longitudinal strips distributed circumferentially around the longitudinal axis of the mixer by circumferentially placing the second strips on either side of the first strips, the first and second strips being configured to form the shroud when at rest and grooves in the shroud when operating, the grooves being defined by an alternation of internal lobes and of external lobes.

Abstract: An aircraft propulsion assembly equipped with a thrust-reversing device which includes at least one lateral aperture, passing through the nacelle of the propulsion assembly, configured to be in an open state or a closed state, at least one inflatable barrier configured to be in a deflated state in which the barrier does not protrude into an annular duct channeling a flow of air and an inflated state in which the barrier, substantially rigid or semi-rigid, is deployed and extends across the annular duct to deflect at least a part of the flow of air toward the lateral aperture, and at least one supply system configured to supply the inflatable barrier with fluid. An aircraft is provided which includes at least one such propulsion assembly.

Abstract: A combustor assembly for gas turbine engine, combustor assembly including: combustor wall including interior surface and combustor wall opening, wherein interior surface partially defines combustion chamber, wherein combustor wall opening extends between combustion chamber and its exterior; sealing element disposed partially within combustor wall opening and includes air inlet, air outlet and air passageway, wherein air outlet exits into combustion chamber and delivers flow of air received from exterior of combustion chamber via air inlet and air passageway to combustion chamber; and fuel nozzle coupled to sealing element and configured to deliver fuel into combustion chamber, wherein combustor wall further includes first bearing surface and sealing element further includes second bearing surface, wherein first bearing surface is concave and forms part of first spherical surface such that first and second bearing surfaces are configured to move relative to each other about central point of first spherical sur

Abstract: This summary relates to a thrust reverser for a turbojet engine mounted on an engine nacelle pylon, comprising a fixed part, a movable part, coaxial and mounted in translation with respect to the fixed part and sealingly inserted against the fixed part by blocking an air outlet passage in a closed state whereas the air outlet passage is freed up in the extended state, and a single actuating device including a beam, a slider, and a driving device, configured to drive the movable part in translation with respect to the fixed part, wherein the beam comprises a stop taking the form of a protuberance protruding on a radially outer surface of the beam and configured to have a clearance with respect to said engine nacelle pylon.

Abstract: There is provided a method of operating a gas turbine engine. The gas turbine engine comprises a staged combustor comprising an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method comprises providing fuel to the combustor having a calorific value of at least 43.5 MJ/kg. Also disclosed is a gas turbine engine.