Abstract: A variable area turbine nozzle assembly includes a guide vane including an outer centering pin defining a tab. An inner support ring is spaced radially outward from the guide vane and defines an opening and a protrusion. The protrusion is configured to engage with the tab of the outer centering pin. An outer support ring extends circumferentially around the inner support ring and defines an aperture. The outer support ring has a second coefficient of thermal expansion that is greater than or less than the first coefficient of thermal expansion. At least one linkage joins the inner support ring to the outer support ring and is configured to rotate the inner support ring circumferentially about an axial centerline of the variable area turbine nozzle assembly in response to a change in operational temperature of a combustion gas thus causing the guide vane to rotate.

Abstract: A combustor includes a combustion chamber and a liner that bounds at least a portion of the combustion chamber. The liner includes a first side facing the combustion chamber and a second side facing away from the combustion chamber. The liner is formed of a lay-up of ceramic matrix composite (CMC) plies that have a first CMC ply on the first side, a second CMC ply on the second side, and intermediate CMC plies between the first and second CMC plies. At least one of the intermediate CMC plies has a pattern of voids that define cooling channels in the combustor panel. The cooling channels are bound on lateral channel sides by the at least one of the intermediate CMC plies.

Abstract: The disclosure provides a system to heat a portion of a gas turbine (GT) engine of a power generation system in a standby mode. The system includes a first heated fluid manifold and a first conduit coupling a heated fluid supply external to the GT engine to one or more fatigue-prone components within the GT engine. A first plurality of valves is configured to control an amount of heated fluid flow through the first conduit between the heated fluid supply and the one or more fatigue-prone components. A controller is in communication with the first plurality of valves. The controller adjusts an open extent of at least one of the first plurality of valves to control a temperature of the one or more fatigue-prone components. The external heated fluid supply may be an auxiliary system or a second GT engine operating in non-standby mode.

Abstract: A method of blending at least two fuels includes providing at least two fuels to a cyclonic mixer via a fuel supply system, the fuel supply system including a fuel supply circuit for each fuel of the at least two fuels, mixing, via at least one vortex formed in the cyclonic mixer, the at least two fuels to form a fuel mixture, determining, via one or more sensors, a measured interchangeability index of the fuel mixture, comparing the measured interchangeability index to a predetermined interchangeability index, adjusting, via the fuel supply system, one or more parameters of at least one of the at least two fuels based on the comparison between the measured interchangeability index and the predetermined interchangeability index, and providing the fuel mixture to a combustion system.

Abstract: A turbine engine including an engine core, an engine drive shaft, an accessory gear box (AGB), an air turbine starter (ATS), and a bleed air circuit. The engine core has a fan section, a compressor section, a combustion section, and a turbine section in serial flow arrangement. The engine drive shaft operably couples the fan section, the compressor section, and the turbine section. The AGB and the ATS are operably coupled to the engine drive shaft.

Abstract: A turbojet engine, having a primary duct and a secondary duct surrounding the primary duct, a row of guide vanes extending in the secondary duct downstream of a fan, at least one conduit extending in the secondary duct downstream of the guide vanes and housed in a profiled casing, and at least one air sampling scoop located in the region of the casing and equipped with a control valve. Where at least one guide vane located in the vicinity of the casing has a variable geometry zone, the geometry of which can be adjusted during operation.

Abstract: A system includes one or more control units configured to determine one or more fuel consumption models for an aircraft. A method includes determining, by one or more control units, one or more fuel consumption models for an aircraft. A non-transitory computer-readable storage medium comprising executable instructions that, in response to execution, cause one or more control units comprising a processor, to perform operations including determining one or more fuel consumption models for an aircraft.

Abstract: A gas turbine engine combustor panel assembly includes a shell defining an outer periphery of a combustion chamber with an outer skin with a first plurality of holes, a combustor panel including a first end, a second plurality of holes extending through an inner face to an inner void, to an outer face with a third plurality of holes to allow fluid flow from the first plurality of holes to the combustion chamber, a plurality of bolt holes, a second end with an internal combustion chamber section extending axially from a central axis and from a panel wall of the combustor panel. The assembly includes a sheet metal seal plate abutting the plurality of bolt holes and the interface, and a plurality of bolts extending along a central axis and interfacing with the plurality of bolt holes to connect the combustor panel and shell.

Abstract: The support structure of the present disclosure is a piping support structure that supports piping comprising: a header pipe extending in a first direction within a cross section of a duct and having an inlet formed therein through which a high-temperature fluid flows; and a plurality of nozzle pipes integrally provided with the header pipe, extending in a second direction orthogonal to the first direction, and arranged in the first direction. The support structure comprises: ridges protruding from surfaces on at least one side of the outer peripheral surfaces of the nozzle pipes in a third direction orthogonal to the first direction and the second direction; and a support member fixed to the duct and having formed therein a guide groove to which the ridge is fitted and which guides the ridge in the direction of heat extension of the header pipe and the nozzle pipes.

Abstract: A gas turbine engine fuel system includes a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and a power recovery device configured to extract power from exhaust gasses of the burner downstream of the heat-exchanger.

Abstract: An apparatus, system, and method of reducing film cooling scrubbing in a combustor are provided. The apparatus includes a combustor liner that includes a field of a plurality of sparsely-spaced film holes that extend through the combustor liner wherein the field includes a boundary. The combustor liner also includes a patch of relatively more densely-spaced shield holes that extend through the combustor liner at specific predetermined locations within the boundary of the field.

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating assembly and a sensor. The rotating assembly is configured to rotate about an axis. The rotating assembly includes an engine shaft and a phonic wheel threaded onto the engine shaft. The phonic wheel includes a lubricant scoop. The sensor is configured to measure fluctuations in a magnetic field induced by the phonic wheel during rotation of the rotating assembly about the axis.

Abstract: A seal monitoring apparatus is provided. The seal monitoring apparatus includes one or more pressure sensors and a controller in communication with the one or more pressure sensors. The one or more pressure sensors is mounted at a seal of a turbine to measure pressure on a first side of the seal and/or a second side of the seal. The controller is configured to receive pressure data from the one or more pressure sensors and determine a condition of the seal based at least in part on the pressure data. The controller is configured to output a signal upon determining a change in the condition of the seal.

Abstract: An actuation system includes: a generating rotating electrical machine including a first stator including at least one output stator winding; a driving rotating electrical machine including a second rotor, intended to actuate a lubrication pump, and a second stator including at least one input stator winding; an array of switches electrically connected to the at least one output stator winding and to the at least one input stator winding, for electrically connecting them together or disconnecting them from each other depending on the state of the array of switches; and a control device configured to control the array of switches, depending on the state of a control signal, in such a way as to electrically connect or not connect the at least one output stator winding and the at least one input stator winding together.

Abstract: A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, and a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with the fan and with a low pressure turbine of the turbine section. The low pressure turbine includes a rotating drum to which a first airfoil structure is connected and extends radially inward toward the rotary member. A torque frame connects the rotating drum to the rotary member and transfers torque from the first airfoil structure mounted to the rotating drum to the rotary member. The torque frame includes an inner disk mounted to the rotary member, an outer ring and a second airfoil structure formed separately from the outer ring and connected thereto by a releasable connecting structure. The second airfoil structure extends radially inward from the outer ring toward the inner disk.

Abstract: An accessible debris separator for a gas turbine engine is provided. The accessible debris separator including a chamber formed by a entrance wall, an exit wall, an inner wall, and an outer wall; a plurality of inlet openings in the entrance wall; a plurality of outlet openings in the exit wall; and a component within the chamber, the component forcing cooling air with debris particulates entering the chamber via an inlet opening of the plurality of inlet openings to take a circuitous path thereby separating the debris particulates from the cooling air prior to the cooling air exiting the chamber via an outlet opening of the plurality of outlet openings.

Abstract: An aircraft propulsion system includes a core engine including a compressor, combustor, and turbine section. An inlet airflow is compressed communicated to the combustor, mixed with fuel, and ignited to generate an exhaust gas flow that is expanded through the turbine section. A water recovery system includes a series of condensers that include a series of passages for water containing exhaust gas flow. The series of passages are defined circumferentially about the propulsor axis such that recovered water drains downward in a direction of gravity through the series of passages.

Abstract: A shaft assembly having: a first shaft with an inner facing surface having: an inner facing spline extending between spline ends; an inner facing projection, adjacent to one of the spline ends, extending to an inner facing projection surface; and an inner facing projection seat, adjacent to another of the spline ends; and a second shaft having an outer facing surface having: an outer facing spline extending between spline ends; an outer facing projection, adjacent to one of the spline ends of the outer facing spline, extending to an outer projection surface; and an outer facing projection seat, adjacent to another of the ends of the outer facing spline; the first shaft at least partially surrounds the second shaft; the inner facing projection surface is disposed against the outer facing projection seat; and the outer facing projection surface disposed against the inner facing projection seat.

Abstract: An aircraft propulsion system includes a compressor section where an inlet airflow is compressed, a combustor section where the compressed inlet airflow is mixed with fuel and ignited to generate an exhaust gas flow that is communicated through a core flow path, a turbine section through which the exhaust gas flow expands to generate a mechanical power output, the exhaust gas flow is split into at least a first exhaust gas flow and a second exhaust gas flow, and a condenser where water is extracted from the second exhaust gas flow, and a first heat exchanger where at least a portion of the extracted water is utilized for cooling a core flow along the core flow path.