Abstract: An aircraft propulsion system is provided that includes a turbine engine and an intercooler (IC). The turbine engine includes a source of non-hydrocarbon fuel, and a low pressure compressor having an LPC air inlet and outlet, and a high pressure compressor having an HPC air inlet and outlet. The intercooler has an IC air inlet, an IC air outlet, a U-shaped air passage, and crossflow tubes. The IC air inlet is in fluid communication with the LPC air outlet. The IC air outlet is in fluid communication with the HPC air inlet. The U-shaped air passage is in fluid communication with the IC air inlet and outlet. The crossflow tubes extend through the U-shaped air passage. The crossflow tubes are in fluid communication with the source of non-hydrocarbon fuel, and are configured to contain a flow of the non-hydrocarbon fuel therethrough.

Abstract: A turbine engine with an axis is provided. This turbine engine includes a fan section, a turbine engine core, a bypass flowpath, an engine housing, an evaporator, a condenser and a core flowpath. The turbine engine core is configured to power the fan section. The turbine engine core includes a core compressor section, a core combustor section and a core turbine section. The bypass flowpath is fluidly coupled with and downstream of the fan section. The engine housing includes a cavity radially outboard of and axially overlapping the fan section and/or the bypass flowpath. The evaporator module is within the cavity. The condenser module is within the cavity. The core flowpath extends sequentially through the core compressor section, the core combustor section, the core turbine section, the evaporator module and the condenser module.

Abstract: An assembly for rotational equipment includes a shaft, a first rolling-element bearing, a second rolling-element bearing, and a mid-shaft foil bearing. The shaft extends along a rotational axis between and to a first axial end and a second axial end. The shaft is configured for rotation about the rotational axis. The first rolling-element bearing is disposed at the first axial end. The second rolling-element bearing is disposed at the second axial end. The mid-shaft foil bearing is disposed axially between the first rolling-element bearing and the second rolling-element bearing.

Abstract: An aircraft engine comprises an exhaust duct receiving an engine gas flow at a first pressure during an operating condition and at a second, lower pressure at shutdown. A heat exchanger duct connected to the exhaust duct has a heat exchanger disposed therein. A cover in the heat exchanger duct includes a closure movable between open and closed positions to fluidly connect or disconnect the heat exchanger duct to the exhaust duct. A biasing member disposed in the exhaust duct, exposed to the gas flow, and operatively connected to the closure biases the closure toward the closed position with a biasing force. First and second forces opposing the biasing force are generated on the biasing member by the gas flow at the first and second pressures, the first being greater than the second. The biasing force is greater than the second force and less than the first force.

Abstract: The subject matter of this specification can be embodied in, among other things, a system that includes a flexible structure, a power inverter, a first electromechanical device configured to urge movement of a first portion of the flexible structure based on power received from the power inverter, a second electromechanical device configured to urge movement of a second portion of the flexible structure based on power received from the power inverter, a feedback sensor configured to provide a feedback signal indicative of alignment between the first portion and the second portion, and a controller configured to control operation of at least one of the first electromechanical device and the second electromechanical device based on the feedback signal.

Abstract: A vertical type once-through heat recovery steam generator (HRSG) capable of improving operational stability during rapid startup and lifetime of a steam turbine while reducing environmental pollution caused by emissions, and a combined power generation system including the HRSG are provided. The HRSG includes a medium-pressure section including a medium-pressure side desuperheater installed at the rear of a medium-pressure superheater to lower the temperature of the steam supplied to a medium-pressure steam turbine, and a high-pressure section including a high-pressure side desuperheater installed at the rear of a high-pressure superheater to lower the temperature of the steam supplied to a high-pressure steam turbine.

Abstract: Systems and methods for operating systems are provided. For example, a system comprises a heat source for providing a flow of a hot fluid and a fuel flowpath for a flow of a fuel. The fuel flowpath includes a fuel accumulator and a heat exchanger for heat transfer between the hot fluid and fuel. The heat exchanger includes a hot fluid inlet for receipt of the hot fluid at an inlet temperature and a fuel inlet for receipt of the fuel at an inlet temperature. The hot fluid inlet temperature is greater than the fuel inlet temperature such that the fuel is heated through heat transfer with the hot fluid in the heat exchanger. The fuel accumulator accumulates at least a portion of the heated fuel. An exemplary system is selectively operated to heat and circulate the fuel through the fuel flowpath for consumption and/or accumulation in the fuel accumulator.

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with a gaseous fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. A fuel system supplies a fuel to the combustor through a fuel flow path, a first heat exchanger thermally communicates a first heat load into a cooling flow, a turboexpander where a heated cooling flow from the first heat exchanger is expanded to generate shaft power and cooled to provide a cooled cooling flow, and a second heat exchanger thermally communicates a second heat load to the cooled cooling flow that is communicated from the turboexpander, cooling flow from the second heat exchanger is communicated to the combustor section.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a fuel injector assembly, and the fuel injector assembly includes an inner air swirler and a fuel swirler. The inner air swirler is configured to swirl inner air in a direction about an axis to provide inner swirled air. The fuel injector assembly is configured to inject the inner swirled air as an inner air flow along an inner air flow trajectory parallel with the axis. The fuel swirler is configured to swirl fuel in the direction about the axis to provide swirled fuel. The fuel injector assembly is configured to inject the swirled fuel as an annular fuel flow along a fuel flow trajectory parallel with the axis. The annular fuel flow is adjacent and circumscribes the inner air flow.

Abstract: Turnbuckle linkages are disclosed herein. A turnbuckle linkage includes a hollow rod including an inner portion, a first outer portion, and a second outer portion between the first outer portion and the inner portion, the inner portion including a first thickness, the second outer portion including a second thickness less than the first thickness, and the first outer portion including a conical thickness that varies from the second thickness to a third thickness less than the second thickness, the first outer portion and the second outer portion including first threads, a threadless rod to be partially positioned in the hollow rod, and a conical nut including second threads to engage the first threads, the conical nut to couple the hollow rod and the threadless rod.

Abstract: A gas turbine engine having a pre-swirler adjustability is presented. The pre-swirler includes a pre-swirler insert installed in a component enclosed by a cover. The component includes an inner compressor exit diffusor enclosed by an outer casing or a shaft cover enclosed by the inner compressor exit diffusor. The pre-swirler is adjustable by replacing the pre-swirler insert. An access port including an access window is arranged on the cover. The access port gives access to the pre-swirler insert for replacement through the access window. The access window includes a manhole or combustor assembly installation hole on the outer casing, or a cutout on the inner compressor exit diffusor. The access port allows adjusting the pre-swirler by replacing the pre-swirler insert installed in the component without lifting the cover enclosing the component.

Abstract: A method and system for identifying combustion dynamics in a combustion chamber, includes an optical sensor that receives energy from a flame within the combustion chamber. A processor is configured to receive a first signal from the sensor indicative of energy at a first wavelength and a second signal indicative of energy at a second wavelength. The processor can generate a data set of combustion quality indicators from the first signal and the second signal. The processor can convert the data set of combustion quality indicators in a time domain to a combustion quality spectrum in a frequency domain. The processor can analyze the combustion quality spectrum to determine anomalies, wherein the anomalies indicate at least one frequency where combustion dynamics occur in the combustion chamber and output a signal indicative of the at least one frequency where combustion dynamics occur.

Abstract: An aircraft turbomachine includes a lubricating circuit notably including a reservoir of lubricating fluid, a pump drawing lubricating fluid from the reservoir to inject it toward several components of the turbomachine, and an intake pipe connecting the reservoir to an intake orifice of the pump, the turbomachine further including a pneumatic starting circuit including a starting tube through which there circulates a flow of compressed air intended to supply a pneumatic starter of the turbomachine, wherein it includes a pneumatic line extending from the starting tube as far as the reservoir so as to supply the reservoir with compressed air so as to cause the lubricating fluid to circulate through the intake pipe towards the pump.

Abstract: A turbine engine that includes an engine core, an inner cowl, an outer cowl and an accessory gearbox. The engine core includes at least a compressor section, a combustion section, and a turbine section in axial flow arrangement. The accessory gearbox is operably coupled to the engine core and includes a first portion and a second portion.

Abstract: Dissimilarly shaped aircraft nozzles with tandem mixing devices, and associated systems and methods are disclosed. An ejector nozzle in a representative embodiment includes a nozzle duct having a nozzle flow axis, a first axial position and a second axial position. The nozzle duct has a first cross-sectional shape at the first axial position, and a second cross-sectional shape at the second axial position, with the second shape being geometrically non-similar to the first shape. The nozzle further includes a fan flow duct portion and a core flow duct portion, both upstream of the first axial position. An ejector duct is positioned in fluid communication with the nozzle duct, and has at least one portion with a cross-sectional shape geometrically similar to the second cross-sectional shape.

Abstract: A turbine engine is provided that includes a fan section, a turbine engine core, an evaporator, a condenser, a core flowpath and a bypass flowpath. The turbine engine core is configured to power the fan section. The turbine engine core includes a core compressor section, a core combustor section and a core turbine section. The evaporator is arranged axially along an axis between the fan section and the turbine engine core. The condenser is arranged axially along the axis between the fan section and the turbine engine core. The core flowpath extends sequentially through the core compressor section, the core combustor section, the core turbine section, the evaporator and the condenser. The bypass flowpath is fluidly coupled with and downstream of the fan section. The bypass flowpath is radially outboard of and extends axially along the evaporator and the condenser.

Abstract: A combustor for a turbomachine engine includes a combustion chamber, a liner forming a boundary of the combustion chamber, and multiple dilution holes through the liner to permit airflow into the combustion chamber. The dilution holes include converging dilution holes and diverging dilution holes, the converging dilution holes having a converging cross-sectional profile and the diverging dilution holes having a diverging cross-sectional profile.

Abstract: A system is provided for a powerplant. This powerplant system includes a core compressor section, a core combustor section, a core turbine section and a core flowpath extending longitudinally through the core compressor section, the core combustor section and the core turbine section between an inlet into the core flowpath and an exhaust from the core flowpath. The powerplant system also includes a first rotating assembly, a second rotating assembly and a clutch. The first rotating assembly includes a first compressor rotor and a first turbine rotor. The first compressor rotor is arranged in the core compressor section. The first turbine rotor is arranged in the core turbine section. The second rotating assembly includes a second compressor rotor and a second turbine rotor. The second turbine rotor is arranged in the core turbine section. The clutch is configured to selectively couple the first rotating assembly to the second rotating assembly.

Abstract: An assembly is provided for a gas turbine engine. This engine assembly includes an exhaust nozzle extending circumferentially about an axial centerline. The exhaust nozzle extends axially along the axial centerline to a trailing edge. The exhaust nozzle extends radially between an exterior inner surface and an exterior outer surface. The exhaust nozzle includes an inner skin, an outer skin and an internal cavity. The inner skin forms the exterior inner surface and includes a plurality of perforations. The internal cavity is disposed radially between the inner skin and the outer skin. The internal cavity is fluidly coupled with the perforations.

Abstract: A propulsion system for an aircraft includes a propulsor rotor, an engine core, a housing, a flowpath and a heat exchange system. The engine core is configured to power rotation of the propulsor rotor. The housing includes an inner structure and an outer structure. The outer structure is radially outboard of and axially overlaps the inner structure. The flowpath is downstream of the propulsor rotor and radially between the inner structure and the outer structure. The heat exchange system includes a duct, an exterior orifice, an interior orifice, a flow regulator and a heat exchanger disposed within the duct. The exterior orifice is formed by the outer structure and disposed along an exterior of the housing. The interior orifice is formed by the outer structure and disposed along the flowpath. The flow regulator is configured to selectively fluidly couple the exterior orifice and the interior orifice to the duct.