Abstract: Combustion sections and sealing systems for fuel ignition assemblies of gas turbine engine combustion sections are provided. For example, a sealing system comprises a ferrule positioned on an outer surface of a ceramic matrix composite (CMC) combustor liner an aperture defined in the CMC liner; a sleeve positioned within an adapter of the fuel ignition assembly such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall that forms an inner boundary of a cavity defined by the sleeve; and a biasing member positioned within the cavity. The biasing member extends between a bushing and the end wall of the sleeve. The biasing member continuously urges the sleeve into contact with the ferrule to seal the aperture against fluid leakage therethrough. The exemplary sealing system may be part of a fuel ignition assembly of a gas turbine engine combustion section.

Abstract: Propulsion assembly comprising: an inner casing (13); an outer casing (3); an inter-duct casing (15) delimiting a primary duct (12) between the inner casing (13) and an outer wall (14), and a secondary duct (16) between the outer casing (3) and an outer wall (17); a fan capable of generating an air flow (24) circulating from downstream to upstream in the secondary duct (16); the assembly further comprising: at least one duct (27) for bleeding air from said flow (24), this bleed duct (27) comprising an inlet port (28) in the outer wall (17) and an outlet port (29) in the inner wall (14); an outer flap (30) movable between an open position and a closed position of the inlet port (28); an inner flap (31) movable between an open position and a closed position of the outlet port (29).

Abstract: A method for starting and operating a NMHC fueled gas turbine combined cycle power plant includes injecting gaseous NMHC fuel into a gaseous NMHC fuel treatment system, injecting at least one of auxiliary steam, HRSG steam, or HRSG water into the gaseous NMHC fuel treatment system, and mixing the at least one of auxiliary steam, HRSG steam, or HRSG water with the gaseous NMHC fuel in the NMHC fuel treatment system to form a gaseous NMHC fuel mixture. The method further includes injecting the gaseous NMHC fuel mixture into a gaseous NMHC fuel distribution system, and providing the gaseous NMHC fuel mixture through the gaseous NMHC fuel distribution system to a combustor of the NMHC fueled gas turbine.

Abstract: A finely distributed combustion system for a gas turbine engine is provided. A combustor body may extend along a longitudinal axis. A premixer space may be formed within the combustor body to premix air and fuel. The premixer space is in communication with an array of finely distributed perforations arranged in a wall of the combustor body to eject an array of premixed main flamelets throughout a contour of the combustor body between the upstream base of the combustor body and the downstream base of the combustor body. The array of finely distributed perforations—potentially comprising thousands or even hundreds of thousands of perforations spatially distributed on a miniaturized scale—for ejecting the premixed main flamelets is technically advantageous compared to conventional distributed combustion systems, where injection of relatively longer main flames occurs just at a few discrete axial locations.

Abstract: A gas turbine includes a compressor, a turbine, and a combustor disposed downstream from the compressor and upstream from the turbine. The combustor includes an end cover. The combustor also includes a flange. The flange includes an internal fluid passage defined within the flange and the flange is coupled to an internal face of the end cover. A fuel port is integrally joined with the flange. The fuel port extends through the end cover between the flange and an inlet positioned outside of the end cover. The inlet of the fuel port is in fluid communication with the internal fluid passage of the flange.

Abstract: A seal assembly to seal a gas turbine hot gas path flow at an interface of a combustor liner and a downstream component, such as a stage one turbine nozzle, in a gas turbine. The seal assembly including a piston ring seal housing, defining a cavity, and a piston ring disposed within the cavity. The piston ring disposed circumferentially about the combustor liner. The piston ring is responsive to a regulated pressure to secure sealing engagement of the piston ring and outer surface of the combustor liner. The seal assembly includes at least one of one or more sectional through-slots, bumps or channel features to provide for a flow therethrough of a high-pressure (Phigh) bypass airflow exiting a compressor to the cavity. The high-pressure (Phigh) bypass airflow exerting a radial force on the piston ring.

Abstract: A gas turbine engine comprises a compressor section and a turbine section, the compressor section having a last compressor stage. High pressure cooling air is tapped from a location downstream of the last compressor stage and passed through a heat exchanger. Lower pressure air passes across the heat exchanger to cool the high pressure cooling air. A housing surrounds the compressor section and the turbine section and there being a space radially outwardly of the housing, and a mixing chamber received in the space radially outwardly of the housing, the mixing chamber receiving the high pressure cooling air downstream of the heat exchanger, and further receiving air at a temperature higher than a temperature of the high pressure cooling air downstream of the heat exchanger. Mixed air from the mixing chamber is returned into the housing and utilized to cool at least the turbine section.

Abstract: The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

Abstract: A system includes a turbine system that contains a heat recovery steam generator (HRSG) having a flow path that receives an exhaust gas, and having a fluid path that receives a fluid. The fluid path is adjacent to the flow path such that the fluid is heated by the exhaust gas. The HRSG includes an electrical heater that provides heat to the fluid path during a start-up mode of the HRSG, during a shutdown mode of the HRSG, or both.

Abstract: An acoustic damper for a rotary machine includes at least one wall, at least one cooling air inlet, at least one outlet, and at least one cap. The at least one wall extends from a back side of a burner front panel such that the at least one wall partially defines a dampening chamber. The at least one cooling air inlet is defined within the back side of the front panel and is configured to channel a flow of cooling air into the dampening chamber. The at least one outlet is defined within the back side of the front panel and is configured to channel the flow of cooling air out of the dampening chamber. The at least one cap is positioned at least partially over the at least one cooling air inlet and is configured to reduce a velocity of the flow of cooling air within the dampening chamber.

Abstract: A grommet may define a dilution hole in a combustor panel. The grommet may comprise a ridge having a stepped geometry formed about an inner diameter of the grommet, the ridge comprising a passage. The passage may comprise an outlet. The ridge may further comprise a fillet about the inner diameter of the grommet, wherein the outlet is configured to direct a cooling flow circumferentially along the fillet and fill the ridge with the cooling flow.

Abstract: A gas turbine engine and methods of operation include a low pressure electric motor-generator arranged for selective operation in a generator mode to generate electrical power or a drive mode to assist rotation of a low pressure drive shaft of the engine.

Abstract: A vane platform cooling system may comprise a combustor shell and a combustor panel. A cavity may be located between the combustor shell and the combustor panel. A surface of the cavity may be angled toward the combustor shell. An orifice may be formed in a vane platform located aft of the combustor panel. A standoff located in the cavity may direct air toward the vane platform. An aperture may be formed in a surface of the vane platform. A channel formed through the vane platform may connect the orifice and aperture.

Abstract: Disclosed is a tangential on-board injector (TOBI) system that includes an annulus and a plurality of cooling airflow passages disposed about the annulus. Each cooling airflow passage of the plurality of cooling airflow passages includes an inlet opening having a polygonal inlet cross-section, the inlet opening having an inlet cross-sectional area. Each cooling airflow passage of the plurality of cooling airflow passages further includes an outlet opening having an outlet cross-section and an outlet cross-sectional area. The inlet cross-sectional area is greater in magnitude than the outlet cross-sectional area. Also disclosed are additive manufacturing methods for manufacturing the tangential on-board injector system and gas turbine engines that incorporate the tangential on-board injector system.

Abstract: Method for constructing impingement/effusion cooling features in a component of a combustion turbine engine is provided. A pocket 102 may be arranged between an outer wall 104 and an inner wall 106 of the component. A lasing device 108 allows drilling through the component to form an effusion hole 110. The lasing device further allows welding closed an opening 117 formed at outer wall 104 of the component during the drilling with the lasing device through the component. Lasing device 108 further allows drilling through outer wall 104 of the component to form an impingement hole 118 for the impingement/effusion cooling feature. The proposed methodology in a multi-panel arrangement, for example, eliminates a need of having to pre-drill such holes in individual panels prior to the bonding and forming of the component, which overcomes various drawbacks commonly associated with such pre-drilling.

Abstract: This disclosure provides an air tube for a combustor of a gas turbine engine. The air tube includes an inner tube and an outer tube to deliver discharged compressor air into a combustion chamber of the combustor. The air tube can include struts and fins that can improve the cooling performance of the air tube during operation of the gas turbine engine.

Abstract: A nacelle may comprise an inlet cowling comprising an inlet cowling aft edge having an aft edge length; a boat tail cowling comprising a boat tail cowling forward edge having a forward edge length, wherein the boat tail cowling forward edge is disposed adjacent to the inlet cowling aft edge. The forward edge length may be shorter than the aft edge length, forming a step being defined by a portion of the inlet cowling aft edge that is radially outward of the boat tail cowling forward edge. The nacelle may further comprise a transition fairing coupled to the boat tail cowling, wherein the transition fairing comprises a fairing forward edge disposed adjacent to the inlet cowling aft edge and a fairing ramp surface spanning between the inlet cowling aft edge and a boat tail cowling external surface.

Abstract: A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool independently rotatable about a central axis extending through an accessory gear box (AGB). The LP spool has an LP compressor, which is axially positioned between the HP compressor of the HP spool and the AGB. A tower shaft drivingly connects the HP spool to the AGB.

Abstract: A gas turbine engine has an engine case housing a low pressure compressor drivingly connected to a low pressure turbine by a low pressure compressor shaft extending along an engine axis. The low pressure turbine is disposed forward of the low pressure compressor. A low pressure turbine shaft is drivingly connected to the low pressure turbine and extends forwardly of the low pressure turbine. A reduction gear box (RGB) is drivingly connected to the low pressure turbine shaft. The RGB is offset from the engine axis to free an access to low pressure compressor shaft connection. The offset positioning of the RGB allows to provide an access port in an axially forwardly facing surface of the engine case to access the low pressure compressor shaft and more particularly a connection thereof to the low pressure turbine.

Abstract: A turbofan gas turbine engine includes low pressure (LP) motor-generator operable in either a generation mode or a drive mode, an high pressure (HP) motor-generator operable in a either a generation mode or a drive mode, and a power control module for selectively operating the LP and HP motor-generators according to operational conditions of the engine.