Abstract: A system includes a multi-tube fuel nozzle including a fuel nozzle head that includes an outer wall surrounding a chamber. The outer wall includes a downstream wall portion configured to face a combustion region. The multi-tube fuel nozzle also includes multiple tubes extending through the chamber to the downstream wall portion. Each tube of the multiple tubes includes an upstream portion, a downstream portion, and at least one fuel inlet disposed at the upstream portion, and is configured to receive air and mix the air with fuel from the at least one fuel inlet. The multi-tube fuel nozzle includes a fuel conduit extending through the chamber crosswise to and around the multiple tubes. The fuel conduit includes multiple impingement cooling orifices. A fuel flow path extends through the fuel conduit, through the impingement cooling orifices, through the chamber, and into the at least one fuel inlet of each tube.

Abstract: A turbomachine combustor includes a combustor cap having a cap surface and a wall that extends about the cap surface to define a cap volume, and a plurality of nozzle members that extend from the cap surface. The plurality of nozzle members include a center nozzle member and one ore more outer nozzle members. A combustor dynamics mitigation system is arranged in the combustor cap and includes plurality of divider members that extend from the wall toward the center nozzle member. The plurality of divider members define a plurality of parallel resonator volumes. The combustor dynamics mitigation system also includes a plurality of tubes that extend into corresponding ones of the plurality of parallel resonator volumes.

Abstract: A power plant is provided and includes a gas turbine engine having a combustor in which compressed gas and fuel are mixed and combusted, first and second supply lines respectively coupled to the combustor and respectively configured to supply the compressed gas and the fuel to the combustor and an exhaust gas recirculation (EGR) system to re-circulate exhaust gas produced by the gas turbine engine toward the combustor. The EGR system is coupled to the first and second supply lines and configured to combine first and second portions of the re-circulated exhaust gas with the compressed gas and the fuel at the first and second supply lines, respectively.

Abstract: A forward casing used in a combustor assembly for a turbine engine includes a circumferential sloped surface which reduces the total interior volume within the forward casing. The circumferential sloped surface can be flat or concave shaped.

Abstract: A turbine burner is provided. The turbine burner has a secondary feed unit and a primary feed unit. The primary feed unit has a primary mixing tube and a fuel nozzle that are arranged concentrically around the secondary feed unit. The primary mixing tube and the fuel nozzle have a fluid flow connection. The fuel nozzle has an annular wall that is radially spaced in the axial direction from the secondary feed unit such that a gap height is fainted by the annular wall and the secondary feed unit. The annular wall has an inside wall directed toward the secondary feed unit and having blades with a leading edge on the upstream side. The fuel nozzle has an inlet and the blades have an axial distance from the inlet. The ratio of the distance to the gap height is greater than 1 and less than the gap height.

Abstract: A system includes a gas turbine combustor. The gas turbine combustor includes a combustion liner disposed about a combustion region and a sleeve disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. The gas turbine combustor also includes multiple axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor. The multiple axial injectors are asymmetrically configured to provide an uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.

Abstract: A combustor for a gas turbine includes a plurality of nozzles provided in an array; a baffle plate configured to provide a desired air flow distribution to the array of nozzles; and a casing comprising a plurality of holes in an outer surface. The casing extends from a headend of the combustor to the baffle plate. A method of distributing an air flow in a combustor of a gas turbine includes providing an air flow to the outer surface of the casing; directing the air flow around the baffle plate; and distributing the air flow through the baffle plate to the array of nozzles.

Abstract: A gas turbine system and a method for controlling combustion instability in a combustion section of a gas turbine system are disclosed. The gas turbine system includes a compressor section, a turbine section connected to the compressor section, and a combustor section connected to the compressor section and the turbine section. The combustor section includes a plurality of combustors. The combustor section further includes at least one igniter for igniting a fuel-air mixture within each of the plurality of combustors into a hot gas. The gas turbine system further includes a control system for controlling a velocity of the hot gas in at least one of the plurality of combustors by controlling an operating parameter of the fuel-air mixture.

Abstract: A fuel injector is provided and includes a member defining a flowpath through which a first fluid flows, the flowpath having a cross-section with transverse elongate and short axes, a head defining a plenum storing a supply of a second fluid and a system fluidly coupled to the flowpath and the plenum to inject the second fluid from the plenum and into the flowpath at first and second locations along the elongate axis. The injected second fluid is formed into jets at the first and second locations, the first fluid entrains the jets such that the injected second fluid flows through the flowpath and mixes with the first fluid, and the short axis has a sufficient dimension such that the jets remain spaced from a sidewall of the member.

Abstract: A system for conditioning a working fluid in a combustor includes a primary combustion chamber, a liner circumferentially surrounding the primary combustion chamber, and a primary nozzle in fluid communication with the primary combustion chamber. A secondary combustion chamber located outside of the primary combustion chamber includes a shroud that defines a fluid passage, a secondary nozzle, and means for igniting fuel in the secondary combustion chamber. A method for conditioning a working fluid in a combustor includes flowing the working fluid through a primary combustion chamber and flowing at least a portion of the working fluid through a secondary combustion chamber located outside of the primary combustion chamber. The method further includes flowing a fuel through the secondary combustion chamber, combusting the fuel, and flowing the combustion gases from the secondary combustion chamber into the primary combustion chamber.

Abstract: Disclosed is a fuel nozzle for a combustor of a gas turbine engine includes a nozzle inlet, a combustion area and a swirler disposed between the nozzle inlet and combustion area. The swirler includes a plurality of swirler vanes, each swirler vane capable of creating a pressure difference in fluid flow through the swirler between a pressure side and suction side of the swirler vane. The swirler further includes at least one through airflow hole located in at least one swirler vane of the plurality of swirler vanes. The at least one through airflow hole is capable of utilizing the pressure difference between the pressure side and suction side to promote fluid flow through the at least one airflow hole. Also disclosed is a method for operating a combustor.

Abstract: A combustion nozzle for use in a turbine. The combustion nozzle may include a center body and a cone extending from the center body. The cone may include a number of apertures positioned therein.

Abstract: A gas turbine combustor assembly including a combustor liner and a plurality of fuel nozzle assemblies arranged in an annular array extending within the combustor liner. The fuel nozzle assemblies each include fuel nozzle body integral with a swirler assembly, and the swirler assemblies each a bellmouth structure to turn air radially inwardly for passage into the swirler assemblies. A radially outer removed portion of each of the bellmouth structures define a periphery diameter spaced from an inner surface of the combustor liner, and an interface ring is provided extending between the combustor liner and the removed portions of the bellmouth structures at the periphery diameter.

Abstract: Disclosed herein is a fuel injection system, including: a fuel injection module including a primary fuel injector and one or more secondary fuel injectors disposed around the primary fuel injector; an air supply module for supplying air to the fuel injection module inwardly and outwardly; and a fuel supply module for supplying fuel to the fuel injection module, wherein the fuel injection module serves to generate multistage flames in a burner by forming a rich fuel flame region using the primary fuel injector and forming a lean fuel flame region behind the rich fuel flame region using the secondary fuel injectors through a burning process of gasifying secondary fuel.

Abstract: A fuel injector assembly includes a fuel injector that provides fuel to be mixed with air for ignition in a combustor. The fuel injector assembly includes a fuel shroud and a plurality of spanning members. The spanning members are spaced circumferentially on the fuel shroud between the fuel shroud and a base plate assembly, which supports the fuel injector assembly within the combustor. The spanning members space the fuel shroud from the base plate assembly such that an annular gap is formed therebetween. The gap permits a flow of purge air therethrough to be mixed with compressed air entering the fuel shroud and with fuel provided by the fuel injector. The spanning members comprise elongate members having upstream ends that are circumferentially offset from opposed downstream ends so that the spanning members provide tangential velocity to the purge air flowing through the annular gap.

Abstract: A gas-turbine burner having a plurality of main swirl generators, each having an inlet flow opening formed by the main swirl generator edge, is provided. In order to achieve a uniform flow of combustion air through the main swirl generator, the gas-turbine burner has an inlet-flow guide means with a flow guide surface which runs from one of the inlet-flow openings to an adjacent inlet-flow opening, to which the main swirl generator edges which form the inlet-flow openings are connected, and the flow guide surface widens from there radially upwards. The main swirl generators are central-symmetrically arranged around a pilot burner and the flow guide surface runs radially outside the main swirl generators.

Abstract: A combustor includes a first flow path and a second flow path. The first flow path places a diffuser in indirect fluid communication with a combustor cap assembly by way of an intervening lower combustor annular passageway. The second flow path places the diffuser in direct fluid communication with the combustor cap assembly.

Abstract: A flow balancing valve for a multistage combustor includes a first pressure feedback line, a first burn line, and a first metering port fluidly connected to the first fuel injector. The flow balancing valve further includes a second pressure feedback line, a second burn line, and a second metering port fluidly connected to the second fuel injector. A metering land is located between and defines sizes of the first metering port and the second metering port. An increase in pressure differential between the first pressure feedback line and the second pressure feedback line causes a compensatory movement in the metering land to balance fuel flow for the first fuel injector and the second fuel injector.

Abstract: A multipoint combustion system for a gas turbine engine includes a housing defining a pressure vessel. A master injector is mounted to the housing for injecting fuel along a central axis. A plurality of slave injectors are each disposed outward of the master injector for injecting fuel and air in an ignition plume radially outward of fuel injected through the master injector. The master injector and slave injectors are configured and adapted so the injection plume of the master injector intersects with the ignition plumes of the slave injectors. A primary manifold is included within the pressure vessel for distributing fuel to the slave injectors. An auxiliary manifold is in fluid communication with the auxiliary nozzles of the slave injectors for issuing an auxiliary flow of fuel from the auxiliary nozzles that is separate from the fuel flow of the primary manifold.

Abstract: A combustor includes an end cover having a nozzle. The nozzle has a front end face and a central axis. The nozzle includes a plurality of fuel passages and a plurality of oxidizer passages. The fuel passages are configured for fuel exiting the fuel passage. The fuel passages are positioned to direct fuel in a first direction, where the first direction is angled inwardly towards the center axis. The oxidizer passages are configured for having oxidizer exit the oxidizer passages. The oxidizer passages are positioned to direct oxidizer in a second direction, where the second direction is angled outwardly away from the center axis. The plurality of fuel passages and the plurality of oxidizer passages are positioned in relation to one another such that fuel is in a cross-flow arrangement with oxidizer to create a burning zone in the combustor.

Abstract: A vane assembly has an outer support ring, an inner support ring, an outer liner ring, an inner liner ring, and a circumferential array of vanes. Each vane has a shell extending from an inboard end to an outboard end and at least partially through an associated aperture in the inner liner ring and an associated aperture in the outer liner ring. There is at least one of: an outer compliant member compliantly radially positioning the vane; and an inner compliant member compliantly radially positioning the vane.

Abstract: A fuel injection array for a gas turbine engine includes a plurality of bluff body injectors and a plurality of swirler injectors. A control operates the plurality of bluff body injectors and swirler injectors such that bluff body injectors are utilized without all of the swirler injectors at least at low power operation. The swirler injectors are utilized at higher power operation.

Abstract: A burner for a combustion chamber of a turbine, with an injection device for the introduction of at least one gaseous and/or liquid fuel into the burner is proposed. The injection device has at least one body arranged in the burner with at least two nozzles for introducing the at least one fuel into the burner, the body being configured with a streamlined cross-sectional profile which extends with a longitudinal direction perpendicularly or at an inclination to a main flow direction prevailing in the burner. The carrier air plenum is provided with holes such that carrier air exiting through the holes impinges an inner side of a leading edge portion of the body.

Abstract: A system includes a turbine fuel controller. The turbine fuel controller includes a purge control logic configured to control a purge sequence of mixing a purge gas with a first fuel during a first fuel shutdown, wherein the purge sequence is configured to open a purge valve for the purge gas before fully closing a fuel valve for the first fuel.

Abstract: A unitary distributor is disclosed, comprising at least one main flow passage located within a distributor ring body, the main flow passage having an arcuate portion oriented in a circumferential direction around a distributor axis. Exemplary embodiments of a unitary distributor having an annular ring located at the axially forward end of the distributor ring body are disclosed. A fuel distributor is disclosed, comprising a fuel conduit and a distributor wherein the fuel distributor has a unitary construction.

Abstract: A system that includes a turbine fuel nozzle comprising an air-fuel premixer. The air-fuel premixed includes a swirl vane configured to swirl fuel and air in a downstream direction, wherein the swirl vane comprises an internal coolant path from a downstream end portion in an upstream direction through a substantial length of the swirl vane.

Abstract: The present application provides a premixer for a combustor. The premixer may include a fuel plenum with a number of fuel tubes and a burner tube with a number of air tubes. The fuel tubes extend about the air tubes.

Abstract: A combustor for a turbine is provided. The combustor includes a plurality of fuel nozzles and a combustion zone is aligned with a combustion process associated with each of the fuel nozzles. A combustion liner includes a plurality of turbulator groups, and each of the turbulator groups has or more individual turbulators. Each of the turbulator groups is aligned with a hot streak caused by the combustion zone associated with the fuel nozzle. Each of the turbulator groups are circumferentially spaced from a neighboring turbulator group.

Abstract: The invention relates to a burner, in particular a gas turbine burner, comprises: at least one swirler, the swirler having at least one air inlet opening, at least one air outlet opening positioned downstream to the air inlet opening and at least one swirler air passage extending from the at least one air inlet opening to the at least one air outlet opening which is delimited by swirler air passage walls, the air passage walls comprising downstream wall sections adjoining the at least one air outlet opening; and a fuel injection system which comprises fuel injection openings arranged in at least one swirler air passage wall so as to inject fuel into the swirler air passage; in which at least the downstream section of one air passage wall is corrugated.

Abstract: A rotationally symmetrical wall for an aircraft turbomachine combustion chamber, including primary holes, dilution holes positioned downstream from the primary holes, and a plug installation aperture upstream from the primary holes, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane (D) in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane (D), wherein the other primary holes are distributed equidistantly, at a predefined interval P, and the distance between the two primary holes is greater than the interval P separating two of the other adjacent primary holes.

Abstract: A gas turbine engine combustor is provided and includes an array of fuel nozzles, a combustion casing assembly disposed about the array of fuel nozzles and an end cap assembly disposed within the combustion casing assembly to define with the combustion casing assembly an axis-symmetric annulus through which fluid travels into each of the fuel nozzles, at least one of the combustion casing assembly and the end cap assembly being formed with lobed, three-dimensional contouring.

Abstract: A fuel injector head for a gas turbine engine the head having a pilot injector and a main injector located radially outwardly of the pilot injector. A concentric splitter separates the pilot injector from the main injector and has a toroid chamber which is supplied with air in use to generate a toroidal flow which delays mixing of the pilot and main air flows.

Abstract: The present application and the resultant patent provide a combustor for a turbine engine. The combustor may include a number of fuel nozzles with one or more of the fuel nozzles including a swirler assembly. The swirler assembly may include a number of stages with a number of fueled structures and a number of unfueled structures.

Abstract: A multi-fuel combustion system is provided. The system includes a combustor basket adapted to combust at least two type of fuels. The combustor basket includes a circumferential wall with a plurality of openings. The combustion system further includes a first conduit adapted to provide a first type of fuel directly to the combustor basket and a second conduit adapted to provide a second type of fuel directly to the combustor basket. The combustion system also may include a third conduit adapted to inject at least one of the first type of fuel and the second type of fuel through the openings into the combustor basket.

Abstract: A fuel nozzle assembly for a combustor of a gas turbine including: a nozzle body having a front and an inner tube defining a fuel passage extending through the nozzle body, wherein the front proximate to a combustion section of the combustor; an outer casing around the inner tube, wherein an air passage is defined between the outer casing and the inner tube; a gas conduit arranged in the air passage and having an outlet proximate to the front of the nozzle body, wherein fuel starts flowing through the expandable conduit only after a flashback condition occurs in the combustor, and a premix fuel passage and port discharging fuel to a premix section of the combustor, wherein the gas conduit has an inlet open to the premix fuel passage.

Abstract: An annular expansion ring centered on a main axis and suitable for being mounted on a fuel injector coaxial with the ring is disclosed. The ring presents holes that are distributed around the main axis and open out into its upstream face to enable air to pass towards the zone that is downstream from the ring. The ring includes a conical annular groove that converges downstream, that is open downstream, and that has the holes opening out into the upstream portion thereof. The axis of each of the holes makes an angle relative to the main axis that is strictly greater than the angle between the generator line of the cone defining the annular groove and the main axis, such that the air exiting from the holes impacts against the inner wall of the annular groove, i.e. its wall that is closer to the main axis.

Abstract: The present application and the resultant patent provide a fuel nozzle for mixing a flow of air and a flow of fuel within a combustor. The fuel nozzle may have one or more air passages for the flow of air, one or more fuel pegs for the flow of fuel, and a liner in communication with the air passages and surrounding the fuel pegs. The liner may include a number of ribs thereon so as to promote mixing of the flow of air and the flow of fuel therein.

Abstract: A gas turbine engine combuslor has a trapped dual vortex cavity defined between aft, forward, and bottom walls. Air injection first holes are positioned in the forward wall. Air injection second holes are positioned in the aft walls. Fuel injection holes in the forward wall are located between the bottom wall and a cavity opening located at a top of the cavity. First angled film cooling apertures are disposed through the bottom wall. Second angled film cooling apertures are located in the forward wall between the fuel injection holes and the bottom wall. Third angled film cooling apertures are located in the forward wall between the fuel injection holes and the cavity opening.

Abstract: Combustion systems and the related methods operate to supply a stream of combustion products to a turbine in a gas turbine engine by forming an outer recirculation zone of recirculating combustion products within a combustor and by forming an inner recirculation zone inboard of the outer recirculation zone. The inner and outer recirculation zones are formed by imparting swirl to compressor discharge air passing through an outer air swirler and an inner air swirler radially inboard of the outer air swirler. Fuel is injected from an outer fuel injector into the outer recirculation zone, and fuel is also injected from an inner fuel injector into the inner recirculation zone.

Abstract: A combustor includes two upstream parts of fuel supply system supplying gaseous fuels of two types having different heating values from each other, a three-way fuel transfer valve merging the two upstream parts of fuel supply system with each other, a plurality of gaseous fuel supply subsystems supplying a combustion chamber with the gaseous fuels supplied through the three-way fuel transfer valve and branched, and a plurality of burners injecting, corresponding to each of the gaseous fuel supply subsystems, the gaseous fuel supplied from the gaseous fuel supply subsystem into the combustion chamber.

Abstract: A turbomachine includes a compressor, a combustor operatively connected to the compressor, an end cover mounted to the combustor, and an injection nozzle assembly operatively connected to the combustor. The injection nozzle assembly includes a first end portion that extends to a second end portion, and a plurality of tube elements provided at the second end portion. Each of the plurality of tube elements defining a fluid passage includes a body having a first end section that extends to a second end section. The second end section projects beyond the second end portion of the injection nozzle assembly.

Abstract: A burner for use in a gas turbine engine includes a burner tube having an inlet end and an outlet end; a plurality of air passages extending axially in the burner tube configured to convey air flows from the inlet end to the outlet end; a plurality of fuel passages extending axially along the burner tube and spaced around the plurality of air passage configured to convey fuel from the inlet end to the outlet end; and a radial air swirler provided at the outlet end configured to direct the air flows radially toward the outlet end and impart swirl to the air flows. The radial air swirler includes a plurality of vanes to direct and swirl the air flows and an end plate. The end plate includes a plurality of fuel injection holes to inject the fuel radially into the swirling air flows. A method of mixing air and fuel in a burner of a gas turbine is also provided. The burner includes a burner tube including an inlet end, an outlet end, a plurality of axial air passages, and a plurality of axial fuel passages.

Abstract: A system includes an air flow conditioner configured to mount in an air chamber separated from a combustion chamber of a turbine combustor. The air flow conditioner comprises a perforated annular wall configured to direct an air flow in both an axial direction and a radial direction relative to an axis of the turbine combustor. In addition, the air flow conditioner is configured to uniformly supply the air flow into air inlets of one or more fuel nozzles.

Abstract: A fuel injector tube includes a one piece, unitary, polygonal tube having an inlet end and an outlet end. The fuel injector tube further includes a fuel passage extending from the inlet end to the outlet end along a longitudinal axis of the polygonal tube, a plurality of air passages extending from the inlet end to the outlet end and surrounding the fuel passage, and a plurality of fuel holes. Each fuel hole connects an air passage with the fuel passage. The inlet end of the polygonal tube is formed into a fuel tube. A fuel injector includes a plurality of fuel injector tubes and a plate. The plurality of fuel tubes are connected to the plate adjacent the inlet ends of the plurality of fuel injector tubes.

Abstract: An internal combustion engine in which the power output is controlled by modulating at least one of the compression ratio, expansion ratio, ratio of expansion rate to compression rate, air to fuel ratio, and steam to air ratio. Continuous isobaric catalytic combustion followed by isothermal expansion and the use of separate compressor and expander devices are used. Control dynamically maximizes fuel efficiency for the given power demand conditions. Power output is controlled by modulating flame temperature and/or pressure instead of by throttling. Lean combustion, high compression ratio, exhaust heat recuperation, and high power density and fuel economy are provided. External cooling is minimized or eliminated. Insulation of the engine effectively reduces energy losses to friction. Interchangeable use of gasoline, hydrogen and ammonia at high fuel efficiency is made possible for transitional periods of fuel availabilities. An injector suitable for isothermal expansion is provided.

Abstract: A fuel spraying apparatus of a gas turbine engine includes a first fuel spraying section configured to spray fuel for diffusion combustion and a second fuel spraying section configured to spray fuel for premixed combustion. The second fuel spraying section is disposed so as to surround the first fuel spraying section. The apparatus further includes an air curtain generating section disposed between the first fuel spraying section and the second fuel spraying section. The air curtain generating section is configured to generate an air curtain which defines an outer edge of a diffusion combustion area in a combustion chamber.

Abstract: A gas turbine engine has in flow series: a compressor section, a combustor, and a turbine section. The engine further has a steam generator, and a water and steam delivery system. During an engine start-up period, the delivery system delivers liquid water to the combustor and mixes the delivered water in its liquid state with working gas in the combustor. After the start-up period, the delivery system accepts steam from the steam generator, delivers the steam to the combustor and mixes the delivered steam in its gaseous state with the working gas in the combustor.

Abstract: According to various embodiments, a system includes a staggered multi-nozzle assembly. The staggered multi-nozzle assembly includes a first fuel nozzle having a first axis and a first flow path extending to a first downstream end portion, wherein the first fuel nozzle has a first non-circular perimeter at the first downstream end portion. The staggered multi-nozzle assembly also includes a second fuel nozzle having a second axis and a second flow path extending to a second downstream end portion, wherein the first and second downstream end portions are axially offset from one another relative to the first and second axes. The staggered multi-nozzle assembly further includes a cap member disposed circumferentially about at least the first and second fuel nozzles to assemble the staggered multi-nozzle assembly.

Abstract: An end cover assembly is provided in combination with a gas turbine combustor for capping a back end of a combustion chamber. The end cover assembly includes an end cover plate securable to the back end of the combustion chamber, and a monolith block secured to the end cover plate. The end cover plate is constructed without welded gas passage plates and without brazed flow passage inserts and includes fuel channels therethrough, where the fuel channels are sized and positioned to interact with fuel nozzle inlet ports. The monolith block includes internal fuel gas passages positioned relative to the fuel channels in the end cover plate to commute fuel gas to the fuel channels of the end cover plate.

Abstract: A method of assembling a combustor assembly for use in a turbine engine is provided. A combustor liner that defines a combustion chamber therein is provided. The liner includes a forward end and an aft end, wherein the aft end includes at least one channel extending therethrough. The channel is aligned obliquely with respect to a centerline extending through the aft end. A plurality of fuel nozzles are coupled to the forward end such that the fuel nozzles extend through the forward end. An annular sleeve that includes at least one opening extending radially therethrough is coupled to the aft end, wherein the sleeve substantially circumscribes the aft end such that a cavity is defined therebetween and such that a fluid channeled through the at least one opening impinges against a surface of the aft end prior to the fluid being channeled into the combustion chamber.