Abstract: Fuel injectors for gas turbine engines include a housing and a tube defining a portion of a first fluid passage therein and a second fluid passage is defined between an exterior surface of the tube and an interior surface of the housing. An inner airflow tube is arranged having an inflow vane assembly and a central air passage. A portion of the first fluid passage extends axially at a position radially outward from the inner airflow tube, and the third fluid passage extends axially at a position radially outward from the first fluid passage. A nozzle outlet is configured to receive first, second, and thirds fluids from the respective fluid passages to cause mixing thereof. The inflow vane assembly comprises a number of vanes, with each vane angled relative to a nozzle axis at an angle between 20° and 40°.

Abstract: A combustor for a gas turbine engine comprises an inner annular liner wall and an outer annular liner wall cooperating to form a combustion chamber of the combustor. A first dome wall has a circumferential array of first fuel injection bores. A second dome has a circumferential array of second fuel injection bores. An intermediate wall extends between the first dome wall and the second dome wall. A first combustion stage is defined by the inner liner wall forward end, the first dome wall and the intermediate wall. A second combustion stage is defined at least by the outer liner wall forward end, the second dome wall and the intermediate wall, the first combustion stage communicating with the first fuel injection bores, the second combustion stage communicating with the second fuel injection bores.

Abstract: A nozzle for a combustor burning a hydrogen-containing fuel is provided. The nozzle includes a first tube through which the fuel flows and having a fuel injection hole on a front side to inject the fuel therethrough, a second tube surrounding the first tube and having a premixing injection hole through which the fuel and air are mixed and discharged, and a third tube surrounding the second tube and through which the fuel flows, wherein the second tube includes a plurality of fine injection holes to inject the fuel from the third tube to form fine flames, and the fine injection holes are spaced apart from each other in a circumferential direction of the second tube.

Abstract: The upstream-side wall portion 54 includes, in the circumferential direction thereof, a first region 31 where air inlets 30 are formed at a lower density, and a second region 32 which is disposed at a position offset from the first region 31 in the circumferential direction, and in which the air inlets 30 are formed at a higher density than in the first region 31.

Abstract: The present invention provides methods and system for power generation using a high efficiency combustor in combination with a CO2 circulating fluid. The methods and systems advantageously can make use of a low pressure ratio power turbine and an economizer heat exchanger in specific embodiments. Additional low grade heat from an external source can be used to provide part of an amount of heat needed for heating the recycle CO2 circulating fluid. Fuel derived CO2 can be captured and delivered at pipeline pressure. Other impurities can be captured.

Abstract: A combustor assembly for a gas turbine engine includes a dome defining a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end is received within the slot of the dome. In one exemplary aspect, the combustor assembly includes features that warm the forward end of the liner during transient operation of the engine. Furthermore, the combustor assembly includes features that reduce the thermal gradient between the forward end and the other portions of the liner. In this way, improved durability of the liner may be achieved.

Abstract: A fuel injector system includes an outer support and an inner support, with a feed arm extending radially between the inner support and the outer support. A plurality of outlet openings extend in an axial direction from the feed arm for feeding respective injection nozzles. The outer support and feed arm define a plurality of fuel passages therethrough to convey fluid from an external source through the outer support and feed arm to the outlet openings. A heat shield extends around the feed arm from the outer support to the inner support. The heat shield is spaced apart from the feed arm with an insulative gap therebetween.

Abstract: A combustor provided with: a nozzle having formed therein an air jet section that causes air to be jetted from a nozzle section; a cylindrical part covering the nozzle from the outer peripheral side thereof and forming an air flow path between the cylindrical part and the nozzle; and a pressure loss section provided to the air flow path. The pressure loss section causes a loss of pressure in the air that flows through the air flow path. The nozzle is provided with: at least one air inlet section that takes in air from an outer peripheral surface that is an upstream side from the pressure loss section; and a flow path-forming section forming a discharge air flow path that guides air that is taken in from the at least one air inlet section to the air jet section.

Abstract: A premixer assembly for a combustor includes: at least one ring of premixers, each premixer having a central axis, an annular peripheral wall surrounding a centerbody, and at least one swirler disposed between the centerbody and the peripheral wall, wherein the peripheral wall defines an inlet area of the premixer; and a lip extending forward along the central axis from the peripheral wall, the lip extending at an oblique angle to the axis of symmetry.

Abstract: A combustor includes a combustion chamber; a nozzle supplied with fuel and air to produce a mixed gas of the fuel and the air and configured to inject the mixed gas into the combustion chamber; a fuel supply section coupled to the nozzle casing and configured to supply the fuel to the nozzle; a fuel conditioner that includes a fuel flow path having a serpentine shape and is disposed between the fuel supply section and the nozzle, the fuel conditioner configured to guide the fuel supplied from the fuel supply section to the nozzle along the fuel flow path; and a nozzle casing having an opening into which the nozzle is inserted. The fuel conditioner extends from a first side of the nozzle casing toward a second side. The fuel conditioner and the nozzle casing form a first passage through which fuel may flow to the nozzle.

Abstract: A gas turbine combustor includes a pilot burner disposed on a central portion of a combustion liner; and a plurality of main burners disposed to surround the pilot burner. The pilot burner includes: a pilot nozzle disposed on the central portion of the combustion liner; and a pilot cone including a widened portion widening downstream away from a downstream end of the pilot nozzle, and an annular portion having an annular shape extending outward in a radial direction from a downstream opening end of the widened portion. The annular portion includes a first position in a circumferential direction and a second position different from the first position in the circumferential direction, and the first position and the second position have different widths in the radial direction.

Abstract: An air swirler arrangement comprises a coaxial arrangement of an inner and an outer air swirler passage. Each air swirler passage comprises a radial flow swirler. Air swirler arrangement comprises a coaxial arrangement of first, second and third members. Second member has radially extending upstream portion spaced axially from first member and a convergent portion. Third member has a radially extending upstream portion spaced axially from the upstream portion of second member and a radially inner surface having convergent and divergent downstream portions and a radially outer surface having a divergent downstream portion. First, second and third members the vanes of the radial flow swirlers is a monolithic structure. Plurality of circumferentially spaced passages are provided within the third member and each passage has an inlet in the surface and an outlet arranged to direct fluid onto the divergent portion of the surface or the surface of the third member.

Abstract: A combustor includes a fuel injector to inject fuel; a cooling passage configured to pass compressed air for cooling an outer surface of a liner forming a combustion chamber for containing combustion gas; and a nozzle part that is coupled to the liner and has a rear surface facing the combustion chamber, the nozzle part configured to mix the compressed air with the fuel from the fuel injector and to discharge some of the compressed air in the cooling passage to the rear surface in order to block an introduction of the combustion gas into the nozzle part. Before the compressed air introduced through air introduction holes formed in a nozzle casing is discharged through through-holes formed in a nozzle cap, some of the compressed air may be introduced through bypass holes formed in a bypass tube and then discharged to the combustion chamber through the gap.

Abstract: A turbine combustor assembly includes a fuel/air mixer assembly having a plurality of fuel/air mixer elements and a fuel injector coupled to the fuel/air mixer assembly. Each of the fuel/air mixer elements includes: a fuel/air mixer body having an internal cavity extending axially between an air inlet and an air outlet; and a prefilmer residing within an interior cavity of the fuel/air mixer body, the prefilmer including an axial inner air flow passage extending coaxially with an axial centerline of the internal cavity of the fuel/air mixer body and a radial fuel injection port into the air flow passage. The fuel injector is arranged to direct fuel into a plurality of the fuel/air mixer elements via the fuel injection port of the prefilmer of each fuel/air mixer element.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine comprising a bulkhead wall extended generally along a radial direction relative to a combustor centerline and a mixer disposed concentric to the nozzle centerline. The bulkhead wall defines one or more openings therethrough adjacent to a combustion chamber. Each of the openings defines a nozzle centerline extended along a lengthwise direction. The mixer comprises a retaining wall extended along a radial direction relative to the nozzle centerline and coupled to the upstream side of the bulkhead wall. The mixer further comprises an annular shroud extended at least partially through the one or more openings of the bulkhead wall. The mixer defines a cooling passage through the retaining wall and the annular shroud. The cooling passage defines an inlet opening through the retaining wall and an outlet opening through the annular shroud.

Abstract: An article and method of forming an article are provided. The article includes a body portion separating an inner region and an outer region, an aperture in the body portion, the aperture fluidly connecting the inner region and the outer region, and a baffle extending along at least a portion of an inner surface of the article, the baffle dividing the inner region into a plurality of sub-regions. The method of forming an article includes forming a body portion defining an inner region and an outer region, forming an aperture in the body portion, the aperture fluidly connecting the inner region to the outer region, and forming at least one baffle extending along at least a portion of an inner surface of the body portion, the at least one baffle dividing the inner region into a plurality of sub-regions. Also provided is a component including a cooling article disposed therein.

Abstract: Combustor assemblies are provided. An exemplary combustor assembly comprises an annular ceramic matrix composite (CMC) inner liner including an inner liner flange, an annular CMC outer liner including an outer liner flange, and an annular CMC combustor dome comprising a plurality of tiles positioned circumferentially adjacent one another. Each tile has a first end radially opposite a second end. The CMC inner liner, outer liner, and combustor dome form a combustor, and the CMC combustor dome is positioned at a combustor forward end. The combustor assembly also comprises a support structure for supporting the combustor and including an annular frame having a frame channel defining a groove and an inner and outer support flanges. The first end of each tile is disposed within the frame channel groove. The inner liner flange is secured to the inner support flange and the outer liner flange is secured to the outer support flange.

Abstract: The present disclosure is directed to a radially stacked fuel injection module for a segmented annular combustion system. The fuel injection module includes a bundled tube fuel nozzle portion and a plurality of fuel injection lances. In one instance, a first set of fuel injection lances and a second set of fuel injection lances share a housing with the bundled tube fuel nozzle. The first and second sets of fuel injection lances are radially stacked, each set of lances being in communication with a respective injector fuel plenum. The fuel injection lances deliver fuel to a respective premixing channel formed between the side walls of an integrated combustor nozzle, an aft end of which defines a turbine nozzle. In another instance, the bundled tube fuel nozzle may be provided as a radially inner fuel nozzle and a radially outer fuel nozzle, each with a respective bundled tube fuel plenum.

Abstract: A cooling system for a fuel injector system of a gas turbine engine has a heat exchanger for cooling a portion of diffuser case air and then routing the cooled diffuser case air through a sleeve that surrounds a fuel injector conduit located in at least the diffuser case plenum for minimizing fuel heat-up rates in the conduit. By minimizing fuel temperatures within the injector conduit, coking accumulation is thereby eliminated or reduced.

Abstract: Combustor assemblies for gas turbine engines are provided. For example, a combustor assembly comprises a combustor dome, a first heat shield having an edge, a second heat shield having an edge, and a seal extending from the edge of the first heat shield to the edge of the second heat shield such that the seal spans a gap between the first heat shield and the second heat shield. In another embodiment, the seal has a first contact portion contacting the edge of the first heat shield, a second contact portion contacting the edge of the second heat shield edge, and a connecting portion connecting the first portion and the second portion. The first contact portion and the second contact portion project away from the connecting portion. Methods for sealing between adjacent heat shields of a combustor assembly also are provided.

Abstract: Combustor assemblies having heat shields heat shield attachment features are provided. For example, a combustor assembly includes a dome plate defining first and second apertures, and a heat shield defining first and second openings. The heat shield includes a first cup extending about the first opening and a second cup extending about the second opening. The combustor assembly further includes a collar having a first frame at least partially surrounding the first cup and a second frame at least partially surrounding the second cup. The collar includes a first fastening feature and the dome plate includes a second fastening feature. The first fastening feature mates with the second fastening feature to couple the heat shield to the dome plate. The combustor assembly also may include an attachment piece configured to couple the heat shield to the dome plate. Methods for forming ceramic matrix composite (CMC) heat shields also are provided.

Abstract: A gas turbine system include a first body including a central passage extending from a mixing region to a central outlet at a first downstream end portion of the first body, a fuel outlet in the mixing region, an oxidant outlet in the mixing region, and an outer passage including a first passage extending in a downstream direction toward the first downstream end portion and a second passage extending in an upstream direction from the first downstream end portion into the mixing region.

Abstract: A swirler comprising an anti-rotation feature is provided. The swirler may be installed on a combustor within a gas turbine engine. The swirler may comprise a generally cylindrical profile. In this regard, the swirler may be configured to provide a generally uniform profile within the combustor. The swirler may comprise a floating collar with an anti-rotation feature. The swirler may also include a collar end plate that defines a slot that at least partially encloses the anti-rotation feature and minimizes rotational movement of the floating collar.

Abstract: A combustor assembly for a gas turbine engine defining an axial direction is provided. The combustor assembly generally includes a liner, an annular dome, and a deflector plate. The liner at least partially defines a combustion chamber. The annular dome defines a first cavity and has a plurality of impingement holes, and the deflector plate defines a conical surface and a flat surface, wherein a plurality of cooling holes are defined through the deflector plate. The annular dome and the deflector plate are positioned together to define a second cavity that is in fluid communication with the first cavity through the plurality of impingement holes. In addition, the second cavity is in fluid communication with the combustion chamber through the plurality of cooling holes.

Abstract: A combustion chamber of a gas turbine having a head plate and a combustion chamber head, where the head plate is connected to the combustion chamber, and where several heat shields arranged in the interior of the combustion chamber are bolted to the head plate, characterized in that each heat shield and the combustion chamber head are fastened to the head plate by means of joint bolted connections.

Abstract: A gas turbine engine is provided that includes a compressor section, a combustor section with a multiple of pre-swirlers, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said multiple of pre-swirlers.

Abstract: There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand.

Abstract: A gas turbine engine combustor conical-flat heat shield includes an annular conical section extending upstream from and being integral with a flat section with a flat downstream facing surface which may be generally perpendicular to or canted with respect to a centerline. The flat section includes radially outer and inner edges at least one of which is circular and circumscribed about a centerline and circumferentially spaced apart clockwise and counter-clockwise radial edges having an origin on the centerline. A gas turbine engine combustor includes conical-flat heat shields in one or more circular rows arranged in a non-symmetrical or asymmetrical pattern. Two or more groups (A, B, C) of the conical-flat heat shields in the circular rows may be mounted on a domeplate and one or more of the conical-flat heat shields is different in one or more of the groups (A, B, C).

Abstract: A combustor heat shield assembly comprises a circumferential array of heat shield panels individually mounted to an inner surface of a combustor shell. Each heat shield panel has opposed front and back faces, the back face facing the inner surface of the combustor shell and being spaced therefrom to define an air gap. The front and back faces have a perimeter including opposed lateral edges extending between opposed circumferentially extending edges. The lateral edges of adjacent heat shield panels have complementary non-linear profiles defining an asymmetric heat shield panel interface.

Abstract: A combustor includes an end cap having upstream and downstream surfaces and a cap shield surrounding the upstream and downstream surfaces. First and second sets of premixer tubes extend from the upstream surface through the downstream surface. A first fuel conduit supplies fuel to the first set of premixer tubes. A casing circumferentially surrounds the cap shield to define an annular passage, and a second fuel conduit supplies fuel through the annular passage to the second set of premixer tubes. A method for supplying fuel to a combustor includes flowing a working fluid through first and second sets of premixer tubes, flowing a first fuel into the first set of premixer tubes, and flowing a second fuel through an annular passage surrounding the end cap and into the second set of premixer tubes.

Abstract: A combustor includes a plurality of tubes arranged in a tube bundle and supported by at least one plate that extends radially within the combustor, wherein each tube includes an upstream end axially separated from a downstream end and provides fluid communication through the tube bundle. A flow conditioner extends upstream from the upstream end of one or more of the plurality of tubes, and a radial passage extends through the flow conditioner. A method for distributing fuel in a combustor including flowing a working fluid through a flow conditioner that extends from a tube that is configured in a tube bundle comprising a plurality of tubes and that is supported by at least one plate. The flow conditioner includes at least one radial passage to impart radial swirl to the working fluid. Flowing a fuel through an annular insert that is at least partially surrounded by the flow conditioner.

Abstract: Annular combustion chamber (10) to be fitted on a turbomachine and comprising a chamber end wall (22), a plurality of air and fuel injection systems (32) circumferentially distributed around an axis (34) of the combustion chamber (10) and mounted on said chamber end wall (22), and, an air manifold (100) associated with each injection system (32), comprising at least one wall (96, 98) mounted on the chamber end wall (22) and projecting in the upstream direction to form an obstacle to a circumferential airflow around the axis (34) of the combustion chamber (10), and an air inlet opening (88) formed at the upstream end of the air manifold (100) and opening radially outwards from an axis (44) of said injection system.

Abstract: A burner assembly having a fuel distribution ring, number of fuel nozzles which are mounted on the fuel distribution ring in the direction of low is provided. The fuel distribution ring has a ring-shaped surface in the direction of flow and wherein the fuel distribution ring center and an opposite outer outer side and wherein there is at least one slot on the surface between the fuel nozzles and the at least one slot extends on the surface from the outside to the inside. There is at least one recess arranged on the surface and the at least one recess also partially includes the outside of the fuel distributor.

Abstract: A gas turbine combustor having a burner including a plurality of fuel nozzles for injecting fuel, air hole plates positioned on a downstream side of the fuel nozzles and a plurality of air holes arranged in pairs with each of the fuel nozzles, and a combustion chamber for mixing fuel injected from the fuel nozzles and air injected from the air holes and injecting and burning the mixed fuel. Each of the fuel nozzles configuring the burners is provided with a projection in which a part of an outer edge of a section of the fuel nozzle is protruded outward; and the projection is arranged so as to be directed toward a center of the gas turbine combustor. The projection of the fuel nozzle is positioned on a downstream side of a flow of combustion air flowing around each of the fuel nozzles.

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 fuel injector device for an annular combustion chamber of a turbine engine, including a pilot circuit feeding an injector and a multipoint circuit feeding injection orifices formed in a front face of an annular ring mounted in an annular chamber, together with a thermal insulation mechanism insulating the front face and including an annular cavity formed around the injection orifices between the front face of the annular ring and a front wall of the annular chamber and configured to be filled in operation with air or with coked fuel.

Abstract: A swirler assembly in a gas turbine combustor includes a hub, a shroud, and a plurality of vanes connected between the hub and the shroud. The vanes include a high pressure side on which air and fuel impinge the vanes and a low pressure side. An air circuit is provided in each of the plurality of vanes receiving discharge air from a compressor. Each of the air circuits includes an air entry passage into the vanes and an air exit passage on the low pressure side of the vanes.

Abstract: A gas-turbine burner having a plurality of main swirl generators which each have 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 fuel nozzle assembly for a gas turbine, the assembly including: a cylindrical center body; a cylindrical shroud coaxial with and extending around the center body, and a turning guide having an downstream edge extending in a passage between the center body and an inlet to the shroud, wherein the turning guide extends only partially around the center body.

Abstract: A disclosed fuel injector provides mixing of fuel with airflow by surrounding a swirled fuel flow with first and second swirled airflows that ensures mixing prior to or upon entering the combustion chamber. Fuel tubes produce a central fuel flow along with a central airflow through a plurality of openings to generate the high velocity fuel/air mixture along the axis of the fuel injector in addition to the swirled fuel/air mixture.

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: Advanced gas turbines and associated components, systems and methods are disclosed herein. A gas turbine configured in accordance with a particular embodiment includes a rotor operably coupled to a shaft and a stator positioned adjacent to the rotor. A coolant line extends at least partially through the stator to transfer heat out of an air flow within a compressor section of the gas turbine.

Abstract: An assembly for use in a late lean injection system of a combustor of a combustion turbine engine, wherein the combustor includes an inner radial wall, which defines a primary combustion chamber downstream of a primary fuel nozzle, and an outer radial wall, which surrounds the inner radial wall forming a flow annulus therebetween, the assembly comprising: a boss rigidly secured to the inner radial wall, the boss being configured to define a hollow passageway through the inner radial wall; a transfer tube slideably engaged within the boss; a stop formed on the transfer tube; and damping means positioned between the boss and the stop.

Abstract: A fuel injector comprises a plurality of swirler vane passages defined between swirler vanes. Each vane is leant with respect to the true radius of the swirler. The pressure distribution through the swirler passages is improved and the flow of air over a prefilmer located at the radially outer edges of the swirl vanes is improved and consequently atomization of fuel is improved and levels of NOx is reduced.

Abstract: A system including a first multi-tube fuel nozzle including a plurality of first tubes extending in an axial direction, wherein each first tube of the plurality of first tubes includes a first air inlet, a first fuel inlet, and a first fuel-air mixture outlet, a second multi-tube fuel nozzle including a plurality of second tubes extending in an axial direction, wherein each second tube of the plurality of second tubes includes a second air inlet, a second fuel inlet, and a second fuel-air mixture outlet, and an aft plate including a plurality of first tube apertures and a plurality of second tube apertures, wherein the plurality of first tubes extend to the plurality of first tube apertures, and the plurality of second tubes extend to the plurality of second tube apertures.

Abstract: A system including a plurality of multi-tube fuel nozzles each having a plurality of tubes extending in an axial direction, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, and a fuel nozzle housing, including an outer wall extending circumferentially about a central axis, a plurality of radial walls extending from the outer wall inwardly toward the central axis, a plurality of fuel nozzle receptacles disposed within the outer wall, wherein the plurality of radial walls separate the plurality of fuel nozzle receptacles from one another, and the plurality of multi-tube fuel nozzles are disposed in the plurality of fuel nozzle receptacles a mounting structure including a plurality of radial support arms extending outwardly from the outer wall.

Abstract: A combustor includes a central fuel nozzle assembly and a plurality of outer fuel nozzle assemblies, each of the plurality of outer fuel nozzle assemblies having a center body and an outer shroud, the plurality of outer fuel nozzle assemblies being configured to abut one another in a surrounding relationship to the central cylinder such that no gaps are present between any two abutting ones of the plurality of outer fuel nozzle assemblies. One or more of the plurality of fuel nozzle assemblies may traverse axially back and forth according to embodiments of the invention.

Abstract: A turbomachine combustor includes a combustion chamber; a plurality of micro-mixer nozzles mounted to an end cover of the combustion chamber, each including a fuel supply pipe affixed to a nozzle body located within the combustion chamber, wherein fuel from the supply pipe mixes with air in the nozzle body prior to discharge into the combustion chamber; and wherein at least some of the nozzle bodies of the plurality of micro-mixer nozzles have axial length dimensions that differ from axial length dimensions of other of the nozzle bodies.

Abstract: A system including a plurality of multi-tube fuel nozzles each having a plurality of tubes extending in an axial direction, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, a fuel nozzle housing including a first outer wall extending circumferentially about a central axis, wherein the plurality of multi-tube fuel nozzles are disposed in the fuel nozzle housing, an inlet flow conditioner removably coupled to a first end portion of the first outer wall, wherein the inlet flow conditioner includes a plurality of air openings, and an aft plate assembly removably coupled to a second end portion of the first outer wall, wherein the aft plate assembly includes an aft plate having a plurality of tube apertures, and the plurality of tubes extend to the plurality of tube apertures.

Abstract: The system may include a turbine engine. The turbine engine may include a fuel nozzle. The fuel nozzle may include an air path. The fuel nozzle may also include a fuel path such that the fuel nozzle is in communication with a combustion zone of the turbine engine. Furthermore, the fuel nozzle may include a resonator. The resonator may be disposed in the fuel nozzle directly adjacent to the combustion zone.