Abstract: A system for distributing compressed air in a combustor of a gas turbine engine. The system may include a flow splitter, a center duct, an outer duct, and an inner duct configured to separate and receive compressed air from a prediffuser exit, and to route separate flows of the compressed air to separate downstream locations for the combustion reaction and for cooling. The system may include an axial mixer configured to axially receive a flow of the compressed air and to direct the flow to mix with fuel provided by an injector.

Abstract: A combustion burner including a fuel nozzle, a burner tube that surrounds the fuel nozzle to form an air passage between the burner tube and the fuel nozzle, swirler vanes that are arranged in a plurality of positions in a circumferential direction on an external circumferential surface of the fuel nozzle, each of which extends along an axial direction of the fuel nozzle, and gradually curves from upstream to downstream, a liquid fuel injecting hole that is formed on the fuel nozzle, and from which a liquid fuel is injected to a vane pressure surface of each of the swirler vanes, and a cooling unit that cools a part of the vane pressure surface on which the liquid fuel hits. The cooling unit injects a mixed fuel prepared by mixing water and the liquid fuel evenly from the liquid fuel injecting hole to the vane pressure surface of the swirler vane.

Abstract: One embodiment is a gas turbine engine premix injector including a mixing duct, a plurality of air inlet slots leading to the duct, a plurality of gaseous fuel inlet apertures leading to the duct, a plurality of vanes positioned in the duct downstream from said air inlets, a plurality of liquid fuel inlet apertures leading to the duct, and a plurality of discharge windows positioned between the vanes.

Abstract: A burner for a gas turbine including a main burner and a pilot burner is provided. The main burner has a supporting structure, a heat shield and a holder for the heat shield, and wherein the holder is at least partially located within the supporting structure and the heat shield is at least partially located within the holder. The heat shield is fastened to the holder by a force-fit and/or a frictional connection.

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 fuel nozzle for use in a gas turbine includes a pre-mix flow passage for directing a flow segment of a flow of a working fluid through the fuel nozzle. A first swirler vane and a second swirler vane extend within the pre-mix flow passage. The first swirler vane provides a first wake region within the flow segment. The second swirler vane provides a second wake region within the flow segment. A fuel injection peg is disposed downstream from the first swirler vane and the second swirler vane. The fuel injection peg is positioned within the flow segment between the first wake region and the second wake region.

Abstract: A burner for a gas turbine including a burner housing is provided. Provided is a lean-rich partially premixed low emission burner for a gas turbine combustor providing stable ignition and combustion process at all engine load conditions. At the upstream end of that burner a pilot combustor creates a flow of an unquenched concentration of radicals and heat. Respectively provided is a plurality of quarl sections surrounding the exit of the pilot combustor, a main combustion room defined downstream the pilot combustor and at least a first channel defined as an annular space between an upstream quarl section and the closest downstream quarl section providing air and fuel to a main flame in the combustion room.

Abstract: A system includes a fuel nozzle. The fuel nozzle includes a hub having an axis, a shroud disposed about the hub, an airflow path between the hub and the shroud, multiple vanes extending between the hub and the shroud, and a first fuel path leading to multiple first fuel outlets disposed on the multiple vanes. The fuel nozzle also includes a second fuel path leading to multiple second fuel outlets disposed on at least one of the hub and/or the shroud, wherein the multiple second fuel outlets are disposed at an axial distance upstream from a downstream end of the hub.

Abstract: A system includes a fuel nozzle. The fuel nozzle includes a hub, a shroud disposed about the hub, an airflow path between the hub and the shroud, multiple first fuel outlets disposed on the hub, and multiple swirl vanes disposed in the airflow path downstream from the multiple first fuel outlets.

Abstract: A burner including a gas injector for injecting fuel into the burner for a gas turbine engine is provided. The burner is provided with a quarl, which surrounds a combustion room, at least a first channel emerging into the combustion room for providing the combustion room with air mixed with fuel, a swirler for mixing the air and the fuel located at the inlet of the channel, at least one tube for the provision of fuel at an inlet of at least one channel, wherein the tube is provided with a plurality of diffuser holes distributed along the tube acting as gas injectors for effectively distributing fuel in a flow of air passing through the channel.

Abstract: A method for operating a gas turbine engine including a combustor assembly includes channeling a first fluid through a plurality of first nozzles into the combustor and igniting the first fluid downstream from the first nozzles. The method also includes increasing the operating speed of the engine and attaining a first predetermined percentage of a baseload by channeling the first fluid only through the first nozzles and then channeling a second fluid through a second nozzle into the combustor. The method also includes igniting the second fluid within the combustor downstream from the second nozzle. The method further includes channeling the second fluid to the first nozzles when the engine attains a second predetermined percentage of the baseload. The second predetermined percentage of the baseload is greater than the first predetermined percentage of the baseload. The method also includes terminating a flow of the first fluid through the first nozzles.

Abstract: A combustion chamber may include a first surface and a second surface interconnected by a wall forming a chamber having a central axis. The first surface may define an exhaust opening and the second surface defining a pilot opening, wherein the exhaust opening and the pilot opening align along the central axis. A plurality of inlet ports may be configured to deliver air to the chamber. A plurality of fuel ports may be arranged on an inside of the second surface to deliver fuel to the chamber. The air flow from the inlet ports and fuel from the fuel ports may oppose each other to create a vortex of product proximal to the second surface.

Abstract: The present application and the resultant patent provide a diffusion combustor fuel nozzle for a gas turbine engine. The fuel nozzle may include one or more gas fuel passages for one or more flows of gas fuel, a swirler surrounding the one or more gas fuel passages and positioned about a downstream face of the fuel nozzle, a number of swirler gas fuel ports defined in the swirler, and a number of downstream face gas fuel ports defined in the downstream face of the fuel nozzle. The swirler may include a number of swirl vanes and a number of air chambers defined between adjacent swirl vanes. The present application and the resultant patent further provide a method of operating a diffusion combustor fuel nozzle of a gas turbine engine.

Abstract: A pilot nozzle, a gas turbine combustor and a gas turbine are provided with a nozzle main body having a fuel passage, a cover ring arranged at an outside of a front end-outer peripheral portion of the nozzle main body at a predetermined interval to form an inner air passage and capable of injecting air toward a front side of the nozzle main body, a plurality of nozzle tips that includes a fuel injection nozzle attached to a front end portion of the cover ring at a predetermined interval in a circumferential direction to communicate with the fuel passage and is able to inject fuel toward an outside of injection air from the inner air passage, and a swirling force application unit that applies a swirling force to air injected through the inner air passage.

Abstract: A swirler, such as for swirling air in a fuel injector of a gas turbine engine, includes a swirler body with opposed inlet and outlet ends with a swirler wall extending therebetween along a longitudinal axis. The inlet end of the swirler body defines an inlet opening. A plurality of swirl slots is defined through a portion of the swirler wall that converges toward the longitudinal axis in a direction from the inlet opening toward the outlet end of the swirler body. The swirl slots are radially off-set with respect to the longitudinal axis for imparting swirl on a flow passing from the inlet opening, through the swirl slots, and past the outlet end of the swirler body.

Abstract: A combustion burner 10A according to one embodiment of the present invention includes: a fuel nozzle 110; a burner tube 120 forming the air passage 111 between the burner tube 120 and the fuel nozzle 110; swirler vanes (swirler vanes) 130 arranged in a plurality of positions in the circumferential direction on the external circumferential surface of the fuel nozzle 110, each extending along the axial direction of the fuel nozzle 110, and gradually curving from upstream toward downstream; and a liquid fuel injecting hole 133A from which a liquid fuel is injected to a surface of each of the swirler vanes 130. The combustion burner 10A also includes multi-purpose injecting holes 11-1 to 11-3 as a cooling unit that cools a part of a vane pressure surface 132a of the swirler vane 130 on which the liquid fuel LF hits.

Abstract: A combustor for a gas turbine includes a fuel nozzle having a central swirler that circumferentially surrounds a downstream end of the fuel nozzle. A primary combustion zone is defined within the central swirler. The combustor further includes an outer swirler that circumferentially surrounds at least a portion of the central swirler and a venturi that is disposed downstream from the primary combustion zone. The venturi includes an inner surface. The central swirler imparts angular swirl to a compressed working fluid so as to rapidly mix and react the fuel rich primary zone products with the working fluid. The outer swirler imparts angular swirl to a compressed working fluid so as to provide a cooling boundary layer along the inner surface of the venturi.

Abstract: A system includes a fuel nozzle. The fuel nozzle includes a central hub with a first annular passage extending along a longitudinal axis of the fuel nozzle. A flow conditioner is disposed along the first annular passage. The flow conditioner includes at least one of a straightening vane, a mesh screen, or a multi-passage body having multiple passages generally parallel with the longitudinal axis. An outer shroud is disposed about the central hub to define a second annular passage extending along the longitudinal axis of the fuel nozzle.

Abstract: A fuel nozzle for a gas turbine engine is disclosed, the fuel nozzle comprising at least one unitary component made using a rapid manufacturing process. In one aspect, the rapid manufacturing process is a laser sintering process. Unitary components disclosed include a conduit, swirler, distributor, venturi and a centerbody.

Abstract: The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner.

Abstract: Provided is a gas turbine combustor capable of reducing the size of a low-temperature air layer of pilot air formed between a pilot flame and a premixed flame and of improving the flame stability of the premixed flame. A gas turbine combustor, which is provided with a pilot burner that is provided at the center portion of a combustor main body formed in a cylindrical shape to form a pilot flame, and a plurality of main burners arranged so as to surround the outer periphery of the pilot burner to form a premixed flame, includes, as the ignition improving part, a channel blocking member that reduces the size of the low-temperature air layer of the pilot air formed between the pilot flame and the premixed flame.

Abstract: A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine.

Abstract: A swirl generator, a method for preventing flashback in a burner with a swirl generator and a burner are provided. The swirl generator includes a central fuel distributor element, an outer wall enclosing the central fuel distributor element and bounding an axial flow channel for combustion air, swirl vanes extending in a radial direction to the outer wall and giving the flowing combustion air a tangential flow component, and a separating wall enclosing the central fuel distributor element and being positioned radially within the outer wall. The separating wall divides the flow channel into a radially inner channel segment and a radially outer channel segment. The radially inner channel segment allows the combustion air to pass without giving it a tangential flow component or while giving it a tangential flow component counter to the orientation of the tangential flow component in the radially outer channel segment.

Abstract: A combustor includes: a combustor basket configured to surround a fuel nozzle from an outer circumferential side, and a plurality of connecting members formed in a circumferential direction at intervals to connect a rear end of the combustor basket and a casing and configured to define a flow path through which compressed air introduced into the combustor basket flows. A circulation direction of the compressed air flowing through the flow path is configured to be reversed at a rear end of the combustor basket and the compressed air is introduced into the fuel nozzle. The flow path is partially or entirely inclined in the circumferential direction to blow the compressed air.

Abstract: A gas turbine combustor of the present invention comprises a fuel injector for injecting a fuel toward a combustion chamber; a swirler which takes-in compressed air generated in a compressor and swirls the compressed air, in the vicinity of the fuel injector; a tubular guide member for guiding the compressed air taken-in from the swirler, to the combustion chamber; and a heat shield having a cylindrical portion located outward relative to the guide member; wherein the cylindrical portion has a purge hole; and air is introduced through the purge hole and is supplied to a space formed between the guide member and the cylindrical portion.

Abstract: A combustion method and combustor for use in a jet engine, the jet engine having a compressor portion and a turbine portion. The combustor includes an outer tube having a central axis that extends longitudinally intermediate the compressor portion and turbine portion and is positioned to receive air discharged by the compressor portion. An inner tube is positioned within the outer tube that includes an associated outer surface spaced from the inner surface of the outer tube thereby defining a combustion chamber. The outer tube and inner tube include fluid directing structure for communicating at least some of the air discharged by the compressor portion to the combustion chamber. The fluid directing structure directs air into the combustion chamber in a direction offset from the central axis, thereby causing rotation or swirling of the air about the central axis.

Abstract: A port (60) for axially staging fuel and air into a combustion gas flow path 28 of a turbine combustor (10A). A port enclosure (63) forms an air path through a combustor wall (30). Fuel injectors (64) in the enclosure provide convergent fuel streams (72) that oppose each other, thus converting velocity pressure to static pressure. This forms a flow stagnation zone (74) that acts as a valve on airflow (40, 41) through the port, in which the air outflow (41) is inversely proportion to the fuel flow (25). The fuel flow rate is controlled (65) in proportion to engine load. At high loads, more fuel and less air flow through the port, making more air available to the premixing assemblies (36).

Abstract: A system for a gas turbine engine includes a combustor assembly coupled to a plurality of first nozzles having outlet passageways coupled to a first fuel source and first passageways coupled to a second fuel source. The system also includes the combustor assembly coupled to a second nozzle that has second passageways coupled to the second fuel source. An associated control system, coupled to the combustor assembly, is configured to channel a first fuel from the first fuel source through the outlet passageways, and channel a second fuel from the second fuel source through the second passageways when the engine attains a predetermined load. The control system is further configured to channel the second fuel from the second fuel source through the first passageways and to reduce a flow of the first fuel entering the first nozzles when the engine is at a load greater than a first predetermined load.

Abstract: An arrangement for injection of an emulsion of a first fluid and a second fluid into a flame of a burner has a central gas duct, an outer gas channel disposed coaxially with the gas duct, and a fluid channel disposed coaxially between the gas duct and the outer gas channel The central gas duct and the fluid channel are separated by a first frustoconical wall. The fluid channel and the outer gas channel are separated by a second frustoconical wall. The arrangement is mounted concentrically surrounding a heat source which provides through the gas duct hot gases being directed into the flame of the burner. Further, the arrangement includes a mixing device for forming an emulsion of the first fluid and the second fluid, for supplying the emulsion into the fluid channel and for injecting the emulsion from the fluid channel into the flame.

Abstract: A combustor generally includes a shroud that that defines at least one inlet passage extends circumferentially inside the combustor. A first plate extends radially inside the shroud downstream from the inlet passage. The first plate defines at least one inlet port, at least one outlet port and at least partially defines at least one fuel nozzle passage. The shroud at least partially surrounds a sleeve that extends around the fuel nozzle passage. A tube at least partially surrounded by the sleeve may extend through the fuel nozzle passage. The tube, the sleeve, and the first plate may at least partially define an outlet passage. A first fluid flow path generally extends from the at inlet passage to the inlet port, and a second fluid flow path extends generally from the outlet port to the outlet passage.

Abstract: In accordance with one aspect of the disclosure, a swirler is disclosed. The swirler may include an outer shroud and inner shroud. The inner shroud may be positioned radially inside the outer shroud. At least one of the outer shroud and inner shroud may have a major diameter greater than that of a minor diameter such that the shrouds define an ovate shape. The swirler may further include a plurality of vanes which may be positioned between the inner and outer shrouds.

Abstract: In accordance with one aspect of the disclosure, a swirler is disclosed. The swirler may include an outer shroud and inner shroud. The inner shroud may be positioned radially inside the outer shroud. At least one of the outer shroud and inner shroud may have a major diameter which is larger than a minor diameter such that the shrouds define an oblong shape. The swirler may further include a plurality of vanes which may be positioned between the inner and outer shrouds.

Abstract: A fuel injector for use in a gas turbine engine combustor assembly. The fuel injector includes a main body and a fuel supply structure. The main body has an inlet end and an outlet end and defines a longitudinal axis extending between the outlet and inlet ends. The main body comprises a plurality of air/fuel passages extending therethrough, each air/fuel passage including an inlet that receives air from a source of air and an outlet. The fuel supply structure communicates with and supplies fuel to the air/fuel passages for providing an air/fuel mixture within each air/fuel passage. The air/fuel mixtures exit the main body through respective air/fuel passage outlets.

Abstract: The present invention relates to a premix burner of a combustion chamber of a gas turbine with at least one annular duct for supplying air and fuel, including a radially outer and a radially inner combustion chamber wall relative to a burner central axis and with at least one swirler arranged in the duct, said swirler including several flow-guiding elements distributed around the circumference of the duct cross-section, characterized in that at least one radially inner duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type.

Abstract: A gas turbine engine lean burn fuel injector includes a fuel injector head which has a first air swirler, a second air swirler arranged around the first air swirler, a pilot fuel injector arranged radially between the first air swirler and the second air swirler. A third air swirler arranged around the second air swirler, a fourth air swirler arranged around the third air swirler and a main fuel injector arranged radially between the third air swirler and the fourth air swirler. A shroud is arranged around the fourth air swirler. A downstream end of the shroud is generally circular in cross-section in a plane perpendicular to the axis (Y) of the fuel injector head and an upstream end of the shroud is generally elliptical in cross-section in a plane perpendicular to the axis (Y) of the fuel injector head.

Abstract: A burner for a gas turbine engine is provided. The burner includes a radial swirler for creating a swirling fuel/air mix, a combustion chamber where combustion of the swirling fuel/air mix occurs, and a pre-chamber located between the radial swirler and the combustion chamber. The radial swirler includes a plurality of vanes arranged in a circle, generally radially inwardly extending flow slots are defined between adjacent vanes in the circle, each flow slot includes a radially outer inlet end, a radially inner outlet end, first and second generally radially inwardly extending sides provided by adjacent vanes, and a base and top. A flow slot includes a first gas fuel injection hole in its base and a flow slot includes a second gas fuel injection hole in its first side wherein the amounts of gas fuel injected via the first and second gas fuel injection holes are independently variable.

Abstract: A combustor dome assembly includes a dome plate with upstream and downstream faces and an aperture therein; a deflector with body and a tube section, the deflector fixedly attached to the dome plate adjacent the downstream face with the tube section in the aperture; a retainer having an annular body with a central opening therein, an annular flange, and at least one hook defining a slot, the retainer being fixedly attached to the dome plate adjacent the upstream face of the dome plate with the flange received in the aperture; and an annular mixer with a radial array of swirl vanes and an annular mounting flange, with a retention tab extending radially outward from the mounting flange. The mixer is disposed adjacent the upstream face with the mounting flange against the retainer. The retention tab is engaged in the slot to prevent axial withdrawal of the mixer from the retainer.

Abstract: A pilot fuel injector tip includes concentric primary and secondary pilot fuel nozzles having a circular primary exit axially aft and downstream of an annular secondary exit respectively. A fuel nozzle assembly includes a pilot swirler flowpath section having an annular inwardly tapering conical flowpath section surrounding primary and secondary exits. An inwardly tapering conical wall section radially inwardly bounding flowpath section defines a conical surface. Exits are located at or axially forward or upstream of the conical surface. An annular secondary fuel supply passage in secondary pilot fuel nozzle includes a secondary fuel swirler with an array of helical spin slots that may have rectangular cross sections. A chamfered leading edge of an annular wall section disposed between an outer pilot swirler and an inlet to an injector cooling flowpath surrounding the second pilot swirler includes a radially inwardly facing conical chamfered surface for deflecting dirt from cooling flowpath.

Abstract: A combustor (10) includes a liner (12) that defines a combustion chamber (18) first pre-mix chamber (14) is upstream of the combustion chamber, and a fuel plenum (40) in fluid communication with the first pre-mix chamber surrounds a least a portion of the first pre-mix chamber. A method of supplying a fuel to combustor includes flowing the fuel over an outer surface of a first pre-mix chamber and into the first pre-mix chamber.

Abstract: A swirler for a gas turbine engine fuel injector comprises a swirler body extending from an upstream end to a downstream end. A fuel injector extends into the body, and has a downstream end for injecting fuel in a downstream direction. A first flow path directs air in a first circumferential direction about a central axis of the swirler body. A second flow path directs air to intermix with the air in the first flow path, and then to mix with fuel injected by the fuel injector. The first and second flow paths are positioned to inject air upstream of the downstream end of the fuel injector where fuel is injected. The first flow path is provided in a greater volume than the volume provided in the second flow path. The second flow path directs air at a location downstream of the first flow path.

Abstract: A swirler arrangement for injecting a fluid into a tubular swirling chamber is provided. The swirler arrangement includes a first radial swirler device and a second radial swirler device. The swirler arrangement is fixed around an internal circulation zone of the tubular swirling chamber. The first radial swirler device includes first vanes, wherein the first vanes are formed to inject the fluid into the internal circulation zone with a first injecting angle. The second radial swirler device includes second vanes, wherein the second vanes are formed to inject the fluid into the internal circulation zone with a second injecting angle. The first injecting angle and the second injecting angle are defined by an angle between an injecting direction of the fluid and the tangential direction along the inner surface. A method of injecting a fluid into a tubular swirling chamber by the swirler arrangement is also provided.

Abstract: A burner for a gas turbine is provided. The burner has a pilot combustor, a supply module providing pilot fuel and air into a pilot combustion room enclosed by a pilot burner housing having a tapered exit throat discharging radicals and heat generated in a pilot combustion zone into a main combustion room. An equalizer has holes for main flow air entering a cavity in a radial direction with regard to a burner axis defined by centers of pilot combustion zone and main combustion zone. A fuel injector is downstream the equalizer to supply flow fuel into flow air. A swirler is downstream the injector to give flow distribution to the flow fuel and air entering the main combustion room. A channel leads from the equalizer to the swirler arranged circumferentially around the pilot burner housing to direct the main air flow from the equalizer in axial and circumferential direction.

Abstract: A fuel nozzle guide plate has mirror symmetry about a first plane perpendicular to the fuel nozzle guide plate and mirror symmetry about a second plane parallel to the fuel nozzle guide plate. The fuel nozzle guide plate is attached to an air swirler, which is mounted in an opening of a bulkhead for mixing air and fuel at an upstream side of a combustor.

Abstract: Disclosed is a method of operating a secondary fuel nozzle for a turbomachine combustor including delivering a flow of pilot fuel through a pilot fuel channel toward a combustion zone and delivering a flow of air through a plurality of transfer passages arranged around the pilot fuel channel toward the combustion zone. The flow of pilot fuel and the flow of air are combusted in the combustion zone, and a flow of transfer fuel is delivered through the plurality of transfer passages for combustion. A secondary fuel nozzle includes a pilot fuel channel extending along the fuel nozzle to deliver a flow of pilot fuel to a combustion zone. A plurality of transfer passages are arranged around the pilot fuel channel and are configured to deliver a flow of air for combustion with the flow of pilot fuel and to deliver a flow of transfer fuel to the combustion zone.

Abstract: A swirling device for injecting a medium into a turbine is provided. The swirling device includes a central axis, a central passage in an axial direction along the central axis and an outer perimeter. The swirling device further includes a first duct and a second duct. The first duct and the second duct are adapted for guiding the medium from a region surrounding the outer perimeter to the central passage. The first duct includes a first depth in the axial direction and the second duct includes a second depth in the axial direction. The first depth and the second depth are different.

Abstract: A combustor 500 is composed of a plurality of premixed combustion burners 100 each comprising a fuel nozzle 110 provided in a burner tube 120, the fuel nozzle 110 having a plurality of swirl vanes 130 on an outer peripheral surface thereof. Injection holes 133a, 133b are formed in each swirl vane 130. Staging control is exercised such that when a gas turbine is in a full load state, a fuel is injected through the injection holes 133a, 133b of all the swirl vanes 130, and when the gas turbine is under a partial load, the fuel is injected only through the injection holes 133a, 133b of a specific number of the swirl vanes 130 adjacent in a circumferential direction, and fuel injection through the injection holes 133a, 133b of the remaining swirl vanes 130 is stopped.

Abstract: A dual fuel nozzle for a gas turbine combustor includes a hub defining a fuel inlet and a plurality of liquid fuel jets disposed at a downstream end of the hub. The fuel jets are oriented to eject liquid fuel radially outward from the hub. An annular air passage includes a swirler that imparts swirl to air flowing in the annular air passage, and a splitter ring is disposed in the annular air passage and surrounds the plurality of liquid fuel jets. The nozzle allows liquid fuels to be injected into a swirling annular airstream and then atomized, dispersed and vaporized inside a lean premixing dual fuel nozzle for a gas turbine combustor.

Abstract: A gas turbine combustor includes a burner disposed upstream of the combustion chamber for supplying a gas and air to an interior of the combustion chamber to hold a flame provided with a first swirler in which a gas hole and an air hole are alternately formed in a circumferential direction. A first gas is supplied to the gas hole in the first swirler and air is supplied to the air hole. A swiveling flow path is formed in the gas hole and the air hole in the burner to swivel the gas and the air and supply the gas and the air to the interior of the combustion chamber, a second gas hole is formed in the swiveling flow path in at least one of the air hole and the gas hole, and a second gas is supplied through the second gas hole.

Abstract: A flameless burner for a gas turbine engine includes a burner body having a longitudinal axis, an upstream section and a downstream section. The upstream section of the burner body defines a primary swirl generating chamber having air swirlers associated therewith. The primary swirl generating chamber is adapted and configured to receive compressor discharge air through the air swirlers. The strong swirl of the compressor discharge air forms a recirculation zone that entrains combustion product gases toward the burner body. Fuel injectors are operatively connected to the downstream section of the burner body for issuing fuel into the recirculated combustion product gases.

Abstract: The invention refers to burner arrangement for producing hot gases to be expanded in a gas turbine, including a burner inside a plenum, where the burner has means for fuel injection, means for air supply and means for generating an ignitable fuel/air mixture inside the burner, and a combustion chamber following downstream said burner having an outlet being fluidly connected to the gas turbine. The invention is characterized in that the means for air supply includes at least two separate flow passages, and that the one of the two flow passages is fed by a first supply pressure and the other flow passage is fed by a second supply pressure.