Abstract: A combustor for a gas turbine engine includes a combustor liner and a dilution chute integral and conformal with the combustor liner to provide an outlet into the combustor for fuel, wherein the dilution chute has at least one wall with a tapered edge extending into an interior of the combustor.

Abstract: A combustor and a gas turbine including the same which can reduce a loss of pressure and enhance a cooling efficiency of a liner and transition piece are provided.

Abstract: A method is disclosed for controlling fuel injection in a reheat combustor of a gas turbine combustor assembly including a combustor casing defining a gas flow channel and a plurality of injection nozzles distributed in or around the gas flow channel; the method includes the step of distributing fuel among the injection nozzles according to a non-uniform distribution pattern.

Abstract: Various embodiments include a trapped vortex combustor and a method for operating trapped vortex combustor. In one embodiment, the trapped vortex combustor comprises a trapped vortex combustion zone and at least one secondary combustion zone disposed downstream of the trapped vortex combustion zone. The trapped vortex combustion zone is operable to receive and combust a first fuel and a first air and produce a first combustion product flowing toroidally therein. The at least one secondary combustion zone is operable to receive and combust the first combustion product and at least one second injection consisting of fuel and/or air and produce at least one second combustion product therein. The combustor may reduce the residence time of the highest temperature combustion products and achieve the lower NOx emission.

Abstract: A sliding part with a wear resistant coating includes a sliding part, and a wear resistant coating provided on a sliding surface of the sliding part, and made of a cobalt alloy containing chromium and silicon. In the wear resistant coating, oxide particles are dispersed which include an oxide containing chromium and silicon, and have a particle size of 100 ?m or less when a cross section of the wear resistant coating is observed using an optical microscope with a magnification of 100 times.

Abstract: A main mixer including a swirler along an axis, the swirler including an outer swirler with a multiple of outer vanes, and a center swirler with a multiple of center vanes and a swirler hub along the axis, the swirler hub including a fuel manifold and an inner swirler with a multiple of inner vanes that support a centerbody, the multiple of inner vanes interconnect the fuel manifold and the centerbody.

Abstract: A combustor includes: a combustion liner having a combustion chamber formed therein; and a fuel injector mounted to a top portion of the combustion liner, and including a fuel injection member having a plurality of fuel injection annular portions and an air guide member including a plurality of combustion air annular portions that guide air for combustion. The fuel injection annular portions and the combustion air annular portions are arranged concentrically and alternately. The fuel injector injects fuel and air into the combustion chamber. Each of the fuel injection annular portions includes a plurality of fuel injection holes that are open in a radial direction thereof, and each of the combustion air annular portions includes a plurality of air guide grooves that are open in an axial direction thereof, and guide the air to the fuel jetted from the fuel injection holes.

Abstract: A combustor can include: a combustion chamber; a cap covering the combustion chamber; a first swirler inlet cup passing through the cap and having a first length and a first bell mouth; a second swirler inlet cup passing through the cap and having a second length and a second bell mouth; and a third swirler inlet cup passing through the cap and having a third length and a third bell mouth, wherein the first length is different from the third length, and wherein the first bell mouth is different from the third bell mouth. The first, second, and third swirler inlet cups can have different diameters, respectively.

Abstract: A method of starting a gas turbine is provided, comprising the step of providing an apparatus for regulating the flow of fuel in a gas turbine, the apparatus comprising a main line connecting a fuel source to a nozzle array; and an auxiliary line connecting the fuel source to the nozzle array. The method further comprises the steps of keeping the main line sealed while increasing the auxiliary line fuel flow rate; firing the gas turbine while keeping the main line sealed; and after the combustion has started in the gas turbine, opening the main line for increasing the main line fuel flow rate; wherein the auxiliary line maximum flow rate is less than the main line maximum flow rate.

Abstract: A gas turbine facility 10 of an embodiment includes: a combustor 20; a cylinder 80 dividing a space between a combustor casing 70 and the combustor 20; and a turbine 25 rotated by a combustion gas exhausted from the combustor 20. The gas turbine facility 10 includes: a heat exchanger 24 which cools the combustion gas; a pipe 42 through which a part of the combustion gas cooled in the heat exchanger 24 passes in the heat exchanger 24 to be heated and is guided to a space between the combustor 20 and the cylinder 80; a pipe 44 which guides another part of the combustion gas cooled in the heat exchanger 24 to a space between the combustor casing 70 and the cylinder 80; and a pipe 45 which exhausts a remaining part of the combustion gas cooled in the heat exchanger 24 to the outside.

Abstract: A method of operating a control unit for controlling at least two different input fuel flows to a combustion device, e.g. a gas turbine includes the step of determining on the basis of at least one operating parameter whether the combustion device is in a predefined operating stage. In response hereto, generating a control signal configured for setting a ratio of at least two different input fuel flows to a predetermined value (psc1, psc3) for a predetermined time (dt) in case the combustion device is in the predefined operating stage.

Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array. Each transition duct includes an inlet, an outlet, and a passage defining an interior and extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. Each transition duct further includes an upstream portion and a downstream portion. The turbomachine further includes a late injection assembly disposed between the upstream portion and the downstream portion of a transition duct and which provides fluid communication for an injection fluid to flow into the interior downstream of the inlet of the transition duct. The late injection assembly includes a late injection ring.

Abstract: A ducting arrangement (10) in a combustion stage downstream of a main combustion stage of a gas turbine engine is provided. A duct (18) is fluidly coupled to receive a cross-flow of combustion gases from the main combustion stage. Duct (18) includes a duct segment (23) with an expanding cross-sectional area (24) where one or more injector assemblies (26) are disposed. Injector assembly (26) includes one or more reactant-guiding structures (27) arranged to deliver a flow of reactants into the downstream combustion stage to be mixed with the cross-flow of combustion gases. Disclosed injector assemblies are arranged in expanding cross-sectional area (24) to reduce total pressure loss while providing an effective level of mixing of the injected reactants with the passing cross-flow. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

Abstract: A system is provided with a turbine combustor having a first diffusion fuel nozzle, wherein the first diffusion fuel nozzle has first and second passages that separately inject respective first and second flows into a chamber of the turbine combustor to produce a diffusion flame. The first flow includes a first fuel, and the second flow includes a first oxidant and a first diluent. The system includes a turbine driven by combustion products from the diffusion flame in the turbine combustor. The system also includes an exhaust gas compressor, wherein the exhaust gas compressor is configured to compress and route an exhaust gas from the turbine to the turbine combustor along an exhaust recirculation path.

Abstract: Embodiments of the present disclosure provide an apparatus including: a first injector located on a surface of a turbine nozzle of a turbine stage positioned downstream of a reheat combustor, wherein the turbine stage includes the turbine nozzle and a turbine blade row; a second injector located on a wall of the reheat combustor; and at least one conduit in fluid communication with each of the first injector and the second injector, wherein the at least one conduit delivers at least one of a fuel from a fuel supply line and a carrier gas to an aft section of the reheat combustor through at least one of the first injector and the second injector, and wherein the aft section is positioned downstream of a combustion reaction zone in the reheat combustor.

Abstract: Aspects of the present disclosure provide an apparatus including: an injector in fluid communication with an aft section of a reheat combustor in a power generation system, the aft section being positioned downstream of a combustion reaction zone in the reheat combustor, and positioned upstream of a turbine stage of the power generation system, wherein the turbine stage includes a turbine nozzle and a turbine blade row; and a conduit in fluid communication with the injector, wherein the conduit delivers at least one of a fuel from a fuel supply line and a carrier gas to the injector.

Abstract: A combustor has at least one premixing burner for premixing fuel with air and jetting the mixed gas into a chamber for combustion. A cylinder attached to an outer circumferential portion of an end face of the burner is provided with air supply holes. An interval D1 defined between the adjacent air supply holes and an interval D2 defined between each air supply hole and the end face of the burner are each made narrower than the quenching distance in the premixed gas jetted from the premixing burner.

Abstract: It is described a combustion device control unit and a combustion device, e.g. a gas turbine, which determine on the basis of at least one operating parameter whether the combustion device is in a predefined operating stage. In response hereto, there is generated a control signal configured for setting a ratio of at least two different input fuel flows to a predetermined value (psc1, psc3) for a predetermined time (dt) in case the combustion device is in the predefined operating stage.

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 gas turbine engine comprises a combustion system comprising a secondary annular combustor and a primary combustor in fluid communication with the secondary combustor, a secondary fuel injector associated with the secondary combustor, a primary fuel injector associated with the primary combustor, and a ECU controlling fuel delivery to the secondary and primary fuel injectors. The primary fuel injector delivers fuel to the primary combustor. The ECU allows fuel to be delivered to the secondary fuel injector in addition to the primary fuel injector only when a fuel amount higher is requested delivered by the primary fuel injector. A method of operating a gas turbine engine is also presented.

Abstract: An internal combustion burner including a combustion chamber supplied with fuel and with oxidant and at least two combustion devices supplied with oxidant and with fuel. Combining two combustion devices of distinct configurations, which respectively generate two distinct types of flames with a system for cooling the walls of the burner by introducing air along the walls, makes it possible to obtain a burner that supplies a combustion gas temperature of up to 1700° C., while at the same time being easily cooled and occupying very little space so that it can, for example, be housed in an existing installation used in the manufacture of rock wool or glass wool.

Abstract: A system for cooling a fuel injector extending into a combustion gas flow field includes a liner that defines a combustion gas flow path within a combustor and a fuel injector that extends through the liner. The fuel injector includes a main body having an annular first portion that extends outward from the liner and a second portion that extends from the first portion inward into the combustion gas flow path. The main body defines a cooling channel that is fully circumscribed within the main body and that extends between the first portion and the second portion. A cooling air inlet is defined within the annular first portion and is in fluid communication with the cooling channel. A cooling air outlet is in fluid communication with the cooling channel and is defined within the second portion downstream from the cooling air inlet.

Abstract: A system includes a solids combining system with a solids mixing section having a mixing chamber and a plurality of solids inlets configured to supply one or more solids into the mixing chamber. The system also includes a solids breakup section having a plurality of fluid inlets configured to supply one or more fluids into the mixing chamber and a solids flow control section having a converging-diverging passage downstream of the mixing chamber.

Abstract: The invention discloses a method for operating a gas turbine with sequential combustion, which gas turbine includes a compressor, a first combustor with a first combustion chamber and first burners, which receives compressed air from the compressor, a second combustor with a second combustion chamber and second burners, which receives hot gas from the first combustor with a predetermined second combustor inlet temperature, and a turbine, which receives hot gas from the second combustor. The CO emission for part-load operation is reduced by reducing the second combustor inlet temperature for base-load operation of the gas turbine, and increasing the second combustor inlet temperature when decreasing the gas turbine load (RLGT) from base-load to part-load.

Abstract: A wake reducing structure includes a combustor liner having an inner surface and an outer surface, the inner surface defining a combustor chamber. Also included is an airflow path located along the outer surface of the combustor liner. Further included is a wake generating component disposed in the airflow path and proximate the combustor liner, wherein the wake generating component generates a wake region located downstream of the wake generating component. Yet further included is a wake generating component boss operatively coupled to the combustor liner and disposed within a combustor liner aperture. Also included is a cooling channel extending through the wake generating component boss, the cooling channel having an air inlet on an upstream region of the wake generating component boss and an air outlet on a downstream region of the wake generating component boss, the cooling channel configured to supply air to the wake region.

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 combustor (10) includes a cap (16), a liner (20), a transition piece (24), and a combustion chamber (22) located downstream from the cap (16) and defined by the cap and liner. A secondary nozzle (40) circumferentially arranged around the liner (20) or transition piece (24) includes a center body, a fluid passage through the center body, a shroud circumferentially surrounding the center body, and an annular passage between the center body and the shroud. A method for supplying fuel to a combustor (10) includes flowing fuel through a primary nozzle radially disposed in a breech end of the combustor and flowing fuel through a secondary nozzle (40) circumferentially arranged around and passing through at least one of a liner (20) or a transition piece. The secondary nozzle (40) includes a center body, a fluid passage through the center body, a shroud circumferentially surrounding at least a portion of the center body (44), and an annular passage between the center body and the shroud.

Abstract: A combustion liner for a gas turbine combustor includes an annular main body having a forward end axially separated from an aft end, and a transitional intersection defined between the forward end and the aft end. The main body extends continuously from the forward end to the aft end. A plurality of fuel injector passages extend radially through the main body upstream from the transitional intersection. The main body comprises a conical section having a circular cross section that diverges between the forward end and the transitional intersection, and a transition section having a non-circular cross section that extends from the transitional intersection to the aft end of the main body.

Abstract: A slinger combustor has an annular combustor shell defining a combustion chamber having a radially inner fuel inlet for receiving a spray of fuel centrifuged by a fuel slinger. The combustion chamber has a fuel atomization zone extending radially outwardly from the fuel inlet and merging into a radially outwardly flaring expansion zone leading to a combustion zone. A plurality of nozzle air inlets are defined in the fuel atomization zone of the combustor shell. The nozzle air inlets have a nozzle axis intersecting the stream of fuel and a tangential component in a direction of rotation of the fuel slinger. A plurality of dilution holes are defined in the combustor shell and have a dilution axis intersecting the combustion zone. The dilution axis of at least some of the dilution holes has a tangential component opposite to the direction of rotation of the fuel slinger.

Abstract: A swirler assembly for a gas turbine engine includes an outer annular injector which at least partially surrounds an inner injector. In sonic embodiments, a combustor section for a gas turbine engine comprises an inner injector which defines an axis, an outer annular injector which surrounds said inner injector, and a combustor vane along said axis.

Abstract: A combustor assembly defining a main combustion zone where fuel and air are burned to create hot combustion products includes a liner, a transition duct, and a transition inlet cone. The liner defines an interior volume including a first portion of the main combustion zone, and has an inlet and an outlet spaced from the inlet in an axial direction. The transition duct includes an inlet section and an outlet section that discharges gases to a turbine section. The inlet section is adjacent to the outlet of the liner and defines a second portion of the main combustion zone. The transition inlet cone is affixed to the transition duct and includes a frusto-conical portion extending axially and radially inwardly into the main combustion zone. The transition inlet cone deflects combustion products that are flowing in a radially outer portion of the main combustion zone toward a combustor assembly central axis.

Abstract: A fuel nozzle is provided. The fuel nozzle includes a nozzle body, a plurality of swirler vanes, and at least one outlet. The nozzle body includes a back plate, a front plate, and a mixing zone defined therebetween. The back plate includes at least one inlet defined therein and the front plate includes at least one discharge defined therein. The plurality of swirler vanes are positioned between the back plate and the front plate and spaced circumferentially about the mixing zone. Each of the plurality of swirler vanes direct air obliquely into the mixing zone. The at least one outlet is defined within at least one of the nozzle body and the plurality of swirler vanes, the at least one outlet configured to inject fuel into said mixing zone.

Abstract: A combustor for a gas turbine engine comprises an inner annular liner and an outer annular liner. First and second combustion stages are defined between the liners. Each combustion stage has a plurality of fuel injection bores distributed in a liner wall defining the respective stage. A lobed mixer extends into the combustor, the lobed mixer arranged to receive combustion gases from each combustion stage for mixing flows of said combustion gases.

Abstract: A reheat burner (1) includes a channel (2) with a lance (3) protruding thereinto to inject a fuel over an injection plane (4) perpendicular to a channel longitudinal axis (15). The channel (2) and lance (3) define a vortex generation zone (6) upstream of the injection plane (4) and a mixing zone (9) downstream of the injection plane (4) in the hot gas (G) direction. The mixing zone (9) includes a high speed area (16) with a constant cross section, and a diffusion area (17) with a flared cross section downstream of the high speed area (16) in the hot gas (G) direction.

Abstract: A turbomachine combustor assembly includes a combustor including a combustor body extending from a head end to a second end through a combustion zone. At least one injector is configured and disposed to introduce a first combustible fluid into the combustion zone at the head end of the combustor body. At least one late lean injector (LLI) is configured and disposed to introduce a second combustible fluid downstream of the first combustible fluid. An external fluid delivery system is fluidically connected to the at least one LLI. The external fluid delivery system includes at least one combustible fluid delivery conduit that extends from a first end fluidically connected to the LLI to a second end external to the turbomachine combustor assembly. The second end is configured and disposed to deliver the second combustible mixture to the LLI.

Abstract: A combustor includes an end cap having an upstream surface axially separated from a downstream surface. A cap shield circumferentially surrounds the upstream and downstream surfaces, tubes extend from the upstream surface through the downstream, and a plenum is inside the end cap. A first baffle extends radially across the plenum toward the cap shield, and a plate extends radially inside the plenum between the first baffle and the upstream surface. A method for supplying fuel to a combustor includes flowing a working fluid through tubes, flowing a fuel into a plenum between upstream and downstream surfaces, radially distributing the fuel along a first baffle, and axially flowing the fuel across a plate that extends radially inside the plenum.

Abstract: A fuel nozzle assembly for use in a combustor apparatus of a gas turbine engine. An outer housing of the fuel nozzle assembly includes an inner volume and provides a direct structural connection between a duct structure and a fuel manifold. The duct structure defines a flow passage for combustion gases flowing within the combustor apparatus. The fuel manifold defines a fuel supply channel therein in fluid communication with a source of fuel. A fuel injector of the fuel nozzle assembly is provided in the inner volume of the outer housing and defines a fuel passage therein. The fuel passage is in fluid communication with the fuel supply channel of the fuel manifold for distributing the fuel from the fuel supply channel into the flow passage of the duct structure.

Abstract: An exemplary burner of a gas turbine includes a tubular body with an inlet for an entrance of an air flow, downstream of inlet vortex generators, and a lance projecting into the tubular body and having a terminal portion extending along a longitudinal axis of the burner which is provided with nozzle groups for injecting fuel into the tubular body. The nozzle groups can lay in an injection plane perpendicular to the axis of the terminal portion of the lance. Downstream of the lance, the burner has an outlet. A ratio x/L between an axial distance x between the side trailing edge of the vortex generator and the injection plane, and the length L of the tubular body can be less than approximately 0.1052.

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 late lean fuel injection nozzle for a gas turbine includes a first outer air supply tube having a relatively large inner diameter and an outlet at a distal end thereof. The first outer air supply tube is adapted to supply air to a combustion chamber, and at least one fuel injection tube having a relatively smaller diameter enters a distal end portion of the first outer air supply tube and extends within the first outer air supply tube substantially to the outlet.

Abstract: Apparatus and methods for combustion of fuel includes, in some embodiments, a fuel nozzle which injects fuel into a combustion channel of a wave rotor combustor or a pulse detonation combustor. In some embodiments the combustion process includes a backward-propagating detonation wave within a substantially closed channel which compresses discrete quantities of combustible and noncombustible mixture. Yet other embodiments include a precombustion chamber integrated into the wave rotor, the outlet stator or both.

Abstract: A fuel injector having an injector body, a mixing circuit, and at least one injector is provided. The injector body has a plurality of manifolds, an inlet, and an outlet. The manifolds are configured for receiving fuel, and the inlet is configured for receiving air. The mixing circuit is positioned within the injector body. The mixing circuit is configured for receiving fuel from at least one of the manifolds, and air from the inlet to create an air-fuel mixture that exits the outlet. The least one fuel injector is positioned radially outwardly from the mixing circuit. The at least one injector receives fuel from at least one of the plurality of manifolds and injects fuel to the outlet.

Abstract: The present application provides a combustor. The combustor may include an air flow path with a flow of air therein. A flow obstruction may be positioned within the air flow path and cause a wake or a recirculation zone downstream thereof. A number of fuel injectors may be positioned downstream of the flow obstruction. The fuel injectors may inject a flow of fuel into the air flow path such that the flows of fuel and air in the wake or the recirculation zone do not exceed a flammability limit.

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 for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a plurality of fuel injectors circumferentially arranged around the combustion chamber. The plurality of fuel injectors provide fluid communication into the combustion chamber, and a separate cap for each fuel injector defines a separate volume around a different fuel injector outside of the combustion chamber.

Abstract: A multi-zone combustor is provided and includes a pre-mixer configured to output a first mixture to a primary zone of a combustor section and a stepped center body disposable in an annulus defined within the pre-mixer. The stepped center body includes an outer body configured to output at a first radial and axial step a second mixture to a secondary zone of the combustor section and an inner body disposable in an annulus defined within the outer body and configured to output at a second radial and axial step a third mixture to a tertiary zone of the combustor section.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber disposed downstream from the fuel nozzle, an inner flow sleeve that circumferentially surrounds the combustion chamber and a plurality of injectors circumferentially arranged around the inner flow sleeve. The plurality of injectors provide for fluid communication through the inner flow sleeve and into the combustion chamber downstream from the fuel nozzle. The system further includes an outer air shield that defines an injection air plenum that surrounds the plurality of injectors. An inlet passage extends through the outer air shield to define a flow path into the injection air plenum. An outer sleeve is slidingly engaged with the outer air shield. The outer sleeve has a first position that restricts flow through the inlet passage and a second position that increases flow through the inlet passage.

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 reheat combustor for a gas turbine engine includes a fuel/gas mixer for mixing fuel, air and combustion gases produced by a primary combustor and expanded through a high pressure turbine. Fuel injectors inject fuel into the mixer together with spent cooling air previously used for convectively cooling the reheat combustor. The fuel mixture is burnt in an annular reheat combustion chamber prior to expansion through low pressure turbine inlet guide vanes. The fuel/gas mixer and optionally the combustion chamber define cooling paths through which cooling air flows to convectively cool their walls. The fuel injectors are also convectively cooled by the cooling air after it has passed through the fuel/gas mixer cooling paths. The low pressure turbine inlet guide vanes may also define convective cooling paths in series with the combustion chamber cooling paths.

Abstract: A late lean injection (LLI) manifold is provided and includes a central nozzle and first and second side nozzles positioned at circumferential locations defined around a vessel, a connector, a first leg, to which the connector is connected, formed to define a tube extending from the central nozzle to the first side nozzle such that fuel is communicable between the connector, the central nozzle and the first side nozzle and a second leg. The second leg is formed to define a tube extending from the central nozzle to the second side nozzle such that the fuel is communicable between the central nozzle and the second side nozzle.