Abstract: A gas turbine combustor includes: multiple gaseous fuel nozzles to inject gaseous fuel; a manifold to distribute gaseous fuel to the multiple gaseous fuel nozzles; a burner flange to secure a burner to an end cover; and a burner body to connect the manifold and the burner flange. Gaseous fuel passages are provided in the end cover, the burner flange, and the burner body. The gaseous fuel passage communicates with the manifold. A second manifold is provided in the burner flange. The gaseous fuel passage is provided in the burner body to make the second manifold communicate with the first mentioned manifold.

Abstract: An integrated plate is provided for use with a combustor including a casing, a fuel plenum extending circumferentially about the casing, and a fuel nozzle extending axially through the casing. The integrated plate includes a plurality of fuel injection pegs that extend radially between the fuel plenum and the fuel nozzle.

Abstract: A nozzle formed of one piece for a jet engine includes a mixing tube, a fuel conduit integrally formed with the mixing tube, and an opening through the fuel conduit and directed radially into the mixing tube.

Abstract: In one embodiment, a system includes a fuel nozzle that includes a fuel injector that includes a fuel port and a premixer tube. The premixer tube includes a wall disposed about a central passage, multiple air ports extending through the wall into the central passage, and a catalytic region. The catalytic region includes a catalyst, disposed inside the wall along the central passage, configured to increase a reaction of fuel and air.

Abstract: In one embodiment, a system includes a variable geometry resonator configured to couple to a fluid path upstream from a combustor of a turbine engine. The variable geometry resonator is configured to dampen pressure oscillations in the fluid path and the combustor.

Abstract: A swirler for fuel injection in a gas turbine engine includes a frustoconical swirler body. A first and a second air flow path direct air in generally opposed circumferential directions into the swirler. These air paths intermix and create turbulence. As this turbulence encounters fuel droplets, the fuel is atomized, and uniformly distributed within the air flow. A shear layer is created adjacent an inner surface of the swirler body. In a separate feature, a third air flow path is directed into the air.

Abstract: A fuel-air premixer for a combustor of a gas turbine engine includes a central passage disposed along an axis and operable to communicate a first airflow and an outer annular passage about the axis operable to communicate a second airflow. The fuel-air premixer further includes an inner annular passage about the axis and between the central passage and the outer annular passage for communicating fuel flow. The inner annular passage including an inner exit angled for directing fuel flow toward the axis into a mixer passage downstream of the outer and inner exits for mixing the fuel flow with the first and second airflows.

Abstract: A combustor basket (20) is provided comprising a generally tubular wall (22) defining a cavity (24) and a liner (26) of a ceramic matrix composite material disposed within the cavity (24), wherein an interference fit exists between an outer surface (34) of the liner (26) located in the cavity (24) and an inner surface (28) of the generally tubular wall (22). Advantageously, the novel combustor basket (20) greatly reduces the amount of air or other medium needed to cool the combustor basket body and increases the maximum allowable flame temperature that the combustor basket (20) can withstand. A method of manufacturing the combustor basket (20) is also provided.

Abstract: A combustion chamber including a first fuel injector and a second fuel injector, the first and second fuel injectors being arranged to inject fuel into a mainstream flow of air with the second fuel injector arranged downstream of the first fuel injector. A method of mixing fuel and air in a combustion chamber, including injecting fuel into a mainstream flow of air with a first fuel injector; injecting fuel into the mainstream flow of air with a second fuel injector, which is arranged downstream of the first fuel injector; injecting fuel into the mainstream flow with the first fuel injector such that the resulting mixture between the first and second fuel injectors has an equivalence ratio less than the lean flame stability limit; and injecting fuel into the mainstream flow with the second fuel injector such that a combustion zone is provided downstream of the second fuel injector.

Abstract: Disclosed is a gas turbine combustor equipped with a burner constructed to fire a plurality of combustors at the same time at a fuel/air ratio suitable for gas turbine ignition. A plurality of supports mounted for fixing an air injection hole plate to the combustor are provided at substantially the same phase position as that of cross fire tubes, and combustion air allocations are adjusted. Thus, flame propagation during gas turbine ignition is accelerated and all combustor cans are fired. In addition, porous plates are disposed downstream of the supports and the combustion air allocations are readjusted. Furthermore, air injection holes proximate to the cross fire tubes are particularly reduced in diameter to further adjust the combustion air allocations.

Abstract: Methods and systems are provided for premixing combustion fuel and air within gas turbines. In one embodiment, a combustor includes an upstream mixing panel configured to direct compressed air and combustion fuel through a premixing zone to form a fuel-air mixture. The combustor also includes a downstream mixing panel configured to mix additional combustion fuel with the fuel-air mixture to form a combustion mixture.

Abstract: A late lean injection sleeve assembly allows the injection of fuel at the aft end of a gas turbine liner, before the transition piece, into the combustion gases downstream of a turbine combustor's fuel nozzles. The late lean injection enables fuel injection downstream of the fuel nozzles to create a secondary/tertiary (with quaternary injection upstream of the fuel nozzles) combustion zone while reducing/eliminating the risk of fuel leaking into the combustor discharge case. The fuel is delivered by the flow sleeve into one or more nozzles that mix the fuel with CDC air before injecting it into the combustor's liner.

Abstract: The present application provides a combustor for combusting a number of flows of air and a number of flows of fuel. The combustor may include a central swirler for producing a high swirl quench air flow, a number of trapped vortex cavities surrounding the central swirler for producing a flow of combustion gases, a brief severe quench zone downstream of the trapped vortex cavities to quench the flow of combustion gases between an outer quench air flow and the high swirl quench air flow, and an expansion zone downstream of the brief severe quench zone.

Abstract: An improved mixing tube design and fuel nozzle that allows for a more uniform and thorough mixing of fuel and air being fed to the combustor of a gas turbine engine, wherein each of a plurality of mixing tubes comprises a pair of concentric hollow cylinders that define a ring-like, annular path for the flow of fuel between the two hollow cylinders in each mixing tube, a plurality of air injection slots formed in the concentric hollow cylinders defining corresponding air flow paths from the outside into the interior of each mixing tube, and one or more fuel injection ports formed in selected ones of the plurality of air injection slots that allow for the flow of fuel from the annular path formed by the hollow cylinders into the air flow path, resulting in significantly better mixing and improved thermodynamic behavior of the fuel and air mixture downstream of the nozzle and upstream of the combustor.

Abstract: Disclosed is a premixer for a combustor including an annular outer shell and an annular inner shell. The inner shell defines an inner flow channel inside of the inner shell and is located to define an outer flow channel between the outer shell and the inner shell. A fuel discharge annulus is located between the outer flow channel and the inner flow channel and is configured to inject a fuel flow into a mixing area in a direction substantially parallel to an outer airflow through the outer flow channel and an inner flow through the inner flow channel. Further disclosed are a combustor including a plurality of premixers and a method of premixing air and fuel in a combustor.

Abstract: A method for the combustion of hydrogen-rich, gaseous fuels in combustion air in a burner of a gas turbine includes injecting the hydrogen-rich, gaseous fuel at least partially isokinetically with respect to the combustion air such that the partially hydrogen-rich, gaseous fuel is injected at least partially in the same direction and at least partially at the same velocity as the combustion air.

Abstract: A fuel injector is provided and includes an outer body, a converging tubular body including a converging section, defining a converging annular passage therein and being disposed within the outer body to define an outer annular passage between an interior surface of the outer body and an exterior surface of the converging tubular body, the annular passages each being receptive of a fluid at respective inlets thereof such that the fluid is directed to flow toward respective outlets thereof and a fuel line to deliver fuel through at least the converging tubular body at the converging section to the converging annular passage in a substantially tangential direction relative to a circumferential curvature of the converging tubular body.

Abstract: A system for reducing combustion dynamics in a combustor includes an end cap having an upstream surface axially separated from a downstream surface, and tube bundles extend from the upstream surface through the downstream surface. A divider inside a tube bundle defines a diluent passage that extends axially through the downstream surface, and a diluent supply in fluid communication with the divider provides diluent flow to the diluent passage. A method for reducing combustion dynamics in a combustor includes flowing a fuel through tube bundles, flowing a diluent through a diluent passage inside a tube bundle, wherein the diluent passage extends axially through at least a portion of the end cap into a combustion chamber, and forming a diluent barrier in the combustion chamber between the tube bundle and at least one other adjacent tube bundle.

Abstract: A combustor for a gas turbine engine has a head end portion that carries at least one fuel/air nozzle. Each fuel/air nozzle includes a premixed pilot nozzle having premix conduits that are configured with concentric axes that direct the fuel/air mixture axially from the premixed pilot nozzle. The premixed pilot nozzle can include an annular channel disposed radially outwardly from the premix and including air jets that direct air radially outwardly from the premix conduits.

Abstract: The disclosure relates to a burner for a single combustion chamber or first combustion chamber of a gas turbine, with an injection device for the introduction of at least one gaseous and/or liquid fuel into the burner, wherein the injection device has at least one body which is arranged in the burner with at least one nozzle for introducing the at least one fuel into the burner, wherein the at least one body is located in a first section of the burner with a first cross-sectional area at a leading edge of the at least one body with reference to a main flow direction prevailing in the burner, wherein downstream of said body a mixing zone is located with a second cross-sectional area, and at and/or downstream of said body the cross-sectional area is reduced, such that the first cross-sectional area is larger than the second cross-sectional area.

Abstract: A combustor section of a gas turbine includes a combustor liner, a sleeve and a fuel-air mixing tube. The combustor liner defines a combustion chamber. The sleeve surrounds the combustor liner. The combustor liner and the sleeve define an annular flow space. The fuel-air mixing tube is configured to channel a mixture of fuel and air and includes an inlet and an outlet. The inlet is in fluid communication with an exterior of the sleeve, and the outlet is in fluid communication with the combustion chamber. The combustor casing encloses the combustor section upstream relative to the inlet of the mixing tube and extends downstream. The sleeve and the combustor casing define a discharge air space. The discharge space is in fluid communication with the mixing tube. The fuel supplying device is located exteriorly of the combustor casing and is configured to inject fuel into the mixing tube.

Abstract: A flowsleeve of a turbomachine component is provided. The flowsleeve includes an annular body including an upstream casing and a downstream casing. The upstream casing defines a fuel feed, and the downstream casing defines an airway opening, and a premixing passage. The premixing passage is fluidly coupled to the fuel feed and the airway opening and has a passage interior in which fuel and air receivable from the fuel feed and the airway opening, respectively, are combinable to form a fuel and air mixture.

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. A vane angle of the plurality of vanes is adjustable.

Abstract: A burner arrangement includes a tubular mixing tube extending along an axis and configured to mix fuel and air, the mixing tube opening into a combustion chamber in a direction of flow; and a film air ring disposed in the mixing tube and arranged concentrically to the axis, the film air ring configured to prevent a backfiring of the fuel disposed in the mixing tube and to generate or reinforce an air film near the wall, wherein the film air ring has at least one annular gap concentric to the axis and configured to inject air.

Abstract: Embodiments of this invention provide a premixed pilot assembly for use with a fuel nozzle for a turbine. The level of nitrogen oxide (NOx) emitted by the pilot is reduced by mixing the fuel and air fast, at the end of the pilot nozzle, thereby avoiding significant zones of rich fuel and air mixtures. The air is mixed with the fuel through the use of openings in a first cylinder and a second cylinder, one of which carries air and one of which carries pilot fuel. The openings can be configured as necessary to obtain a desired effect on the pilot.

Abstract: An air/fuel premixer comprising a peripheral wall defining a mixing chamber, a nozzle disposed at least partially within the peripheral wall comprising an outer annular wall spaced from the peripheral wall so as to define an outer air passage between the peripheral wall and the outer annular wall, an inner annular wall disposed at least partially within and spaced from the outer annular wall, so as to define an inner air passage, and at least one fuel gas annulus between the outer annular wall and the inner annular wall, the at least one fuel gas annulus defining at least one fuel gas passage, at least one air inlet for introducing air through the inner air passage and the outer air passage to the mixing chamber, and at least one fuel inlet for injecting fuel through the fuel gas passage to the mixing chamber to form an air/fuel mixture.

Abstract: A fuel nozzle for use with a turbine engine is described herein. The fuel nozzle includes a housing that is coupled to a combustor liner defining a combustion chamber. The housing includes an endwall that at least partially defines the combustion chamber. A plurality of mixing tubes extends through the housing for channeling fuel to the combustion chamber. Each mixing tube of the plurality of mixing tubes includes an inner surface that extends between an inlet portion and an outlet portion. The outlet portion is oriented adjacent the housing endwall. At least one of the plurality of mixing tubes includes a plurality of projections that extend outwardly from the outlet portion. Adjacent projections are spaced a circumferential distance apart such that a groove is defined between each pair of circumferentially-apart projections to facilitate enhanced mixing of fuel in the combustion chamber.

Abstract: Disclosed is a gas turbine fuel injector and swirler assembly, including: a delivery tube structure arranged on a central axis of the fuel injector and swirler assembly, a first fuel supply channel arranged in the delivery tube structure, a shroud surrounding the delivery tube structure, swirl vanes arranged between the delivery tube structure and the shroud, a radial passage in each swirl vane, communicating with the first fuel supply channel, a set of apertures open between the radial passage and the exterior surface of said each swirl vane, wherein a second fuel supply channel is arranged in the delivery tube structure extending to a downstream end of the delivery tube structure and a mixer with lobes for fuel injection is arranged at the downstream end. Further disclosed is an assembly method for assembling a fuel injector and swirler assembly.

Abstract: A combined flow straightener and mixer is disclosed as well as a burner for a combustion chamber of a gas turbine having such a mixing device. At least two streamlined bodies are arranged in a structure comprising the side walls of the mixer. The leading edge area of each streamlined body has a profile, which is oriented parallel to a main flow direction prevailing at the leading edge position, and with reference to a central plane of the streamlined bodies, the trailing edges are provided with at least two lobes in opposite transverse directions. The periodic deflections forming the lobes from two adjacent streamlined bodies are out of phase.

Abstract: Disclosed is a triple swirl gas turbine combustor using various fuels such as coal gas, DiMethyl Ether, and waste gas generated in an ironworks in a gas turbine. The triple swirl gas turbine combustor combusts three different fuels simultaneously or individually and various fuels such as LCV gas and HCV gas so that fuel flexibility can be improved. Swirl generated in a second swirler of the triple swirler is reversely jetted to increase a mixing degree of a fuel-air mixture and an intensity of a turbulent flow so that combustion efficiency can be increased, harmful exhaust gas can be reduced and vibration can be reduced.

Abstract: A highly-reliable combustor is provided that allows flash back of flame into a premixer to be suppressed. The combustor has a mixing chamber forming member 110 that forms a mixing chamber thereinside. The mixing chamber includes a first mixing chamber 200 broadening toward a downstream side. The member 110 includes air introduction holes 202, 203, 204 formed in a plurality of rows in an axial direction, with the air introduction holes being arranged plurally in a circumferential direction of the mixing chamber. The member 110 includes a fuel ejection hole 206 provided in a wall surface which forms the air introduction hole. The air introduction holes 202, 203, 204 are circumferentially eccentrically installed. The air introduction holes 202 located in the most upstream row are more inclined toward the downstream side than the air introduction holes 203, 204 located in the rows other than the most upstream row.

Abstract: A combustor includes an end cover and a combustion chamber downstream of the end cover. The combustor further includes nozzles disposed radially in the end cover and a shroud surrounding at least one of the nozzles and extending downstream into the combustion chamber. The shroud includes an inner wall surface and an outer wall surface. A method for operating a combustor includes flowing compressed working fluid through nozzles into a combustion chamber, flowing fuel through each nozzle in a first subset of the nozzles into the combustion chamber, and igniting the fuel from each nozzle in the first subset of nozzles in the combustion chamber. In addition, the method includes extending into the combustion chamber a separate shroud around each nozzle in a second subset of the nozzles and isolating fuel to each nozzle in the second subset of nozzles.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique fuel injection system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and fuel injection systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A combustor (1) for a gas turbine engine, particularly for a gas turbine having sequential combustion, includes a combustor wall (4) defining a mixing region (5) and a combustion region (6). The mixing region (5) has at least one first inlet (2) for introducing combustion air into the mixing region (5) and at least one second inlet for introducing fuel into the mixing region (5), the combustion region (6) extending downstream of the mixing region. The mixing region (5) crosses over to the combustion region (6) in a transition region (14). A baffle (9) extends from the transition region (14) generally in the downstream direction (15), forming at least one space (10) between the combustor wall (4) and the baffle (9).

Abstract: In one embodiment, a system includes a fuel nozzle that includes a fuel injector configured to output a fuel flow and a premixer tube disposed about the fuel flow output from the fuel injector. The premixer tube includes a perforated portion and a non-perforated portion, and the non-perforated portion is downstream of the perforated portion.

Abstract: An arrangement for preparation of a fuel for combustion including a burner is provided. The arrangement includes a combustion chamber associated with the burner and in which combustion of a fuel is to take place, as well as means for supplying liquid fuel to the arrangement through an internal passage in the burner for the combustion, solid portions of the burner body being heated by the combustion, wherein the internal passage is located inside the solid portions of the burner body for receiving heat energy evaporating the fuel from these body portions, and that it includes means for conveying vaporized fuel to the combustion chamber to take part in the combustion.

Abstract: A low-emissions high-pressure multi-fuel burner includes a fuel inlet, for receiving a fuel, an oxidizer inlet, for receiving an oxidizer gas, an injector plate, having a plurality of nozzles that are aligned with premix face of the injector plate, the plurality of nozzles in communication with the fuel and oxidizer inlets and each nozzle providing flow for one of the fuel and the oxidizer gas and an impingement-cooled face, parallel to the premix face of the injector plate and forming a micro-premix chamber between the impingement-cooled face and the in injector face. The fuel and the oxidizer gas are mixed in the micro-premix chamber through impingement-enhanced mixing of flows of the fuel and the oxidizer gas. The burner can be used for low-emissions fuel-lean fully-premixed, or fuel-rich fully-premixed hydrogen-air combustion, or for combustion with other gases such as methane or other hydrocarbons, or even liquid fuels.

Abstract: A fuel injection head for a fuel nozzle used in a gas turbine combustor includes a substantially hollow body formed with an upstream end face, a downstream end face and a peripheral wall extending therebetween. A plurality of pre-mix tubes or passages extend axially through the hollow body with inlets at the upstream end face and outlets at the downstream end face. An exterior surface of the downstream end face is formed with three-dimensional surface features that increase a total surface area of the exterior surface as compared to a substantially flat, planar downstream end face.

Abstract: The present application provides a fuel nozzle for use in a gas turbine. The fuel nozzle may include a mounting flange, a number of premixers attached to each other, and a number of gas pathways extending from the mounting flange to the number of premixers.

Abstract: A fuel-air mixer includes an annular shroud, inner and outer counter-rotating swirlers, a hub separating the inner and outer swirlers, a center body extending axially along the annular shroud, a fuel shroud disposed radially outwardly from the outer swirler circumferentially around the annular shroud, and a radial swirler disposed downstream of the inner and outer swirlers, the radial swirler being configured to allow an independent radial rotation of a second gas stream entering the annular shroud from a region outside a wall of the annular shroud separately from a stream flowing through the inner and outer swirlers, the radially rotating second gas stream entering the annular shroud at a region adjacent an outer wall of the annular shroud.

Abstract: A liner cap assembly is disclosed for use in a gas turbine engine combustor. The assembly includes an outer ring that extends along an axis. Multiple struts are circumferentially arranged about an inner diameter of the outer ring and extend radially inwardly therefrom. A plate is supported by and axially aligned with the struts. The plate includes multiple circumferential openings that support a collar and a premix tube at each of the openings. The plate is arranged between leading and trailing edges of the struts to provide a stiffened liner cap assembly that is robust and resistant to the vibrations typically found in dry low NOx systems.

Abstract: A combustor for a turbine engine is disclosed. The combustor may have a can-like dilution zone, a primary combustion zone, and a secondary combustion zone. The primary combustion zone may be disposed radially about the can-like dilution zone. The secondary combustion zone may be disposed at an end of the can-like dilution zone to fluidly communicate the primary combustion zone and the can-like dilution zone.

Abstract: A burner and a method for operating a burner are provided. The burner includes a channel having a mixing zone and having a feed for an oxidation means, particularly an air feed, and at least one fuel feed for injecting fuel, wherein a separating means which divides the channel over a wide range of the channel into at least two separated channels, namely a first channel and a second channel, is provided in the channel. The method for operating a burner having a channel includes a mixing zone into which an oxidation mass flow and fuel are injected, wherein two substantially separate flow paths are formed by means of a separating means in the channel and the at least two separated first and second channels, formed by the separating means.

Abstract: A fuel nozzle assembly for a gas turbine engine includes a flange and a pre-mix tube. The flange includes a first end that is configured to couple to an end cover of the combustor, and a second end that is opposite said first end. The pre-mix tube is coupled at a first end to the flange second end. The flange and the pre-mix tube are fabricated to operate at a natural frequency that is different from an operating frequency of the gas turbine engine.

Abstract: One embodiment is a unique gas turbine engine combustion chamber including primary and secondary burning zones. Other embodiments include unique gas turbine engine apparatuses, systems, methods, and combinations.

Abstract: A premixing device includes an air inlet, at least one fuel inlet slot having a wall profile configured to form a fuel boundary layer along a portion of a wall of the premixing device, a mixing chamber, and at least one diverging fuel injection slot jet disposed inside the at least one fuel inlet slot, the slot jet being configured to create a flow separation region in a diverging portion thereof to generate mixing turbulence at an outlet of the slot jet to aerodynamically enhance a mixing of the fuel from the boundary layer with compressed air without causing a boundary layer flow separation and a flame holding in the mixing chamber. Low-emission combustors, gas turbine combustors, methods for premixing a fuel and an oxidizer in a combustion system, a gas turbine, and a gas to liquid system using the premixing device are also disclosed.

Abstract: An air/fuel premixer comprising a peripheral wall defining a mixing chamber, a nozzle disposed at least partially within the peripheral wall comprising an outer annular wall spaced from the peripheral wall so as to define an outer air passage between the peripheral wall and the outer annular wall, an inner annular wall disposed at least partially within and spaced from the outer annular wall, so as to define an inner air passage, and at least one fuel gas annulus between the outer annular wall and the inner annular wall, the at least one fuel gas annulus defining at least one fuel gas passage, at least one air inlet for introducing air through the inner air passage and the outer air passage to the mixing chamber, and at least one fuel inlet for injecting fuel through the fuel gas passage to the mixing chamber to form an air/fuel mixture.

Abstract: Disclosed is a premixer for a combustor including an annular outer shell and an annular inner shell. The inner shell defines an inner flow channel inside of the inner shell and is located to define an outer flow channel between the outer shell and the inner shell. A fuel discharge annulus is located between the outer flow channel and the inner flow channel and is configured to inject a fuel flow into a mixing area in a direction substantially parallel to an outer airflow through the outer flow channel and an inner flow through the inner flow channel. Further disclosed are a combustor including a plurality of premixers and a method of premixing air and fuel in a combustor.

Abstract: A mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer, and a fuel manifold. The pilot mixer includes: an annular pilot housing having a hollow interior, a primary fuel injector mounted in the pilot housing, a plurality of axial swirlers positioned upstream from the primary fuel injector, and a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity, a plurality of fuel injection ports, and at least one swirler positioned upstream from the plurality of fuel injection ports. The fuel manifold is in flow communication with the plurality of secondary fuel injection ports in the pilot mixer and the plurality of fuel injection ports in the main mixer.

Abstract: A gas turbine system includes a fuel reformer comprising: a fuel inlet; an oxygen inlet; a pre-mixing zone configured to mix the fuel and the oxygen in a pre-mixing device to form a gaseous pre-mix; wherein the pre-mixing device comprises a flow conditioning device configured to pre-condition the fuel stream, wherein the flow conditioning device is disposed upstream of the oxygen inlet; a diffuser disposed downstream of the flow conditioning device; a catalytic partial oxidation zone disposed downstream of the diffuser, wherein the catalytic partial oxidation zone comprises a catalyst composition configured to react the fuel and the oxygen to generate a syngas. The generated syngas is then mixed with rest of the fuel to form a hydrogen-enriched fuel mixture, which is then sent to the combustion chamber of a gas turbine to reduce the NOx emission and extend the lean blow out limit.