Abstract: An apparatus and method for lean/rich combustion in a gas turbine engine (10), which includes a combustor (12), a transition (14) and a combustor extender (16) that is positioned between the combustor (12) and the transition (14) to connect the combustor (12) to the transition (14). Openings (18) are formed along an outer surface (20) of the combustor extender (16). The gas turbine (10) also includes a fuel manifold (28) to extend along the outer surface (20) of the combustor extender (16), with fuel nozzles (30) to align with the respective openings (18). A method (200) for axial stage combustion in the gas turbine engine (10) is also presented.

Abstract: A fuel injector for a combustor generally includes an annular outer body having an inlet and an outlet. The outer body at least partially defines an outer flow passage. An inner flow passage extends at least partially through the outer flow passage and a radial swirler is disposed at the inlet of the outer body. The radial swirler includes a first radial passage separated from a second radial passage. The first radial passage has a first plurality of swirler vanes and the second radial passage has a second plurality of swirler vanes. The first radial passage is in fluid communication with the outer flow passage and the second radial passage is in fluid communication with the inner flow passage.

Abstract: A combustor liner with an input end and an output end includes an annular inner wall and an annular outer wall. At least one of the inner wall and outer wall is three-dimensionally contoured. The inner wall and the outer wall form a combustion chamber with the contours creating alternating expanding and constricting regions inside the chamber causing combustion gases to flow in the circumferential and axial directions.

Abstract: A system comprising a fuel nozzle. The fuel nozzle includes a mounting base and an inlet flow conditioner extending directly from the mounting base in a downstream direction. Moreover, the inlet flow conditioner structurally supports the fuel nozzle without a central support member extending directly from the mounting base inside the inlet flow conditioner.

Abstract: A fluid manifold apparatus includes: (a) an array of spaced-apart manifold fittings, each manifold fitting aligned in a predetermined angular orientation. Each manifold fitting includes: (i) a tubular neck; (ii) a pair of spaced-apart tubular arms extending away from a first end of the neck; and (iii) a coupling connected to a second end of the neck; and (b) a plurality of curved tubes, each tube being coupled to one arm of each of two adjacent manifold fittings.

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: 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: Fuel control arrangements for gas turbine engines generally comprise an injector and a fuel control valve. Typically the fuel control valve is controlled in terms of fuel demand through fuel pressure presented to the valve. Fuel demand may vary and in such circumstances stagnation of fuel adjacent to the valve may cause degradation of the fuel and therefore spurious operational performance. By providing a dedicated working fluid, and typically hydraulic, pressure to the valve a variable aperture port can be displaced to alter the fuel flow configuration within the valve. In such circumstances different fuel valve conditions can be generated by altering the available area of aperture in the port to divert or present fuel to the injector between and across first or primary fuel paths and second or pilot fuel paths as required. Thus more flexibility with regard to fuel presentation to the injector is achieved as well as consistency with respect to avoiding fuel degradation as fuel demand varies.

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 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.

Abstract: A method for operating a burner including a burner outlet opening with at least two sectors, each sector is assigned at least one fuel nozzle, is provided. The method is characterized in that fuel is supplied separately to the fuel nozzles of different sectors. Also described is a burner which includes at least two sectors wherein each sector is assigned at least one fuel nozzle. The burner includes at least two separate fuel supply lines and a device for adjusting the fuel mass flow which flows through the respective fuel supply line. The fuel supply lines supply fuel to the fuel nozzles of different sectors. Also described is a gas turbine which is fitted with at least one burner.

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 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 turbomachine including a combustor in which fuel is combustible to produce a working fluid, a turbine section, which is receptive of the working fluid for power generation operations, a transition piece in which additional fuel is combustible, the transition piece being disposed to transport the working fluid from the combustor to the turbine section and a staged combustion system coupled to the combustor and the transition piece. The staged combustion system is configured to blend components of the fuel and the additional fuel in multiple modes.

Abstract: A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel supplied thereto is combustible, a turbine, including rotating turbine blades, into which products of at least the combustion of the first fuel are receivable to power a rotation of the turbine blades, a transition zone, including a second interior in which a second fuel supplied thereto and the products of the combustion of the first fuel are combustible, the transition zone being disposed to fluidly couple the combustor and the turbine to one another, and a plurality of fuel injectors, which are supported by the transition zone and coupled to a fuel circuit, and which are configured to supply the second fuel to the second interior in any one or more of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages.

Abstract: A fuel control method for a gas turbine with a combustor being formed of at least two groups of a pluralities of main nozzles for supplying fuel, and that supplies fuel from the main nozzles of all groups upon ignition of the combustor (S1), and supplies fuel from three main nozzles of a group A during subsequent acceleration of the gas turbine (S3). Because fuel is injected from a small number of the main nozzles during acceleration, the fuel flow rate per one main nozzle is increased, thereby increasing the fuel-air ratio (fuel flow rate/air flow rate) in a combustion region and improving the combustion characteristics. Accordingly, the generation of carbon monoxide and unburned hydrocarbon is reduced, whereby no bypass valve is required and manufacturing costs are reduced.

Abstract: A gas turbine engine is provided and includes a fuel circuit, including multiple fuel circuit branches, a combustor having a first interior in which a first fuel supplied thereto by any one of the multiple fuel circuit branches is combustible, a turbine, a transition zone, including a second interior in which a second fuel supplied thereto by any one of the multiple fuel circuit branches, the second fuel including gas receivable by the fuel circuit from an external source, and the products of the combustion of the first fuel are combustible, the transition zone being disposed to fluidly couple the combustor and the turbine to one another, and a plurality of fuel injectors which supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages.

Abstract: An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of said DMRJ functioning as an aerodynamic choke.

Abstract: A combustor of the present invention includes: a plurality of nozzles which is arranged around the central axis of a combustor and of which each base end is connected to an oil supply tube; a nozzle tube base which supports the plurality of nozzles and includes an oil chamber supplying fuel to the nozzles via an oil inflow portion provided in each oil supply tube; a first guide portion which faces the center of the combustor inside the oil chamber and guides the oil remaining inside the oil chamber to the oil inflow portion; and a second guide portion which faces the base end of the combustor inside the oil chamber and guides the oil remaining inside the oil chamber to the oil inflow portion.

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: In one embodiment, a premixer is provided for use, e.g., with trapped vortex (TVC) combustors, that injects premixed fuel and air directly into the recirculating vortex, in a manner compatible with natural vortex flow in the cavity.

Abstract: An embodiment of the present invention provides a method and system of operating a combustion system that has Lean direct injection (LDI) functionality. The method and system provides a passive cooling system for an injector of the LDI system. The method and system may also provide a means to direct the flow of the fluid exiting the injector of the LDI system.

Abstract: In one embodiment, alternative combustion cavity layouts are provided for practicing fuel staging in a trapped vortex (TVC) combustion apparatus comprising an inlet premixer, for injecting fuel-air mixture into the inlet of the combustion apparatus and one or more vortex premixers, for injecting fuel-air mixture into the recirculating vortex within each of one or more trapped vortex cavities. A plurality of TVC cavities, may, for example, be laid out axially, radially, peripherally, internally, or in combinations of such arrangements. These layouts may be used in conjunction with a fuel staging method whereby the relative proportion of mixture introduced through the inlet and the respective vortex premixers can be varied as a function of operating conditions.

Abstract: The present invention discloses a novel apparatus and way for controlling a velocity of a fuel-air mixture entering a gas turbine combustion system. The apparatus comprises a hemispherical dome assembly which directs a fuel-air mixture along a portion of the outer wall of a combustion liner and turns the fuel-air mixture to enter the combustion liner in a manner coaxial to the combustor axis and radially outward of a pilot fuel nozzle so as to regulate the velocity of the fuel-air mixture.

Abstract: An annular type gas turbine combustor having a plurality of fuel nozzle assemblies (10) on a circumference includes a pilot nozzle unit (12) for spraying a fuel for diffusive combustion from a pilot outer peripheral nozzle (34) into a combustion chamber (8), a main nozzle unit (14) provided so as to surround the pilot nozzle unit (12) for spraying a fuel for premix combustion, and a flow guide (27) disposed on a downstream side of each of the fuel nozzle assemblies (10) and having a sectional area of a passage for air and air-fuel mixture from each of the fuel nozzle assemblies (10), which gradually increase in a downstream direction.

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: A combustor having a combustion chamber is provided with an external flow sleeve and a combustor liner surrounding the combustion chamber. A plurality of flow channels are provided on the combustor liner and a plurality of nozzles are disposed at predetermined locations on the flow channels. The locations of the nozzles are selected to provide different mixing times for fuel injected through the nozzles.

Abstract: A unitary fuel injection manifold for a secondary fuel nozzle improves fuel-air mixing and offers flexibility to alter the mixing profile through adaptability to a variety of number, types, and orientation of discharge outlets to the combustion air mixing space around the secondary fuel nozzle. An aerodynamic surface with reduced extension into the mixing space reduces pressures drop and interference with design airflow. Manifold integrity is enhanced by elimination of fillet welds to mount external pegs.

Abstract: An airblast fuel nozzle assembly (10) comprising a sleeve structure having a series of coaxial sleeves forming an inner-air circuit, an outer-air circuit, a main-fuel-feed circuit, and a pilot-fuel-feed circuit. The pilot-fuel-feed circuit includes a channel (44), and a discharge region (45) with exits (46). The exits (46) have a combined cross-sectional area that is substantially less than the cross-sectional area of the channel (44) upstream of the discharge region (45). In this manner, the pilot-fuel-feed circuit itself can provide a relatively large pressure drop across the channel region (44), and thereby assist in self atomization during ignition stages of engine operation.

Abstract: This invention provides a gas turbine combustor including: a main burner at a head portion of a combustor cylinder; and a pre-mixing type supplemental burner at a downstream portion of the combustor cylinder and extending through a circumferential wall thereof. The supplemental burner includes: an introducing passage configured to deflect a part of the compressed air radially inward, the compressed air flowing from an air passage between the circumferential wall of the combustor cylinder and a housing surrounding the circumferential wall toward the head portion of the combustor cylinder, and introduce the compressed air into the combustor cylinder; and a fuel nozzle configured to supply the fuel from fuel injection holes to the compressed air introduced into the introducing passage to produce a pre-mixed gas in the introducing passage.

Abstract: A method for controlling a gas turbine combustion system includes supplying an ultra low calorific fuel to a combustor of the combustion system through a first fuel circuit, controlling a supply of the ultra low calorific fuel through a second fuel circuit as required to control the volumetric flow of the ultra low calorific fuel through the combustor, and combusting the ultra low calorific fuel in the combustor.

Abstract: A gas-turbine lean combustor includes a combustion chamber (2) and a fuel nozzle (1) which includes a pilot fuel injection (17) and a main fuel injection (18). The main fuel injection (18) includes central recesses (23) for a controlled inhomogeneous fuel injection, the number of said recesses on the circumference ranging from 8 to 40 and said recesses having an angle of inclination ?2 in circumferential direction of 10°??2?60° and an axial angle of inclination ?1 relative to the combustor axis (4) between ?10°??1?90°.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles and a combustion zone downstream of the fuel nozzles. The combustion zone may include a number of steps such that the combustion zone expands in a radial direction downstream of the fuel nozzles.

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 burner for a gas turbine includes a burner housing, and a pilot combustor having a supply module providing pilot fuel and pilot air into a pilot combustion room being enclosed by a pilot combustor housing having a tapered exit with a throat of a defined length into the resulting main flow direction. The throat discharges a concentration of radicals are and heat generated in the pilot combustion room into a main combustion room enclosed by the burner housing. The interior cross section area of the throat deviates from a circle by means of flow guiding elements provided as protrusions with a defined radial height or as recesses with a defined depth extending longitudinally along the direction of the burner axis to give the discharging flow a defined velocity distribution with regard to a circumferential direction.

Abstract: The invention refers to a reheat burner that includes a flow channel for a hot gas flow with a lance arranged along said flow channel, protruding into the flow channel for injecting a fuel over an injection plane perpendicular to a channel longitudinal axis, wherein the channel and lance define a vortex generation zone upstream of the injection plane and a mixing zone downstream of the injection plane in the hot gas flow direction. The mixing zone provides at least one axially region having different cross sectional areas along its longitudinal axis with continuously changing shape, or having non circular cross section areas which change location along its longitudinal axis by continuously rotation around the longitudinal axis.

Abstract: A combustion chamber for a gas turbine is proposed. The combustion chamber has a wall section and a brim element. The wall section has an inlet aperture for injecting a cooling medium into the combustion chamber. The brim element is mounted to an inner face of the wall section. The brim element is formed in such a way that a projected area of the brim element onto the inner face along a direction of a normal of the inner face at least partially covers the inlet channel.

Abstract: A gas turbine engine combustor includes an annulus with one or more circular rows of burners and a means for providing a number of equiangular spaced apart flame temperature nonuniformities around the annulus during engine operation. The number of the flame temperature nonuniformities being equal to a circumferential acoustic mode to be attenuated in the combustor (i.e three, five, or seven). Fuel lines and/or water lines in supply communication with the burners and metering orifices in a portion of the fuel lines and/or the water lines may be used to produce the flame temperature nonuniformities. The annulus of the burners may have an equal number of equiangular spaced apart first and second arcuate segments of the burners and a means for operating the burners in the first segments and operating the burners in the second segments at different first and second flame temperatures respectively.

Abstract: A fuel nozzle arrangement includes an annular combustor having four quadrants. Fuel injectors are located in the four quadrants and each fuel injector has either a first or second type nozzle attached to it. The first and second type fuel nozzles are located in an array so that at least two of one nozzle type are located together to allow for non-uniform injector spacing with uniform downstream temperatures.

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 combustor cap assembly for a turbine engine includes a combustor cap and a plurality of fuel nozzles mounted on the combustor cap. One or more of the fuel nozzles would include two separate fuel circuits which are individually controllable. The combustor cap assembly would be controlled so that individual fuel circuits of the fuel nozzles are operated or deliberately shut off to provide for physical separation between the flow of fuel delivered by adjacent fuel nozzles and/or so that adjacent fuel nozzles operate at different pressure differentials. Operating a combustor cap assembly in this fashion helps to reduce or eliminate the generation of undesirable and potentially harmful noise.

Abstract: A combustion system includes a first combustion chamber and a second combustion chamber. The second combustion chamber is positioned downstream of the first combustion chamber. The combustion system also includes a pre-mixed, direct-injection secondary fuel nozzle. The pre-mixed, direct-injection secondary fuel nozzle extends through the first combustion chamber into the second combustion chamber.

Abstract: An object of the present invention is to provide a gas turbine combustor that supports hydrogen-containing gas having a high burning velocity and is capable of performing low NOx combustion without reducing reliability of a burner. A first fuel nozzle is provided upstream of a combustion chamber and supplies fuel for activation and hydrogen-containing gas. The combustor has a primary combustion zone, a reduction zone and a secondary combustion zone. In the primary combustion zone, the fuel supplied from the first fuel nozzle is combusted under a fuel rich condition to form a burned gas containing a low concentration of oxygen. In the reduction zone, a hydrogen-containing gas is injected into the combustion chamber through a second fuel injection hole from a second fuel nozzle so that NOx generated in the primary combustion zone is reduced by an oxygen reaction of the hydrogen.

Abstract: A combustor includes a chamber mixing and burning fuel and air and an air hole plate disposed on a wall surface of the chamber. The air hole plate includes a plurality of rows disposed concentrically of a plurality of air holes jetting coaxial jets of fuel and air into the chamber. A first fuel nozzle and a second fuel nozzle are disposed near a fuel hole for jetting fuel into an air hole row on an inner peripheral side. The first fuel nozzle is structured to suppress turbulence of surrounding air flow and the second fuel nozzle is structured to promote turbulence of a surrounding air flow. In accordance with the aspect of the present invention, good flame stability can be maintained while further reducing NOx in a combustor using coaxial jets.

Abstract: Electric fields control an electric combustion. Plasmas are created by a dielectrically impeded discharge. At least one part of the fuel is mixed with water vapor in the supply line and is subjected to a plasma discharge in a compact reactor that is in close contact with the burner in terms of reaction kinetics. In the associated device, a plasma reactor for creating a reactive plasma gas with at least one supply line for the fuel gas-water vapor mixture and an electric energy supply is integrated into the burner. The device provides suitable flow guidance of the fuel gas-air mixture, of the fuel gas-water vapor mixture, and of the reactive plasma gas.

Abstract: A combustion system including a combustor; a combustor liner disposed within the combustor is provided. At least one primary fuel nozzle is provided to provide fuel to a primary combustion zone disposed proximate to the upstream end of the combustor liner. A transition duct is coupled to the downstream end of the combustor liner. A secondary nozzle assembly is disposed proximate to the downstream end of the combustor to provide fuel to a secondary combustion zone at locations predetermined to reduce peak thermal loads on the surface area of the transition duct.

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: A combustor and a combustion method for the combustor, which can suppress backfire and ensure stable combustion. The combustor comprises a mixing-chamber forming member for forming therein a mixing chamber in which air for combustion and fuel are mixed with each other, and a combustion chamber for burning a gas mixture mixed in the mixing chamber and producing combustion gases. A channel for supplying the air for combustion to the mixing chamber from the outer peripheral side of the mixing-chamber forming member is provided inside the mixing-chamber forming member. The fuel and the air are premixed in the channel, and a resulting premixed gas mixture is supplied to the mixing chamber.