Abstract: A secondary fuel stage (14) of a combustor of a gas turbine engine. The combustor has a main combustion zone (43) upstream of the secondary fuel stage to ignite working gas The secondary fuel stage includes a nozzle (18) with dual outlets (20, 22) oriented with a circumferential component to inject an air-fuel mixture (24) into the combustor. The nozzle is effective to entrain the air-fuel mixture with the working gas (44) such that a peak temperature (46) of the working gas at a location downstream of the secondary fuel stage is less than a peak temperature (50) of working gas if the air-fuel mixture were injected into the combustor with a single outlet nozzle (118).

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 flow conditioning device (30, 50, 70, 100, 150) for a can annular gas turbine engine, including a plurality of flow elements (32, 34, 52, 54, 72, 74, 102) disposed in a compressed air flow path (42, 60, 80, 114, 122) leading to a combustor (12), configured such that relative adjustment of at least one flow directing element (32, 52, 72, 110) with respect to an adjacent flow directing element (34, 54, 74, 112, 120) during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path (42, 60, 80, 114, 122).

Abstract: A flow conditioning device (30, 50, 70, 100, 150) for a can annular gas turbine engine, including a plurality of flow elements (32, 34, 52, 54, 72, 74, 102) disposed in a compressed air flow path (42, 60, 80, 114, 122) leading to a combustor (12), configured such that relative adjustment of at least one flow directing element (32, 52, 72, 110) with respect to an adjacent flow directing element (34, 54, 74, 112, 120) during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path (42, 60, 80, 114, 122).

Abstract: A secondary fuel stage (14) of a combustor of a gas turbine engine. The combustor has a main combustion zone (43) upstream of the secondary fuel stage to ignite working gas The secondary fuel stage includes a nozzle (18) with dual outlets (20, 22) oriented with a circumferential component to inject an air-fuel mixture (24) into the combustor. The nozzle is effective to entrain the air-fuel mixture with the working gas (44) such that a peak temperature (46) of the working gas at a location downstream of the secondary fuel stage is less than a peak temperature (50) of working gas if the air-fuel mixture were injected into the combustor with a single outlet nozzle (118).

Abstract: A gas turbine engine (202) including a secondary fuel stage (218) which also functions as a dual frequency resonator. The engine includes a combustor (210) and a casing (205) enclosing the combustor to define a volume (214). The secondary fuel stage includes a nozzle (217) sized to be effective as a transverse resonator at a high frequency. The nozzle and the volume (214) of the casing are sized to be effective as a longitudinal resonator at an intermediate frequency.

Abstract: A fuel gas cooling system for a combustion basket spring clip seal support is disclosed. The fuel gas cooling system may be formed from one or more fuel gas supply channels terminating proximate to a spring clip at the intersection between a combustor basket and a transition section such that fuel gas may be supplied to the hot gas path proximate to the intersection between the combustor basket and the transition section. The fuel gas supply channel may create an intermediate fuel gas burn at this intersection, which may reduce the firing temperature at the fuel nozzles and reduce NOx emissions.

Abstract: An emissions control system for a gas turbine engine including a flow-directing structure (24) that delivers combustion gases (22) from a burner (32) to a turbine. The emissions control system includes: a conduit (48) configured to establish fluid communication between compressed air (22) and the combustion gases within the flow-directing structure (24). The compressed air (22) is disposed at a location upstream of a combustor head-end and exhibits an intermediate static pressure less than a static pressure of the combustion gases within the combustor (14). During operation of the gas turbine engine a pressure difference between the intermediate static pressure and a static pressure of the combustion gases within the flow-directing structure (24) is effective to generate a fluid flow through the conduit (48).

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 (36) for alternate fuels (26A, 26B) with different energy densities. Vanes (47B) extend radially from a fuel delivery tube structure (20B) with first and second fuel supply channels (19A, 19B). Each vane has first and second radial passages (21A, 21B) communicating with the respective fuel supply channels, and first and second sets of apertures (23A, 23B). The first fuel supply channel, first radial passage, and first apertures form a first fuel delivery pathway providing a first fuel flow rate at a given fuel delivery pathway backpressure that is essentially common to both sets of fuel delivery pathway apertures. The second fuel supply channel, second radial passage, and second apertures form a second fuel delivery pathway providing a second fuel flow rate that may be at least 1 about twice the first fuel flow rate at the given fuel delivery pathway backpressure.

Abstract: A fuel shroud assembly (100) into which fuel (118) is injected for mixing with an air stream (120) in a fuel manifold. The shroud assembly (100) comprises a plurality of parallel fuel scoops (102) each receiving the injected fuel (118). The fuel stream (118) flows through each scoop (102), exiting at an open scoop end (114A). The air stream (120) flows between scoops, creating a shear region proximate each scoop end (114A) where the fuel exits. The shear causes mixing of the air (120) and the fuel (118) streams, wherein the degree of mixing is not dependent on the momentum ratio of the air (120) or fuel (118) streams.

Abstract: A fuel shroud assembly (100) into which fuel (118) is injected for mixing with an air stream (120) in a fuel manifold. The shroud assembly (100) comprises a plurality of parallel fuel scoops (102) each receiving the injected fuel (118). The fuel stream (118) flows through each scoop (102), exiting at an open scoop end (114A). The air stream (120) flows between scoops, creating a shear region proximate each scoop end (114A) where the fuel exits. The shear causes mixing of the air (120) and the fuel (118) streams, wherein the degree of mixing is not dependent on the momentum ratio of the air (120) or fuel (118) streams.

Abstract: An emissions control system for a gas turbine engine including a flow-directing structure (24) that delivers combustion gases (22) from a burner (32) to a turbine. The emissions control system includes: a conduit (48) configured to establish fluid communication between compressed air (22) and the combustion gases within the flow-directing structure (24). The compressed air (22) is disposed at a location upstream of a combustor head-end and exhibits an intermediate static pressure less than a static pressure of the combustion gases within the combustor (14). During operation of the gas turbine engine a pressure difference between the intermediate static pressure and a static pressure of the combustion gases within the flow-directing structure (24) is effective to generate a fluid flow through the conduit (48).

Abstract: A catalytic oxidation module for a catalytic combustor of a gas turbine engine is provided. The catalytic oxidation module comprises a plurality of spaced apart catalytic elements for receiving a fuel-air mixture over a surface of the catalytic elements. The plurality of catalytic elements includes at least one primary catalytic element comprising a monometallic catalyst and secondary catalytic elements adjacent the primary catalytic element comprising a multi-component catalyst. Ignition of the monometallic catalyst of the primary catalytic element is effective to rapidly increase a temperature within the catalytic oxidation module to a degree sufficient to ignite the multi-component catalyst.

Abstract: A fuel gas cooling system for a combustion basket spring clip seal support is disclosed. The fuel gas cooling system may be formed from one or more fuel gas supply channels terminating proximate to a spring clip at the intersection between a combustor basket and a transition section such that fuel gas may be supplied to the hot gas path proximate to the intersection between the combustor basket and the transition section. The fuel gas supply channel may create an intermediate fuel gas burn at this intersection, which may reduce the firing temperature at the fuel nozzles and reduce NOx emissions.

Abstract: A pre-mixer assembly associated with a fuel supply system for mixing of air and fuel upstream from a main combustion zone in a gas turbine engine. The pre-mixer assembly includes a swirler assembly disposed about a fuel injector of the fuel supply system and a pre-mixer transition member. The swirler assembly includes a forward end defining an air inlet and an opposed aft end. The pre-mixer transition member has a forward end affixed to the aft end of the swirler assembly and an opposed aft end defining an outlet of the pre-mixer assembly. The aft end of the pre-mixer transition member is spaced from a base plate such that a gap is formed between the aft end of the pre-mixer transition member and the base plate for permitting a flow of purge air therethrough to increase a velocity of the air/fuel mixture exiting the pre-mixer assembly.

Abstract: A catalytic oxidation module for a catalytic combustor of a gas turbine engine is provided. The catalytic oxidation module comprises a plurality of spaced apart catalytic elements for receiving a fuel-air mixture over a surface of the catalytic elements. The plurality of catalytic elements includes at least one primary catalytic element comprising a monometallic catalyst and secondary catalytic elements adjacent the primary catalytic element comprising a multi-component catalyst. Ignition of the monometallic catalyst of the primary catalytic element is effective to rapidly increase a temperature within the catalytic oxidation module to a degree sufficient to ignite the multi-component catalyst.

Abstract: A combustor (22) for a gas turbine (10) includes a main burner oxidizer flow path (34) delivering a first portion (32) of an oxidizer flow (e.g., 16) to a main burner (28) of the combustor and a pilot oxidizer flow path (38) delivering a second portion (36) of the oxidizer flow to a pilot (30) of the combustor. The combustor also includes a flow controller (42) disposed in the pilot oxidizer flow path for controlling an amount of the second portion delivered to the pilot.

Abstract: A fuel injector (36) for alternate fuels (26A, 26B) with energy densities that differ by at least about a factor of two. Vanes (47B) extend radially from a fuel delivery tube structure (20B) with first and second fuel supply channels (19A, 19B). Each vane has first and second radial passages (21A, 21B) communicating with the respective fuel supply channels, and first and second sets of apertures (23A, 23B) between the respective radial passages and the surface (49) of the vane. The first fuel supply channel, first radial passage, and first apertures form a first fuel delivery pathway providing a first fuel flow rate at a given backpressure. The second fuel supply channel, second radial passage, and second apertures form a second fuel delivery pathway providing a second fuel flow rate that may be at least about twice first fuel flow rate at the given backpressure.

Abstract: A pre-mixer assembly associated with a fuel supply system for mixing of air and fuel upstream from a main combustion zone in a gas turbine engine. The pre-mixer assembly includes a swirler assembly disposed about a fuel injector of the fuel supply system and a pre-mixer transition member. The swirler assembly includes a forward end defining an air inlet and an opposed aft end. The pre-mixer transition member has a forward end affixed to the aft end of the swirler assembly and an opposed aft end defining an outlet of the pre-mixer assembly. The aft end of the pre-mixer transition member is spaced from a base plate such that a gap is formed between the aft end of the pre-mixer transition member and the base plate for permitting a flow of purge air therethrough to increase a velocity of the air/fuel mixture exiting the pre-mixer assembly.