Abstract: An injector module includes an injector stem that extends along an injector longitudinal axis between an inlet end and an outlet end of the injector module. The injector module also includes a first fuel line of a first fuel circuit at least partly extending through the injector stem. The first fuel line has a first outlet disposed at the outlet end of the injector stem. The injector module further includes a second fuel line of a second fuel circuit at least partly extending through the injector stem. The second fuel line has a second outlet disposed at the outlet end of the injector stem. The first outlet and the second outlet are spaced apart and have different orientations relative to the injector longitudinal axis. The first fuel line is thermally coupled to the second fuel line.

Abstract: A gas turbine engine component includes a hot side surface that is configured for exposure to a hot gas flow path, a second surface that is opposite the hot side surface and not exposed to the hot gas flow path, and an effusion cooling aperture positioned along the hot side surface. The effusion cooling aperture includes a recessed portion including a void area beginning at the hot side surface and extending inwardly therefrom, a forward surface, an entirety of which is angled at 90 degrees or greater than 90 degrees with respect to the hot side surface, defining a forward end of the recessed portion, an inward surface, angled with respect to the hot side surface, and connecting the forward surface to the hot side surface, and an overhang portion connected with the hot side surface and extending aftward from the forward surface and over the void area.

Abstract: A gas turbine engine component includes a hot side surface that is configured for exposure to a hot gas flow path, a second surface that is opposite the hot side surface and not exposed to the hot gas flow path, and an effusion cooling aperture positioned along the hot side surface. The effusion cooling aperture includes a recessed portion including a void area beginning at the hot side surface and extending inwardly therefrom, a forward surface, an entirety of which is angled at 90 degrees or greater than 90 degrees with respect to the hot side surface, defining a forward end of the recessed portion, an inward surface, angled with respect to the hot side surface, and connecting the forward surface to the hot side surface, and an overhang portion connected with the hot side surface and extending aftward from the forward surface and over the void area.

Abstract: A gas turbine engine combustor is provided. An inner liner has an upstream end and a downstream end and extends in an axial direction between the upstream and downstream ends. A dual wall outer liner has a hot wall, a cold wall at least partially surrounding the hot wall, an upstream end, and a downstream end. The outer liner extends in the axial direction between the upstream and downstream ends. A dome assembly is coupled between the upstream ends of the inner and outer liners to define a combustion chamber between the inner liner and the hot wall of the outer liner. Effusion cooling holes are disposed in the hot wall, including a first row disposed at a tangential angle of between about 70° and about 90° and a second row disposed at a tangential angle of between about 0° and about 20°.

Abstract: In accordance with an exemplary embodiment, a fuel injector assembly is provided for a gas turbine engine with a compressor section and a combustion section. The fuel injector assembly includes a stem defining a fuel path for fuel and a flow guide coupled to and extending along the stem. The flow guide receives an air flow. The assembly further includes a swirler apparatus coupled to the stem and configured to receive the fuel. The swirler apparatus is further coupled to the flow guide and configured to receive the air flow. The swirler apparatus is configured to mix the fuel and the air flow and direct the mixture into a combustor of the combustor assembly.

Abstract: A combustor dome and heat-shield assembly comprises a heat-shield having a first opening therethrough. A swirler extends through the first opening and captures the heat-shield. A dome having a second opening therethrough and has an upstream side and a downstream side. The swirler extends through the second opening from the downstream side to the upstream side to capture the heat-shield on the downstream side, and a retaining clip engages the swirler to secure the swirler and the heat-shield on the dome.

Abstract: In accordance with an exemplary embodiment, a fuel injector assembly is provided for a gas turbine engine with a compressor section and a combustion section. The fuel injector assembly includes a stem defining a fuel path for fuel and a flow guide coupled to and extending along the stem. The flow guide receives an air flow. The assembly further includes a swirler apparatus coupled to the stem and configured to receive the fuel. The swirler apparatus is further coupled to the flow guide and configured to receive the air flow. The swirler apparatus is configured to mix the fuel and the air flow and direct the mixture into a combustor of the combustor assembly.

Abstract: A combustor for a gas turbine engine includes an inner liner and an outer liner circumscribing the inner liner and forming a combustion chamber therewith. The outer liner is a dual walled liner with a first wall and a second wall. A fuel igniter includes a tip portion configured to ignite an air and fuel mixture in the combustion chamber. An igniter support assembly positions the fuel igniter relative to the combustion chamber. The igniter support assembly defines a plurality of holes configured to direct cooling air toward the tip portion of the fuel igniter. The igniter support assembly includes first and second floating seals that are configured to accommodate radial and axial relative movements.

Abstract: A combustor dome and heat-shield assembly comprises a heat-shield having a first opening therethrough. A swirler extends through the first opening and captures the heat-shield. A dome having a second opening therethrough and has an upstream side and a downstream side. The swirler extends through the second opening from the downstream side to the upstream side to capture the heat-shield on the downstream side, and a retaining clip engages the swirler to secure the swirler and the heat-shield on the dome.

Abstract: A combustor for a turbine engine is provided. The combustor includes a first liner and a second liner forming a combustion chamber. The combustion chamber is configured to receive an air-fuel mixture for combustion therein and having a longitudinal axis that defines axial and radial directions. The first liner is a first dual walled liner having a first hot wall facing the combustion chamber and a first cold wall that forms a first liner cavity with the first hot wall, the first liner cavity having first and second ends. A first liner seal is configured to seal the second end of the first liner cavity and to accommodate relative movement of the first hot wall and first cold wall generally in the axial and radial directions.

Abstract: A combustor includes a first liner; and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein, and the first liner is a first dual wall liner having a first hot wall facing the combustion chamber and a first cold wall. The first cold wall has a first cold wall orifice and the first hot wall has a first hot wall orifice. A first insert is mounted in the first hot wall orifice and is configured to receive a first air jet of pressured air from the first cold wall orifice, and guide the first jet through the first hot wall orifice and into the combustion chamber.

Abstract: A combustor for a turbine engine is provided. The combustor includes a first liner and a second liner forming a combustion chamber. The combustion chamber is configured to receive an air-fuel mixture for combustion therein and having a longitudinal axis that defines axial and radial directions. The first liner is a first dual walled liner having a first hot wall facing the combustion chamber and a first cold wall that forms a first liner cavity with the first hot wall, the first liner cavity having first and second ends. A first liner seal is configured to seal the second end of the first liner cavity and to accommodate relative movement of the first hot wall and first cold wall generally in the axial and radial directions.

Abstract: A combustor for a gas turbine engine includes an inner liner and an outer liner circumscribing the inner liner and forming a combustion chamber therewith. The outer liner is a dual walled liner with a first wall and a second wall. A fuel igniter includes a tip portion configured to ignite an air and fuel mixture in the combustion chamber. An igniter support assembly positions the fuel igniter relative to the combustion chamber. The igniter support assembly defines a plurality of holes configured to direct cooling air toward the tip portion of the fuel igniter. The igniter support assembly includes first and second floating seals that are configured to accommodate radial and axial relative movements.

Abstract: A gas turbine engine combustor is provided. An inner liner has an upstream end and a downstream end and extends in an axial direction between the upstream and downstream ends. A dual wall outer liner has a hot wall, a cold wall at least partially surrounding the hot wall, an upstream end, and a downstream end. The outer liner extends in the axial direction between the upstream and downstream ends. A dome assembly is coupled between the upstream ends of the inner and outer liners to define a combustion chamber between the inner liner and the hot wall of the outer liner. Effusion cooling holes are disposed in the hot wall, including a first row disposed at a tangential angle of between about 70° and about 90° and a second row disposed at a tangential angle of between about 0° and about 20°.

Abstract: A cooled hybrid structure is provided for deployment within a gas turbine engine. In one embodiment, the cooled hybrid structure includes a woven oxide fiber sheet and an insulative oxide coating. The woven oxide fiber sheet includes an outer cold wall and an inner hot wall, which is integrally woven with the outer cold wall and which cooperates therewith to define a plurality of elongated cooling channels extending within the woven oxide fiber sheet. A plurality of impingement apertures is formed through the outer cold wall and conducts airflow into the plurality of elongated cooling channels and against the inner hot wall to convectively cool the woven oxide fiber sheet. A plurality of effusion channels is formed through the inner hot wall and through the insulative oxide coating and conducts airflow through the insulative oxide coating to provide convective cooling thereof.

Abstract: A combustor includes a first liner; and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein, and the first liner is a first dual wall liner having a first hot wall facing the combustion chamber and a first cold wall. The first cold wall has a first cold wall orifice and the first hot wall has a first hot wall orifice. A first insert is mounted in the first hot wall orifice and is configured to receive a first air jet of pressured air from the first cold wall orifice, and guide the first jet through the first hot wall orifice and into the combustion chamber.

Abstract: A retaining ring is provided for use in conjunction with a carburetor assembly having a flow passage therethrough for conducting a fuel-air mixture into a combustion chamber, the carburetor assembly comprising an air flow modifier and a fuel injector receiving element adjacent thereto. The retaining ring comprises a collar configured to engage the fuel injector receiving element so as to maintain its position adjacent the air flow modifier, and a retaining portion coupled to the collar and configured to engage the air flow modifier to secure the airflow modifier to the combustion chamber. The retaining portion cooperates with the collar to form an aperture through the retaining ring for receiving the carburetor assembly.

Abstract: A system is provided for aerodynamically coupling air flow from a centrifugal compressor to an axial combustor. The system includes a diffuser, a deswirl assembly, combustor inner and outer annular liners, a combustor dome, and a curved annular plate. The diffuser has an inlet that communicates with the centrifugal compressor, an outlet, and a flow path that extends radially outward. The deswirl assembly has an inlet that communicates with the diffuser outlet to receive air flowing in a radially outward direction, an outlet, and a flow path configured to redirect the air in a radially inward and axial direction through the deswirl assembly outlet at an angle toward a longitudinal axis. The curved annular plate is coupled to combustor inner and outer annular liner upstream ends to form a combustor subplenum therebetween and has a first opening and a second opening formed therein, the first opening aligned with the deswirl assembly outlet to receive air discharged therefrom.

Abstract: A layered structure includes a substrate comprising a layer of an oxide/oxide ceramic based composite material, a first oxide layer disposed directly on the substrate and formed from a material that has no greater than about 10% porosity and is substantially impermeable by water vapor, and a second oxide layer disposed directly on the first oxide layer and having a greater porosity than the first oxide layer. Either or both the first and second oxide layers of the coating system may be deposited using a plasma spraying process, a slurry deposition process which is followed by a sintering step, or an EB-PVD process.

Abstract: A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.