Abstract: A combustor for a gas turbine engine is provided, and includes an inner liner; an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner; a fuel igniter comprising a tip portion configured to ignite an air and fuel mixture in the combustion chamber; and an igniter tube positioning the fuel igniter relative to the combustion chamber, the igniter tube having a plurality of channels configured to direct cooling air toward the tip portion of the fuel igniter.

Abstract: A gas turbine engine combustor includes a plurality of main fuel injector assemblies, and a plurality of pilot fuel injector assemblies, that are arranged and configured to reduce exhaust gas emissions during engine operation. The plurality of main fuel injector assemblies are arranged in a substantially circular pattern of a first radius, and each includes an outlet port having a first divergence angle. The plurality of pilot fuel injector assemblies are arranged in a substantially circular pattern of a second radius. Each pilot fuel injector assembly is disposed between at least two main fuel injector assemblies, and each includes an outlet port having a second divergence angle.

Abstract: A combustor liner assembly includes a liner and a first group of cooling holes formed in the liner and having an increasing density in a downstream direction. The first group of cooling holes include a generally circumferential first row of cooling holes, a generally circumferential second row of cooling holes immediately downstream from, consecutive to, and separated from the first row at a first distance, a generally circumferential third row of cooling holes immediately downstream from, consecutive to, and separated from the second row at a second distance, and a generally circumferential fourth row of cooling holes immediately downstream from, consecutive to, and separated from the third row at a third distance. The first distance is greater than the second distance and the third distance, and the second distance is greater than the third distance.

Abstract: A gas turbine engine combustor includes a plurality of main fuel injector assemblies, and a plurality of pilot fuel injector assemblies, that are arranged and configured to reduce exhaust gas emissions during engine operation. The plurality of main fuel injector assemblies are arranged in a substantially circular pattern of a first radius, and each includes an outlet port having a first divergence angle. The plurality of pilot fuel injector assemblies are arranged in a substantially circular pattern of a second radius. Each pilot fuel injector assembly is disposed between at least two main fuel injector assemblies, and each includes an outlet port having a second divergence angle.

Abstract: A combustion system includes a fuel manifold adapted to receive a flow of fuel from a fuel source, the fuel manifold including at least one orifice from which the received fuel is discharged. The combustion system further includes a rotary fuel slinger disposed adjacent to the fuel manifold. The rotary fuel slinger includes a coupler shaft and a slinger disc coupled to the coupler shaft. The slinger disc includes a first portion disposed generally perpendicular to the coupler shaft and a second portion disposed at an angle to the first portion. The fuel manifold is positioned relative to the rotary fuel slinger such that the discharged fuel impinges on the second portion. The rotary fuel slinger is configured to rotate such that the flow of fuel is centrifuged radially outwardly from the second portion onto the first portion and subsequently off the rotary fuel slinger to atomize the received fuel.

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 combustor liner assembly includes a liner and a first group of cooling holes formed in the liner and having an increasing density in a downstream direction.

Abstract: A gas turbine engine combustor includes a plurality of main fuel injector assemblies, and a plurality of pilot fuel injector assemblies, that are arranged and configured to reduce exhaust gas emissions during engine operation. The plurality of main fuel injector assemblies are arranged in a substantially circular pattern of a first radius, and each includes an outlet port having a first divergence angle. The plurality of pilot fuel injector assemblies are arranged in a substantially circular pattern of a second radius. Each pilot fuel injector assembly is disposed between at least two main fuel injector assemblies, and each includes an outlet port having a second divergence angle.

Abstract: A combustor for a gas turbine engine is provided, and includes an inner liner; an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner; a fuel igniter comprising a tip portion configured to ignite an air and fuel mixture in the combustion chamber; and an igniter tube positioning the fuel igniter relative to the combustion chamber, the igniter tube having a plurality of channels configured to direct cooling air toward the tip portion of the fuel igniter.

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 combustion system includes a fuel manifold adapted to receive a flow of fuel from a fuel source, the fuel manifold including at least one orifice from which the received fuel is discharged. The combustion system further includes a rotary fuel slinger disposed adjacent to the fuel manifold. The rotary fuel slinger includes a coupler shaft and a slinger disc coupled to the coupler shaft. The slinger disc includes a first portion disposed generally perpendicular to the coupler shaft and a second portion disposed at an angle to the first portion. The fuel manifold is positioned relative to the rotary fuel slinger such that the discharged fuel impinges on the second portion. The rotary fuel slinger is configured to rotate such that the flow of fuel is centrifuged radially outwardly from the second portion onto the first portion and subsequently off the rotary fuel slinger to atomize the received fuel.

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 premix chamber for a combustor of a gas turbine engine comprises a cylindrical chamber having a premix chamber wall, the cylindrical chamber having a chamber inlet end longitudinally separated from a chamber outlet end along a central axis, a chamber inlet plate in communication with the premix chamber wall at the chamber inlet end, the chamber inlet plate having a fuel nozzle inlet hole disposed through the chamber inlet plate, the chamber inlet plate further comprising a plurality of swirler passages disposed through the chamber inlet plate, and the chamber outlet end being open. A method of producing turbine gas is also disclosed.

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.