Patent ID: 9366438
Filing Date: 2016-06-14
CPC Classification: F23R

Claim Text:
1. A combustor assembly in a gas turbine engine comprising: a liner defining a combustion zone where fuel and air are mixed and burned to create a hot working gas that flows through the combustion zone generally in a first direction toward a turbine section of the engine; at least one fuel injector for providing the fuel to be burned in the combustion zone; a flow sleeve located radially outwardly from the liner, wherein an inner surface of the flow sleeve defines an outer boundary for an air flow passageway where the air to be burned in the combustion zone flows generally in a second direction opposite to the first direction, wherein upon the air reaching a head end of the combustor assembly at an end of the air flow passageway the air turns 180 degrees to flow generally in the first direction into the combustion zone where the air is burned with the fuel; an inlet assembly positioned radially between the liner and the flow sleeve, the inlet assembly defining an inlet to the air flow passageway and comprising a plurality of overlapping conduits that are arranged such that the air entering the air flow passageway passes through radial spaces between adjacent conduits; and wherein the number of conduits, radial heights between adjacent conduits, and lengths of conduits overlap are each selected to fine tune acoustic losses provided by the inlet assembly.