Abstract: A combustor for a gas turbine engine comprising a combustion liner defining a combustion chamber, wherein the combustion liner has at least one opening into which a combustor liner mount is received. The combustor liner mount can have at least one cooling passage defined therein having an inlet defining an inlet axis and an outlet defining an outlet axis, wherein the inlet axis is not collinear with the outlet axis for providing enhanced cooling to at least one of the combustor liner mount and an adjacent combustion component, such as an igniter or a borescope plug.

Abstract: Shroud assemblies and methods for cooling gas turbine engine shrouds are provided. For example, a shroud assembly comprises a shroud segment having hot and cold side portions. The hot side portion faces a gas flow path, and the cold side portion faces a cooling fluid flow path. The shroud assembly further comprises a shroud hanger for mounting the shroud segment in the gas turbine engine. The shroud hanger includes a conduit for receipt of a flow of cooling fluid that defines an outlet configured for the flow of cooling fluid to exit the conduit substantially tangential to the cold side portion of the shroud segment. The shroud and shroud hanger together define an annular cavity that forms the cooling fluid flow path, and fluid exits a plurality of conduits defined in the shroud hanger into the cavity substantially tangential to the shroud cold side to cool the shroud cold side.

Abstract: An apparatus and method relating to a cooling hole of a component of a turbine engine. The component can include a wall separating the hot gas fluid flow from the cooling fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow and at least one cooling hole comprising at least one inlet at the cooled surface, at least one outlet at the heated surface, with the outlet having a modified outlet shape.

Abstract: A gas turbine engine includes a combustion section space between a compressor section in a turbine section. The combustion section defines a combustion chamber and includes a combustor member defining an opening to the combustion chamber. A mounting assembly extends around or is positioned adjacent to the opening defined by the combustor member. In igniter extends through a ferrule of the mounting assembly and includes a distal end positioned proximate the opening in the combustor member. The igniter defines a plurality of channels each channel extending between a first end and a second end. The first end is positioned away from the distal end of the igniter relative to the second end, and the second end is a terminal end spaced from the distal end of the igniter.

Abstract: An apparatus and method relating to a cooling hole of a component of a turbine engine. The component can include a wall separating the hot gas fluid flow from the cooling fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow and at least one cooling hole comprising at least one inlet at the cooled surface, at least one outlet at the heated surface, with the outlet having a modified outlet shape.

Abstract: An aspect of the present disclosure is directed to a rotor assembly for a turbine engine. The rotor assembly includes an airfoil assembly and a hub to which the airfoil assembly is attached. A wall assembly defines a first cavity and a second cavity between the airfoil assembly and the hub. The first cavity and the second cavity are at least partially fluidly separated by the wall assembly. The first cavity is in fluid communication with a flow of first cooling fluid and the second cavity is in fluid communication with a flow of second cooling fluid different from the first cooling fluid.

Abstract: The present disclosure is directed to a fuel nozzle for a gas turbine engine, the fuel nozzle defining a radial direction, a longitudinal direction, a circumferential direction, an upstream end, and a downstream end. The fuel nozzle includes an aft body coupled to at least one fuel injector. The aft body defines a forward wall and an aft wall each extended in the radial direction, and a plurality of sidewalls extended in the longitudinal direction. The plurality of sidewalls couples the forward wall and the aft wall. The forward wall defines at least one channel inlet orifice. At least one sidewall defines at least one channel outlet orifice. At least one micro channel cooling circuit is defined between the one or more channel inlet orifices and the one or more channel outlet orifices.

Abstract: Shroud assemblies and methods for cooling gas turbine engine shrouds are provided. For example, a shroud assembly comprises a shroud segment having hot and cold side portions. The hot side portion faces a gas flow path, and the cold side portion faces a cooling fluid flow path. The shroud assembly further comprises a shroud hanger for mounting the shroud segment in the gas turbine engine. The shroud hanger includes a conduit for receipt of a flow of cooling fluid that defines an outlet configured for the flow of cooling fluid to exit the conduit substantially tangential to the cold side portion of the shroud segment. The shroud and shroud hanger together define an annular cavity that forms the cooling fluid flow path, and fluid exits a plurality of conduits defined in the shroud hanger into the cavity substantially tangential to the shroud cold side to cool the shroud cold side.

Abstract: A heat engine including a wall assembly is generally provided. The wall assembly includes a plurality of radial walls coupled together via a connecting member. The radial wall defines a flow opening therethrough. A flow cavity is defined between the plurality of radial walls and the connecting member.

Abstract: A combustor assembly for a heat engine is generally provided. The combustor assembly includes a liner wall defining a combustion chamber, and a deflector assembly. The deflector assembly includes a radially extended first wall disposed adjacent to the combustion chamber, and further an axially extended second wall disposed forward of the first wall and adjacent thereto. The second wall is coupled to the liner wall.

Abstract: A combustor assembly for a turbine engine and method for cooling a portion of the combustor assembly, the combustor assembly comprising a combustor liner defining a combustion chamber with an inlet and an outlet, a fuel dome located in the inlet and having a wall adjacent at least a portion of the combustor liner, and a deflector overlying and spaced from at least a portion of the wall.

Abstract: A combustor for a gas turbine engine comprising a combustion liner defining a combustion chamber, wherein the combustion liner has at least one opening into which a combustor liner mount is received. The combustor liner mount can have at least one cooling passage defined therein having an inlet defining an inlet axis and an outlet defining an outlet axis, wherein the inlet axis is not collinear with the outlet axis for providing enhanced cooling to at least one of the combustor liner mount and an adjacent combustion component, such as an igniter or a borescope plug.

Abstract: A gas turbine engine having an improved air delivery system that includes features for pressurizing and/or cooling various components of the engine while minimizing the impact to the cycle efficiency of the engine, reducing the weight of the engine, and reducing the specific fuel consumption of the engine is provided.

Abstract: Axially adjacent annular booster bleed aft and forward plenums with annular common wall therebetween extend radially outwardly from transition duct. Variable bleed valve includes bleed valve door in bleed inlet in transition duct, attached to rotatable valve body rotatable about axis of rotation, operable to open and close bleed inlet to aft plenum. Rotatable plenum door clocked or circumferentially spaced apart from variable bleed valve door and attached to rotatable valve body, operable to close and open up and control flow through an inter plenum aperture in common wall. Aft and forward bleed exhaust ducts extend from booster bleed aft and forward plenums to bypass flow path. One or more heat exchanger, such as from thermal management system, may be disposed in the bleed exhaust ducts. Heat exchangers may be used for cooling oil for power gear box and/or engine bearings, air conditioning, or variable frequency generator.

Abstract: A leakage management assembly for an orifice configured to limit a flow of at least one of a leakage fluid and a cooling fluid past an instrument positioned within the orifice are provided. The leakage management assembly includes a labyrinth ferrule including an annular ferrule body having a central bore extending therethrough. The labyrinth ferrule further includes an annular assembly cone formed at a first end of the labyrinth ferrule. The annular assembly cone includes an annular convergent lip and is configured to facilitate directing the instrument into the central bore. The labyrinth ferrule also includes a plurality of labyrinthine annular restrictions extending from a radially inner surface of the annular body toward the central bore. A combustor and a gas turbine engine including such a seal assembly are also provided.

Abstract: A combustor for a gas turbine engine comprising a combustion liner defining a combustion chamber, wherein the combustion liner has at least one opening into which a combustor liner mount is received. The combustor liner mount can have at least one cooling passage defined therein having an inlet defining an inlet axis and an outlet defining an outlet axis, wherein the inlet axis is not collinear with the outlet axis for providing enhanced cooling to at least one of the combustor liner mount and an adjacent combustion component, such as an igniter or a borescope plug.

Abstract: A combustor assembly for a gas turbine engine defining an axial direction is provided. The combustor assembly generally includes a liner, an annular dome, and a deflector plate. The liner at least partially defines a combustion chamber. The annular dome defines a first cavity and has a plurality of impingement holes, and the deflector plate defines a conical surface and a flat surface, wherein a plurality of cooling holes are defined through the deflector plate. The annular dome and the deflector plate are positioned together to define a second cavity that is in fluid communication with the first cavity through the plurality of impingement holes. In addition, the second cavity is in fluid communication with the combustion chamber through the plurality of cooling holes.

Abstract: The present disclosure is directed to a fuel nozzle for a gas turbine engine, the fuel nozzle defining a radial direction, a longitudinal direction, a circumferential direction, an upstream end, and a downstream end. The fuel nozzle includes an aft body coupled to at least one fuel injector. The aft body defines a forward wall and an aft wall each extended in the radial direction, and a plurality of sidewalls extended in the longitudinal direction. The plurality of sidewalls couples the forward wall and the aft wall. The forward wall defines at least one channel inlet orifice. At least one sidewall defines at least one channel outlet orifice. At least one micro channel cooling circuit is defined between the one or more channel inlet orifices and the one or more channel outlet orifices.

Abstract: Axially adjacent annular booster bleed aft and forward plenums with annular common wall therebetween extend radially outwardly from transition duct. Variable bleed valve includes bleed valve door in bleed inlet in transition duct, attached to rotatable valve body rotatable about axis of rotation, operable to open and close bleed inlet to aft plenum. Rotatable plenum door clocked or circumferentially spaced apart from variable bleed valve door and attached to rotatable valve body, operable to close and open up and control flow through an inter plenum aperture in common wall. Aft and forward bleed exhaust ducts extend from booster bleed aft and forward plenums to bypass flow path. One or more heat exchanger, such as from thermal management system, may be disposed in the bleed exhaust ducts. Heat exchangers may be used for cooling oil for power gear box and/or engine bearings, air conditioning, or variable frequency generator.

Abstract: A combustor for a gas turbine engine comprising a combustion liner defining a combustion chamber, wherein the combustion liner has at least one opening into which a combustor liner mount is received. The combustor liner mount can have at least one cooling passage defined therein having an inlet defining an inlet axis and an outlet defining an outlet axis, wherein the inlet axis is not collinear with the outlet axis for providing enhanced cooling to at least one of the combustor liner mount and an adjacent combustion component, such as an igniter or a borescope plug.