Abstract: Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.

Abstract: A gas turbine engine having a rotor blade stage with a plurality of circumferentially spaced rotors, a nozzle stage adjacent the rotor blade stage and including an outer nozzle end, at least one stop, and a shroud. The shroud having a forward end positioned radially outward from the circumferentially spaced rotor blades, and an aft end axially aft of the forward end. The at least one stop confronting at least a portion of the outer nozzle end.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Abstract: Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

Abstract: Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity.

Abstract: Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component.

Abstract: Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.

Abstract: Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component.

Abstract: Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.

Abstract: Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Abstract: Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component.

Abstract: Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Abstract: Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

Abstract: Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

Abstract: Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.

Abstract: Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Abstract: Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity.