Abstract: The disclosure describes articles and techniques that include an airfoil having a hybrid coating system to provide improved particle impact resistance and improve CMAS attack resistance on the pressure side of the airfoil and improved thermal load protection on the suction side of the airfoil. An example article for a gas turbine engine may include a substrate, and a hybrid environmental barrier coating (EBC) including a relatively dense EBC layer on a first portion of the substrate and a relatively porous EBC layer on a second portion of the substrate, where the first portion of the substrate is different from the second portion of the substrate, and wherein at least a portion of the relatively porous EBC layer overlaps at least a portion of the relatively dense EBC layer in an overlap region.

Abstract: A gas turbine engine component may include an airfoil extending radially from a base to a blade tip, the airfoil including a pressure sidewall and a suction sidewall each extending between a leading edge and a trailing edge opposite the leading edge, an internal cooling circuit extending from the base to the blade tip; and a squealer tip. The squealer tip may include a first and a second squealer tip rail adjacent to the first squealer tip rail, and a squealer tip cap extending between the first and second squealer tip rails. The blade tip, the first and second squealer tip rails, and the squealer tip cap may define an internal squealer tip cooling channel. The blade tip may define a supply aperture that fluidly connect the internal cooling circuit and the internal squealer tip cooling channel.

Abstract: A gas turbine engine component may include an airfoil extending radially from a base to a blade tip, the airfoil including a pressure sidewall and a suction sidewall each extending between a leading edge and a trailing edge opposite the leading edge, an internal cooling circuit extending from the base to the blade tip; and a squealer tip. The squealer tip may include a first and a second squealer tip rail adjacent to the first squealer tip rail, and a squealer tip cap extending between the first and second squealer tip rails. The blade tip, the first and second squealer tip rails, and the squealer tip cap may define an internal squealer tip cooling channel. The blade tip may define a supply aperture that fluidly connect the internal cooling circuit and the internal squealer tip cooling channel.

Abstract: The disclosure describes articles and techniques that include an airfoil having a hybrid coating system to provide improved particle impact resistance and improve CMAS attack resistance on the pressure side of the airfoil and improved thermal load protection on the suction side of the airfoil. An example article for a gas turbine engine may include a substrate, and a hybrid environmental barrier coating (EBC) including a relatively dense EBC layer on a first portion of the substrate and a relatively porous EBC layer on a second portion of the substrate, where the first portion of the substrate is different from the second portion of the substrate, and wherein at least a portion of the relatively porous EBC layer overlaps at least a portion of the relatively dense EBC layer in an overlap region.

Abstract: A vane adapted for use in a gas turbine engine includes a forward following serpentine cooling path having a plurality of divider walls to direct cooling airflow. The serpentine cooling path optionally includes one or more bifurcating walls in the cooling path to split the airflow. The serpentine cooling path is in communication with an aft inlet in the vane and air passing through the serpentine cooling path exits a forward outlet in the vane into a forward rotor/stator cavity.

Abstract: A gas turbine engine component may include an airfoil extending radially from a base to a blade tip, the airfoil including a pressure sidewall and a suction sidewall each extending between a leading edge and a trailing edge opposite the leading edge, an internal cooling circuit extending from the base to the blade tip; and a squealer tip. The squealer tip may include a first and a second squealer tip rail adjacent to the first squealer tip rail, and a squealer tip cap extending between the first and second squealer tip rails. The blade tip, the first and second squealer tip rails, and the squealer tip cap may define an internal squealer tip cooling channel. The blade tip may define a supply aperture that fluidly connect the internal cooling circuit and the internal squealer tip cooling channel.

Abstract: One embodiment of the present invention is a unique method of manufacturing a component for a turbomachine, such as an airfoil. Another embodiment is a unique airfoil. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooled gas turbine engine components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: An airfoil may be provided that includes a CMC body and a support piece. The CMC body has an inner surface that defines a chamber within the CMC body. The support piece may be positioned within the chamber of the CMC body. The support piece comprises a channel in a surface of the support piece, the surface being in contact with the inner surface of the CMC body. The channel and the inner surface of the CMC body define a passageway for a cooling fluid. The passageway may wind about the circumference of the CMC body and extend along the span of the airfoil. Outlets may be positioned through the CMC body allowing fluid communication between the passageway and the outer surface of the CMC body.

Abstract: A vane adapted for use in a gas turbine engine is disclosed. The vane includes a forward following serpentine cooling path having a plurality of divider walls to direct cooling airflow. The serpentine cooling path optionally includes one or more bifurcating walls in the cooling path to split the airflow. The serpentine cooling path is in communication with an aft inlet in the vane and air passing through the serpentine cooling path exits a forward outlet in the vane into a forward rotor/stator cavity.

Abstract: An airfoil may be provided that includes a CMC body and a support piece. The CMC body has an inner surface that defines a chamber within the CMC body. The support piece may be positioned within the chamber of the CMC body. The support piece comprises a channel in a surface of the support piece, the surface being in contact with the inner surface of the CMC body. The channel and the inner surface of the CMC body define a passageway for a cooling fluid. The passageway may wind about the circumference of the CMC body and extend along the span of the airfoil. Outlets may be positioned through the CMC body allowing fluid communication between the passageway and the outer surface of the CMC body.

Abstract: One embodiment of the present invention is a unique method of manufacturing a component for a turbomachine, such as an airfoil. Another embodiment is a unique airfoil. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooled gas turbine engine components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A turbine airfoil arrangement for a gas turbine engine includes an airfoil having an inlet and an exit, the inlet configured to receive a cooling gas flow operable to cool at least part of an other airfoil; and a passage disposed in the airfoil and fluidly coupled to the inlet and the exit, the exit being configured to pass at least some of the cooling gas flow to the other airfoil.

Abstract: A turbine airfoil arrangement for a gas turbine engine includes an airfoil having an inlet and an exit, the inlet configured to receive a cooling gas flow operable to cool at least part of an other airfoil; and a passage disposed in the airfoil and fluidly coupled to the inlet and the exit, the exit being configured to pass at least some of the cooling gas flow to the other airfoil.

Abstract: A turbine airfoil arrangement for a gas turbine engine includes an airfoil having an inlet and an exit, the inlet configured to receive a cooling gas flow operable to cool at least part of an other airfoil; and a passage disposed in the airfoil and fluidly coupled to the inlet and the exit, the exit being configured to pass at least some of the cooling gas flow to the other airfoil.