Abstract: An airfoil with dual-wall cooling for a gas turbine engine comprises a spar having a pressure side wall and a suction side wall meeting at a leading edge and a trailing edge of the airfoil. An interior of the spar comprises a coolant cavity. The suction side wall includes an arrangement of rails on an outer surface thereof, where each rail extends along a non-chordal direction and terminates at or near the trailing edge. The airfoil also comprises a suction side coversheet overlying the suction side wall, where an inner surface of the suction side coversheet is in contact with the arrangement of rails so as to define a plurality of angled channels between the suction side wall and the suction side coversheet. The suction side wall also comprises inlet holes in fluid communication with the coolant cavity for feeding coolant to the angled channels, and the arrangement of rails is configured to direct the coolant through the angled channels and toward the trailing edge of the airfoil.

Abstract: An example system includes a casting mold and a casting core. The casting core includes a substrate. A plurality of support structures integral with and extending from the substrate define a plurality of channels. Respective support structures of the plurality of support structures define respective contact surfaces distal from the substrate. A sacrificial composition substantially fully fills the plurality of cooling channels and leaves the respective contact surfaces substantially uncovered. An example technique includes filling the sacrificial composition in the plurality of cooling channels, and casting a cover layer onto the respective contact surfaces of the plurality of support structures.

Abstract: An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel.

Abstract: An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel.

Abstract: Techniques for forming a dual-walled component for a gas turbine engine that include chemically etching at least one of a hot section part or a cold section part to form an etched part having plurality of support structures and bonding the etched part to a corresponding cold section part or a corresponding hot section part to form a dual-walled component, with the plurality of support structures defining at least one cooling channel between the at least one of the hot section part or the cold section part and the corresponding cold section part or the corresponding hot section part.

Abstract: An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel.

Abstract: An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel.

Abstract: An example gas turbine engine component includes a component configured to separate a cooling air plenum from a heated gas environment. The component includes a substrate defining a surface, and a unitary structure. The unitary structure includes a cooling region and a cover layer. The cover layer defines a hot wall surface configured to face the heated gas environment. The cooling region is disposed between the cover surface and the substrate and includes a plurality of support structures extending between the cover layer and the surface of the substrate. At least some of the support structures define a respective bond surface bonded to the substrate at the surface of the substrate. An example technique for fabricating the gas turbine engine component includes additively depositing the unitary structure on the surface of the substrate.

Abstract: Techniques for forming a dual-walled component for a gas turbine engine that include chemically etching at least one of a hot section part or a cold section part to form an etched part having plurality of support structures and bonding the etched part to a corresponding cold section part or a corresponding hot section part to form a dual-walled component, with the plurality of support structures defining at least one cooling channel between the at least one of the hot section part or the cold section part and the corresponding cold section part or the corresponding hot section part.

Abstract: An example system includes a casting mold and a casting core. The casting core includes a substrate. A plurality of support structures integral with and extending from the substrate define a plurality of channels. Respective support structures of the plurality of support structures define respective contact surfaces distal from the substrate. A sacrificial composition substantially fully fills the plurality of cooling channels and leaves the respective contact surfaces substantially uncovered. An example technique includes filling the sacrificial composition in the plurality of cooling channels, and casting a cover layer onto the respective contact surfaces of the plurality of support structures.

Abstract: A fuel nozzle for a gas turbine engine is provided having swirlers and a mounting device configured to mount the fuel nozzle to the inside of a combustor. In one embodiment the combustor has an interior surface formed to receive the mounting device of the fuel nozzle from within the combustor. The swirlers can be integrally formed in the fuel nozzle or can be separately attached, thus allowing the fuel nozzle to be mounted within the combustor having the relative placement of the swirlers already determined prior to installation. The fuel nozzle tip can be positioned a predetermined distance from the outlet of a swirler to reduce NOx emissions.