A method and apparatus for an airfoil in a gas turbine engine can include an outer surface bounding an interior. At least one flow channel can be defined among one or more full-length and partial-length ribs to further define a cooling circuit within the airfoil. The cooling circuit can have at least one tip turn at the partial-length rib, having at least on fastback turbulator disposed at least partially within the tip turn.

The present disclosure is directed to a rotor blade for a gas turbine engine. The rotor blade includes a platform having a radially inner surface and a radially outer surface. A connection portion extends radially inwardly from the radially inner surface of the platform. An airfoil extends radially outwardly from the radially outer surface of the platform to an airfoil tip. The airfoil includes a leading edge portion and a trailing edge portion. The platform, the airfoil, and the connection portion collectively define a cooling circuit extending from an inlet defined by the connection portion to one or more outlet passages at least partially defined by the trailing edge portion of the airfoil. At least one of the one or more outlet passages include an entrance and an exit radially inwardly offset from the entrance.

A stator vane assembly for a gas turbine engine in which both higher pressure cooling air and lower pressure cooling air are both supplied to the stator vane assembly to cool both an airfoil and inner and outer diameter endwall cavities of the stator vane assembly, in which the spent higher pressure cooling air is then discharged into a combustor of the gas turbine engine. The higher pressure cooling air flows through a closed loop cooling circuit formed within the stator vane assembly while the lower pressure cooling air is discharged through exit holes into the hot gas stream of the turbine.

An abradable sealing element comprises a substrate and a sealing structure. The sealing structure comprises one or more wall structures extending from the substrate and defining at least one open cell which is filled with abradable material. The one or more wall structures are formed by additive-layer, powder-fed, laser-weld deposition onto the substrate. The one or more wall structures are formed from nickel-based superalloy and constitute from about 10% to about 50% of the total volume of the sealing structure.

An airfoil assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil that has an airfoil section extending from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction, and the airfoil section defining a pressure side and a suction side separated in a circumferential direction. A platform includes a first and second platform portions attached to the root section. Each of the first and second platform portions includes a shoe section and a platform section extending in the circumferential direction from the shoe section to establish a gas path surface, and the shoe sections of the first and second platforms are circumferentially arranged on opposed sides of the root section to capture the root section in a root cavity established between the shoe sections. A method of assembly is also disclosed.

An airfoil for a gas turbine engine defining a spanwise direction, a root end, a tip end, a leading edge end, and trailing edge end is provided. The airfoil includes: a body extending along the spanwise direction between the root end and the tip end, the body formed of a composite material; and a sculpted leading edge member attached to the body positioned at the leading edge end of the airfoil, the sculped leading edge member formed at least in part of a metal material and defining a non-linear patterned leading edge of the airfoil.

An airfoil cooling circuit includes an impingement cooling circuit and a serpentine cooling circuit. An airfoil for use in a gas turbine engine having a cooling circuit which includes an impingement cooling circuit and a serpentine cooling circuit.

A fluid cooling arrangement in a gas turbine engine for aerospace propulsion includes an inner structure. Also included is an outer structure disposed radially outwardly of the inner structure, the outer structure and the inner structure defining a bypass flow path. Further included is at least one strut operatively coupling the inner structure to the outer structure. Yet further included is at least one cooling tube formed within the at least one strut, the at least one cooling tube configured to cool a fluid passing through the at least one cooling tube upon convective cooling of the at least one strut as air passes through the bypass flow path and over the at least one strut.

A blade for a turbomachine, extending longitudinally between a base and an apex, comprising a lower surface wall, an upper surface wall, a leading edge and a trailing edge and comprising a plurality of internal ventilation cavities which form a cooling circuit of the blade, wherein at least one ventilation cavity is a ventilation cavity of a first type comprising at least one ascending portion which extends substantially longitudinally between the base and the apex, engaged with the lower surface wall and spaced from the upper surface wall, and at least one discharge portion which extends substantially transversely and which opens at the trailing edge via at least one port of the trailing edge, and wherein at least one ventilation cavity of the first type further comprises at least one descending portion which extends substantially longitudinally from the apex, engaged with the upper surface wall and spaced from the lower

A turbine vane includes a blade body and a shroud. The shroud includes a gas path surface, a front end surface, a front end corner portion which is a corner portion between the gas path surface and the front end surface, a cavity defining surface which defines a cavity allowing cooling air to flow thereinto, a first air passage in which the cooling air flows, and a second air passage in which the cooling air flows. The first air passage includes a first inlet opened at the cavity defining surface and a first outlet opened at the front end corner portion. The second air passage includes a second inlet opened at the cavity defining surface and a second outlet opened at the front end surface.