Abstract: A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

Abstract: A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

Abstract: In one example embodiment, a blade includes an attachment region, an airfoil extending from the attachment region, and a blade cooling arrangement. The blade cooling arrangement includes at least a first feed passage disposed through the attachment region, which is connected to a first cooling passage disposed in the airfoil. A passively actuated first coolant valve is disposed in or proximate the first feed passage. A plurality of such blades can be disposed in a turbine section of an engine.

Abstract: In a featured embodiment, a lost core assembly includes a ceramic component having a tapered shape in a radial direction. A refractory metal component extends radially from the ceramic core component. A method of molding a gas turbine engine component is also disclosed.

Abstract: A gas turbine engine component comprises an airfoil with a suction side and pressure side extending from a leading edge to a trailing edge. There are a plurality of cooling holes adjacent the leading edge, with the cooling holes having a non-circular shape, with a longer dimension and a smaller dimension. The airfoil defines a radial direction from a radially outer end to a radially inner end, and radially outer of the cooling holes spaced toward the radially outer end, which have the longer dimension extending closer to parallel to the radial direction. Radially inner cooling holes closer to the radially inner end having the longer dimension extend to be closer to perpendicular relative to the radial direction compared to the radially outer cooling holes.

Abstract: An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A tip trench is provided in the tip and wrapping at least a portion of the airfoil from the pressure side wall around the leading edge to the suction side wall. The tip trench is provided by a recess.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

Abstract: A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

Abstract: A cooling hole for a component includes a meter section and a diffuser section. The diffuser section has a footprint region defined by five sides, a first side of the five sides extending along substantially an entire height of the diffuser section and second and third sides of the five sides meeting in an obtuse angle opposite the first side. A component having the cooling hole and a method of forming the cooling hole are also disclosed.

Abstract: An airfoil according to an example of the present disclosure includes, among other things, an airfoil body defining a cavity, and a baffle including a baffle body including sidewalls and defining an internal passage for conveying coolant. The baffle body is situated in the cavity such that a majority of external surfaces of the sidewalls abut the cavity.

Abstract: In one example embodiment, a blade includes an attachment region, an airfoil extending from the attachment region, and a blade cooling arrangement. The blade cooling arrangement includes at least a first feed passage disposed through the attachment region, which is connected to a first cooling passage disposed in the airfoil. A passively actuated first coolant valve is disposed in or proximate the first feed passage. A plurality of such blades can be disposed in a turbine section of an engine.

Abstract: A gas turbine engine includes a structure that has walls that provide a cooling passage and a cooling surface. A non-ferrous obstruction is relative to the walls. The obstruction includes a portion spaced from the cooling surface to provide a gap which is configured to receive a cooling fluid.

Abstract: In one example embodiment, a blade includes an attachment region, an airfoil extending from the attachment region, and a blade cooling arrangement. The blade cooling arrangement includes at least a first feed passage disposed through the attachment region, which is connected to a first cooling passage disposed in the airfoil. A passively actuated first coolant valve is disposed in or proximate the first feed passage. A plurality of such blades can be disposed in a turbine section of an engine.

Abstract: A method of manufacturing a component that includes providing a core structure, casting a component about the core structure, removing a first portion of the core structure from the cast component, and leaving a second portion of the core structure in the cast component to provide a reduced cross-section in the cast component.

Abstract: A film cooled component may comprise a cooling chamber and a first ligament centered about a first axis. The first ligament may be in fluid communication with the cooling chamber. A first meter may be disposed at an end of the first ligament. A first diffuser may extend from the first meter to a surface of the film cooled component. The first diffuser may comprise a first tapered sidewall oriented at an angle of between 5 degrees to 15 degrees relative to the first axis. The first diffuser may further comprise a first non-tapered sidewall oriented at an angle less than 5 degrees relative to the first axis.

Abstract: A platform trailing edge seal for a turbomachine airfoil (e.g., a blade or vane) assembly includes a body configured to extend into an aft portion of a mateface gap defined between a circumferentially adjacent pair of turbomachine airfoil platforms to minimize flow from entering a blade-vane cavity through the aft portion of the mateface gap.

Abstract: In a featured embodiment, a lost core assembly includes a ceramic component having a tapered shape in a radial direction. A refractory metal component extends radially from the ceramic core component. A method of molding a gas turbine engine component is also disclosed.

Abstract: In a featured embodiment, a lost core assembly includes a ceramic component having a tapered shape in a radial direction. A refractory metal component extends radially from the ceramic core component. A method of molding a gas turbine engine component is also disclosed.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.