Abstract: An airfoil for a gas turbine engine includes a substrate portion extending from an airfoil leading edge to an airfoil trailing edge portion. The airfoil trailing edge portion includes a flared portion wherein a substrate portion thickness increases along a camber line of the airfoil, and a trailing edge defined as a full constant radius extending from a pressure side of the airfoil to a suction side of the airfoil. A coating portion includes a coating applied over at least a portion of the substrate portion.

Abstract: Disclosed is a blade, having: a blade root; a blade body that extends radially from the blade root to a blade tip, the blade body having a pressure side and a suction side each extending from a leading edge of the blade body to a trailing edge of the blade body; a squealer pocket located in the blade tip; and a first wearable tip-rub-portion located radially above the squealer pocket, and extending between the pressure side and the suction side, the first wearable tip-rub-portion terminating at a first wearable tip-rub-portion distal end.

Abstract: Disclosed is a blade, having: a blade root; a blade body that extends radially from the blade root to a blade tip, the blade body having a pressure side and a suction side each extending from a leading edge of the blade body to a trailing edge of the blade body; a squealer pocket located in the blade tip; and a first wearable tip-rub-portion located radially above the squealer pocket, and extending between the pressure side and the suction side, the first wearable tip-rub-portion terminating at a first wearable tip-rub-portion distal end.

Abstract: A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet, the combustor bypass passage configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. The combustor bypass valve including a valve element configured to rotate about an engine central longitudinal axis A to move one or more valve openings relative to one or more bypass passage openings.

Abstract: A die setting apparatus for a ceramic core is provided. The die setting apparatus includes a first setter abuttable with a first side of the ceramic core and a second setter abuttable with a second side of the ceramic core opposite the first side. At least one of the first and second setters includes two or more pieces respectively arranged to form one or more gaps. Each of the one or more gaps is oriented to thermally adjust in correspondence with thermal changes of the ceramic core during a firing process thereof.

Abstract: A gas turbine engine blade according to an example of the present disclosure includes a blade including internal walls defining a blade cooling passage and a tip section at a radially outer end of the blade. The tip section includes a pocket protruding into the tip section from a radially outermost end of the tip section to a pocket floor. The blade further includes a pocket wall defined between the pocket floor and the radially outermost end of the tip section and including a tapered wall cooling passage.

Abstract: A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet. The combustor bypass passage is configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. A method of operating a gas turbine engine, includes urging a core airflow from a compressor section toward a combustor section, flowing a first portion of the core airflow into the combustor section via a combustor inlet, and flowing a second portion of the core airflow into a combustor bypass passage via a combustor bypass valve, thereby bypassing the combustor with the second portion of the core airflow.

Abstract: A nozzle segment for a gas turbine engine includes an arcuate outer vane platform segment. An arcuate inner vane platform segment spaced from the arcuate outer vane platform segment. A multiple of airfoils between the arcuate inner vane platform segment and the arcuate outer vane platform segment, the arcuate outer vane platform segment includes a scallop slot and a seal that seals the scallop slot.

Abstract: A die setting apparatus is provided for a ceramic core is provided. The die setting apparatus includes a first setter abuttable with a first side of the ceramic core and a second setter abuttable with a second side of the ceramic core opposite the first side. At least one of the first and second setters includes two or more pieces respectively arranged to form one or more gaps. Each of the one or more gaps is oriented to thermally adjust in correspondence with thermal changes of the ceramic core during a firing process thereof.

Abstract: A gas turbine engine blade according to an example of the present disclosure includes a blade including internal walls defining a blade cooling passage and a tip section at a radially outer end of the blade. The tip section includes a pocket protruding into the tip section from a radially outermost end of the tip section to a pocket floor. The blade further includes a pocket wall defined between the pocket floor and the radially outermost end of the tip section and including a tapered wall cooling passage.

Abstract: An airfoil for a gas turbine engine includes a substrate portion extending from an airfoil leading edge to an airfoil trailing edge portion. The airfoil trailing edge portion includes a flared portion wherein a substrate portion thickness increases along a camber line of the airfoil, and a trailing edge defined as a full constant radius extending from a pressure side of the airfoil to a suction side of the airfoil. A coating portion includes a coating applied over at least a portion of the substrate portion.

Abstract: A rim seal for a rotor of a gas turbine engine includes a seal portion extending circumferentially across a rim cavity of a rotor, the sealing portion configured to seal the rim cavity and a first foot portion extending radially inwardly from a first end of the sealing portion. A rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades secured to the rotor disc defining a rim cavity between the rotor disc and a rim portion of the plurality of rotor blades. A rim seal is located in the rim cavity and includes a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity. The seal portion has an increasing radial thickness with increasing distance from a first end of the rim seal and from a second end opposite the first end.

Abstract: A baffle assembly includes a collar with a port and a baffle positioned in the port. The baffle includes two ends, a hollow body, and a weld connecting the baffle to the collar, and the collar continuously surrounds the body of the baffle.

Abstract: A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet, the combustor bypass passage configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. The combustor bypass valve including a valve element configured to rotate about an engine central longitudinal axis A to move one or more valve openings relative to one or more bypass passage openings.

Abstract: A combustor section of a gas turbine engine includes a combustor having a combustor inlet, and a combustor bypass passage having a passage inlet located upstream of the combustor inlet. The combustor bypass passage is configured to divert a selected bypass airflow around the combustor. A combustor bypass valve is located at the combustor bypass passage to control the selected bypass airflow along the combustor bypass passage. A method of operating a gas turbine engine, includes urging a core airflow from a compressor section toward a combustor section, flowing a first portion of the core airflow into the combustor section via a combustor inlet, and flowing a second portion of the core airflow into a combustor bypass passage via a combustor bypass valve, thereby bypassing the combustor with the second portion of the core airflow.

Abstract: A vane seal assembly for a gas turbine engine comprises of a case including a first connector. A notch in the case adjoins the groove. A vane having a second connector mates with the first connector. A seal assembly is provided between the vane and the case to provide a sealed cavity adjoining the notch.

Abstract: A rim seal for a rotor of a gas turbine engine includes a seal portion extending circumferentially across a rim cavity of a rotor, the sealing portion configured to seal the rim cavity and a first foot portion extending radially inwardly from a first end of the sealing portion. A rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades secured to the rotor disc defining a rim cavity between the rotor disc and a rim portion of the plurality of rotor blades. A rim seal is located in the rim cavity and includes a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity. The seal portion has an increasing radial thickness with increasing distance from a first end of the rim seal and from a second end opposite the first end.

Abstract: A baffle assembly includes a collar with a port and a baffle positioned in the port. The baffle includes two ends, a hollow body, and a weld connecting the baffle to the collar, and the collar continuously surrounds the body of the baffle.

Abstract: The present disclosure relates generally to a transducer position guide. The transducer position guide may be attached to a component to be measured, such as an airfoil, and controls a path traversed by the transducer during a measurement process.

Abstract: An airfoil for a gas turbine engine according to one exemplary embodiment includes an airfoil body that extends between a leading edge and a trailing edge. A cooling circuit can be defined within the airfoil body. The cooling circuit can include at least one trip strip disposed within a cavity of the cooling circuit between a leading edge inner wall and a first rib. The at least one trip strip can include an increasing height in a direction from the first rib toward the leading edge inner wall.