Abstract: A rotor assembly has: blades having airfoils and roots protruding from platform segments; a rotor disc having a peripheral face defining recesses, and slots, a recess located between two adjacent ones of the slots and bounded by a step; feather seals located radially between the peripheral face and the platform segments, a feather seal having a core extending from a trailing end to a leading end and overlapping a gap defined between two platform segments and tabs protruding from the core, the tabs including: trailing tabs positioned axially outside the recess; and leading tabs, a leading tab extending from a root to a tip and having one or more of: the tip axially positioned outside of the recess; and a fillet at an intersection between the tip and an edge of the leading tab, the edge extending between the tip and the core, and facing the step.

Abstract: A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Abstract: A turbine vane for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface. The surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Abstract: An exhaust casing for a gas turbine engine comprises a shroud configured to surround an exhaust cone, and a heat shield attached to the shroud. The heat shield has a first end and a second end axially spaced apart from each other. A gap is defined radially between the shroud and the heat shield. The gap is configured to enclose at least a portion of an interface defined between at least one strut and the shroud. The exhaust casing is configured to surround the exhaust cone and configured to be connected to the exhaust cone via the at least one strut.

Abstract: A method of reducing creep in an internally cooled turbine blade, comprising: providing a radially extending intermediate wall to continuously join a localized high stress zone of a concave side wall and a convex side wall in an intermediate cooling air channel through the blade. The intermediate wall distributes stress from the localized zone to a zone of lower stress to balance the creep inducing stress and temperature more evenly.

Abstract: An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

Abstract: A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.

Abstract: The turbine housing assembly includes a housing surrounding a plurality of turbine shroud segments mounted to support members of the housing. Impingement holes extend through the housing and have outlet openings communicating with a cavity between the shroud segments and the support. A deflector rail protrudes axially away from the support members into the cavity. The deflector rail defines a flow-redirecting surface to redirect the cooling air flow from the impingement holes radially outwardly, away from the turbine shroud segments.

Abstract: A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.

Abstract: A method of reducing creep in an internally cooled turbine blade, comprising: providing a radially extending intermediate wall to continuously join a localized high stress zone of a concave side wall and a convex side wall in an intermediate cooling air channel through the blade. The intermediate wall distributes stress from the localized zone to a zone of lower stress to balance the creep inducing stress and temperature more evenly.

Abstract: An exhaust casing for a gas turbine engine comprises a shroud configured to surround an exhaust cone, and a heat shield attached to the shroud. The heat shield has a first end and a second end axially spaced apart from each other. A gap is defined radially between the shroud and the heat shield. The gap is configured to enclose at least a portion of an interface defined between at least one strut and the shroud. The exhaust casing is configured to surround the exhaust cone and configured to be connected to the exhaust cone via the at least one strut.

Abstract: An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

Abstract: An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

Abstract: The turbine housing assembly includes a housing surrounding a plurality of turbine shroud segments mounted to support members of the housing. Impingement holes extend through the housing and have outlet openings communicating with a cavity between the shroud segments and the support. A deflector rail protrudes axially away from the support members into the cavity. The deflector rail defines a flow-redirecting surface to redirect the cooling air flow from the impingement holes radially outwardly, away from the turbine shroud segments.

Abstract: A rotor blade vibration damper for a gas turbine engine includes an elongated damper body including a top portion extending longitudinally between a front end and a rear end. The top portion has a width defined between spaced apart lateral sides and is substantially flat between the front and rear ends and between the lateral sides such as to define a longitudinal plane within which the top portion lies. A front tab extends downwardly from the front end of the top portion relative to the longitudinal plane. The rear end of the top portion is flat and generally contained in the longitudinal plane. A pair of lateral tabs extends downwardly from each of said lateral sides of the top portion relative to the longitudinal plane.

Abstract: An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

Abstract: A method includes displaying an obscured first image to a first user, where the first image is associated with a first alpha-numeric string. The method also includes generating a score based on an amount of the first image revealed or not revealed and whether the first user is able to successfully identify the first alpha-numeric string. The method may further include revealing at least a portion of the first image to the first user and receiving user input defining at least the portion of the first image revealed to the first user. The user input could include one or more swipes across a touchscreen that displays the obscured first image.

Abstract: A rotor blade vibration damper for a gas turbine engine includes an elongated damper body including a top portion extending longitudinally between a front end and a rear end. The top portion has a width defined between spaced apart lateral sides and is substantially flat between the front and rear ends and between the lateral sides such as to define a longitudinal plane within which the top portion lies. A front tab extends downwardly from the front end of the top portion relative to the longitudinal plane. The rear end of the top portion is flat and generally contained in the longitudinal plane. A pair of lateral tabs extends downwardly from each of said lateral sides of the top portion relative to the longitudinal plane.

Abstract: A gas turbine engine includes a combustor having a plurality of fuel nozzles arranged circumferentially equidistant from one another about a first center. The plurality of fuel nozzles produce a plurality of pressure pulses in a fluid contained in the combustor. A turbine section is disposed downstream of the combustor and receiving the fluid from the combustor. The turbine section includes a plurality of vanes arranged circumferentially equidistant from one another about a second center. When projecting the second center onto the first center, the plurality of vanes is angularly offset in a circumferential direction relative to the plurality of fuel nozzles of a predetermined amount. The offsetting positions the vanes in flowpaths of the pressure pulses generated by the plurality of fuel nozzles.

Abstract: A turbine exhaust section comprises a turbine exhaust case having radially outer and inner ducts defining therebetween an annular exhaust portion for the hot exhaust gases, and an exhaust mixer projecting axially rearwardly from the turbine exhaust case for mixing the hot exhaust gases with a cooler bypass duct flow. The upstream end of the exhaust mixer surrounds a downstream end of the outer duct and defines therewith an axially extending overlap joint with a radial play between the outer duct and the exhaust mixer. A sliding attachment in the radial direction is provided between the outer duct and the exhaust mixer for accommodating differences in thermal growth during engine operation. The sliding attachment includes a circumferential array of sliding guides extending radially through the axially extending overlap joint. A resilient sealing ring seals the radial play between the outer duct and the exhaust mixer at the overlap joint.