Abstract: A rim seal between a stator and a rotor in a gas turbine engine is described. The seal includes an axial overlap between platform rims of the stator and the rotor. The axial length of the platform rim of one of the stator and the rotor is accommodated within a radially aligned trench defined in the other of the stator and the rotor.

Abstract: A rim seal located between a stator and a rotor in a gas turbine engine includes an axial overlap between platform rims of the stator and the rotor. The axial length of the platform rim of one of the stator and the rotor is accommodated within a radially aligned trench defined in the other of the stator and the rotor.

Abstract: A power turbine includes a second stage blade having an airfoil with a cold un-coated nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.

Abstract: A turbine blade of a gas turbine engine is described which includes an airfoil extending away from the hub platform to a blade tip. The airfoil defines a leading edge, a trailing edge, and a span-wise length extending between the platform and the blade tip. A hot-corrosion-resistant coating is located on the leading edge of the airfoil within a radially inner portion thereof, the radially inner portion extending away from the hub platform a desired distance along the span-wise length.

Abstract: A turbine shroud segment comprises a body having an upstream end portion and a downstream end portion relative to a flow of gases through the gas path. A core cavity is defined in the body and extends axially from the upstream end portion to the downstream end portion. A plurality of cooling inlets is defined in the upstream end portion of the body for feeding coolant in the core cavity. A plurality of cooling outlets is defined in the downstream end portion of the body for discharging coolant from the core cavity. Pedestals are provided in the core cavity.

Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.

Abstract: A mid-turbine frame of a gas turbine engine has a structural ring assembly comprising an outer ring, an inner ring having a plurality of threaded bosses extending from a radially outer surface thereof, and a corresponding number of structural spokes interconnecting the inner ring to the outer ring. Each spoke has a radially inner threaded end threadably engaged in an associated one of the threaded bosses on the inner ring.

Abstract: A turbine shroud segment comprises a body having an upstream end portion and a downstream end portion relative to a flow of gases through the gas path. A core cavity is defined in the body and extends axially from the upstream end portion to the downstream end portion. A plurality of cooling inlets is defined in the upstream end portion of the body for feeding coolant in the core cavity. A plurality of cooling outlets is defined in the downstream end portion of the body for discharging coolant from the core cavity. Pedestals are provided in the core cavity.

Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.

Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.

Abstract: A power turbine includes a second stage blade having an airfoil with a cold un-coated nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Abstract: A turbine blade for a gas turbine engine comprises a platform, a blade root, an airfoil portion defining pressure and suction sides, and a shroud provided at a tip of the airfoil portion opposite to the blade root. The shroud includes a body having a radially outer face opposite to the airfoil portion, upstream and downstream generally parallel fins extending outwardly from the outer face, and two ribs extending outwardly from the outer face. Each of the fins has an end disposed toward the pressure side and another end disposed toward the suction side. The ribs extend from and connecting the upstream fin to the downstream fin at locations other than the ends of the upstream and downstream fins. The ribs converge toward the downstream fin at an angle of between about 10 and about 45 degrees. A shroud for a blade is also presented.

Abstract: A turbine shroud segment forms part of a circumferentially segmented shroud assembly surrounding a circumferential array of turbine blades of a gas turbine engine. The turbine shroud segment has a platform extending axially from an upstream end portion to a downstream end portion relative to a flow of gas through the gas turbine engine. A heat shield extension projects radially inwardly and in an upstream direction from the upstream end portion of the platform.

Abstract: A gas turbine engine comprises a shaft assembly including a hollow shaft of the gas turbine engine and a plug connected to the inlet end of the shaft. The hollow shaft has a shaft bore having a bore diameter. The hollow shaft has an inlet end for receiving a first portion of an incoming air flow. The plug has a plug bore therethrough, and an inlet end having an inlet diameter. The inlet diameter of the plug is smaller than the bore diameter. The plug includes a deflection surface adapted to deflect a second portion of the incoming air flow away from the shaft bore. A plug for connecting to an end of a hollow shaft of a gas turbine engine and s method of controlling a flow of fluid through a shaft having a bore therethrough of a gas turbine engine are also presented.

Abstract: A rim seal between a stator and a rotor in a gas turbine engine is described. The seal includes an axial overlap between platform rims of the stator and the rotor. The axial length of the platform rim of one of the stator and the rotor is accommodated within a radially aligned trench defined in the other of the stator and the rotor.

Abstract: A gas turbine engine has an engine case and a circumferential array of rotor blades, rotatable about a centerline. A stator vane assembly is located within the engine case and axially spaced from the array of rotor blades. The stator vane assembly is formed by a plurality of stator vane segments, disposed circumferentially one adjacent to another. Each stator vane segment has an outer endwall, a plurality of vanes extending radially from the outer endwall towards the centerline and a shroud segment extending axially from the outer endwall. The shroud segment is configured to extend to and surround the array of rotor blades. The shroud segment has an abradable portion. The vane assembly is secured relative to the engine case.

Abstract: A gas turbine engine comprises a shaft assembly including a hollow shaft of the gas turbine engine and a plug connected to the inlet end of the shaft. The hollow shaft has a shaft bore having a bore diameter. The hollow shaft has an inlet end for receiving a first portion of an incoming air flow. The plug has a plug bore therethrough, and an inlet end having an inlet diameter. The inlet diameter of the plug is smaller than the bore diameter. The plug includes a deflection surface adapted to deflect a second portion of the incoming air flow away from the shaft bore. A plug for connecting to an end of a hollow shaft of a gas turbine engine and s method of controlling a flow of fluid through a shaft having a bore therethrough of a gas turbine engine are also presented.

Abstract: An integrated shroud structure surrounds a circumferential array of stator vanes and a circumferential array of rotor blades of a gas turbine engine. The shroud structure includes a plurality of vane shroud segments and a plurality of blade shroud segments. The blade shroud segments integrally extend downstream from the vane shroud segments and each pair of circumferentially adjacent blade shroud segments defines an inter-segment gap. At least one slot extends axially from a location downstream of the vane shroud segments to an aft end of the blade shroud segment. The inter-segment gaps and slots are sealed by a sealing band mounted around the full circumference of the integrated shroud structure.

Abstract: A gas turbine engine comprises a shaft assembly including a hollow shaft of the gas turbine engine and a plug connected to the inlet end of the shaft. The hollow shaft has a shaft bore having a bore diameter. The hollow shaft has an inlet end for receiving a first portion of an incoming air flow. The plug has a plug bore therethrough, and an inlet end having an inlet diameter. The inlet diameter of the plug is smaller than the bore diameter. The plug includes a deflection surface adapted to deflect a second portion of the incoming air flow away from the shaft bore. A plug for connecting to an end of a hollow shaft of a gas turbine engine and s method of controlling a flow of fluid through a shaft having a bore therethrough of a gas turbine engine are also presented.