A turbine assembly adapted for use with a gas turbine engine includes an outer case, a blade track segment, and a carrier. The outer case extends circumferentially at least partway around an axis of the engine. The blade track segment is configured to define a portion of a gas path of the turbine assembly. The carrier is coupled with the outer case and the blade track segment to support the blade track segment in position radially relative to the axis. The carrier is coupled with the outer case for movement with the outer case in response to thermal expansion and contraction of the outer case during use of the turbine assembly.

A seal structure for a gas turbine blade including a shank including a first side surface, a second side surface, and a bottom surface located radially inward from a fin. The first side surface has a recess movably housing a wedge seal, and a slot on the bottom surface as an insertion opening for a spline seal. The second side surface has a slot on the bottom surface as the insertion opening. The recess has an inclined surface extending straight radially inward and away from the first side surface. The wedge seal includes a wedge portion including a first wedge surface to be opposed to the inclined surface and a second wedge surface to be opposed to the second side surface of the adjacent rotor blade to form a wedge together with the wedge surface, and a weight portion positioned radially inward in the recess from the wedge portion.

A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

A gas path component for a gas turbine engine includes a film cooling hole disposed in the gas path component. The film cooling hole includes a metering section, a diffuser, and a tapered surface extending between the metering section and the diffuser. The tapered surface is oriented between twenty degrees and seventy degrees with respect to a centerline axis of the metering section. The tapered surface is oriented at an obtuse angle with respect to an immediately adjacent surface of the diffuser, the obtuse angle is open towards the centerline axis. The tapered surface is configured to mitigate flow separation in the diffuser.

The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of vanes. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The vanes are arranged circumferentially around the rotational axis. The vanes include a first vane that is integral with the first rotor disk and projects axially to the second rotor disk.

A rotor blade in an embodiment includes: a suction surface side projecting portion projecting from a suction surface on a leading edge side at a blade tip of the blade effective portion; and a pressure surface side projecting portion projecting from a pressure surface on a trailing edge side at the blade tip of the blade effective portion. The suction surface side projecting portion includes: a leading edge side end surface including a contact surface that comes into contact with the pressure surface side projecting portion of the adjacent rotor blade and a non-contact surface that does not come into contact with the pressure surface side projecting portion of the adjacent rotor blade during rotation; a groove portion formed from the non-contact surface to the trailing edge side; and a joining member joined to the groove portion, the joining member being formed of an erosion-resistance material.

An apparatus and method for a turbine engine for can include an engine component. The engine component can include an interior cooling passage at least partially defining a cooling circuit for passing a flow of cooling fluid through the component. Film holes provide for exhausting a portion of the cooling fluid to an exterior of the component, to form a cooling film along an exterior hot surface of the engine component. A deflector can be position within the cooling passage upstream form the film hole.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.

A rotor disk for a gas turbine engine includes a disk body having a central bore extending therethrough. The disk body includes a bore body that extends around the central bore, a web that extends radially outward from the bore body having decreased thickness relative to the bore body and a peripheral rim that is located at an outer end of the web. The peripheral rim includes blade mounting structures for engaging complimentary mounting structures of rotor blades. The bore body has a bore cavity that extends continuously through the bore body and about an entire periphery of the central bore. The bore cavity has a central axis that forms a circle about the central bore.