Disclosed herein is a gas turbine blade. The gas turbine blade includes a turbine blade (33) provided in a turbine, and film cooling elements (100), each including a cooling channel (110) for cooling of the turbine blade (33), an outlet (120) through which cooling air is discharged, and a plurality of ribs (130), wherein the outlet (120) extends from a longitudinally extended end of the cooling channel (110) to an outer surface of the turbine blade (33) and has a width increased from one end of the cooling channel (110) to the outer surface of the turbine blade (33), and the ribs (130) face each other on inner walls of the outlet (120).

Various embodiments provide a rotor disk assembly capable of preventing a slip between rotor disks when rotating, and capable of performing a balancing operation such that the center of gravity thereof is the same as that of a rotating shaft, and a gas turbine including the rotor disk assembly. The rotor disk assembly may include: a plurality of rotor disks disposed parallel to each other; a tie rod passing through the plurality of rotor disks and coupling the plurality of rotor disks to each other; a plurality of coupling depressions formed in each of facing surfaces of the plurality of rotor disks; and a plurality of coupling pins each having opposite ends inserted into the corresponding respective coupling depressions of the facing surfaces.

An airfoil for a turbine engine includes an array of features positioned in an interior portion of the airfoil. Each feature extends from a pressure side to a suction side. The array includes multiple radial rows (A-N) of features with the features in each row (A-N) being interspaced radially to define coolant passages therebetween. The radial rows (A-N) are spaced along a forward-to-aft direction toward an airfoil trailing edge. The coolant passages of the array are fluidically interconnected to lead a pressurized coolant toward the trailing edge via a serial impingement on to the rows of features. The coolant passages are geometrically configured to bias a coolant flow therethrough toward a first side in relation to a second side of the outer wall to effect a greater cooling of the first side than the second side.

Provided is a bearing structure including a recess portion, which is formed in at least one of a first opposed surface of a positioning pin or a second opposed surface of a through hole, and is recessed in a separating direction in which the first opposed surface and the second opposed surface separate apart from each other.

Disclosed herein is a gas turbine blade. The gas turbine blade includes a turbine blade (33) provided in a turbine, and film cooling elements (100), each including a cooling channel (110) for cooling of the turbine blade (33), an outlet (120) through which cooling air is discharged, and a plurality of ribs (130), wherein the outlet (120) extends from a longitudinally extended end of the cooling channel (110) to an outer surface of the turbine blade (33) and has a width increased from one end of the cooling channel (110) to the outer surface of the turbine blade (33), and the ribs (130) face each other on inner walls of the outlet (120).

A method of controlling an extent of a thermal barrier coating (TBC) sheet spall and a hot gas path (HGP) component are disclosed. The method provides an HGP component having a body with an exterior surface. Controlling the extent of the TBC sheet spall includes forming a TBC over a selected portion of the exterior surface of the body. The TBC includes a plurality of segments in a cellular pattern. Each segment is defined by one or more slots in the TBC, and each segment has a predefined area such that the extent of the TBC sheet spall is limited by the predefined area of each of the plurality of segments that constitute the TBC sheet spall.

A turbine wheel assembly adapted for use in a gas turbine engine includes turbine blades made from ceramic matric composite materials. The turbine blades are mounted to a disk and anti-rotation features block movement of the turbine blades around a circumference of the disk.

The present disclosure is directed to a nozzle for a turbomachine. The nozzle includes an inner side wall, an outer side wall radially spaced apart from the inner side wall, and an airfoil extending radially from the inner side wall to the outer side wall. The airfoil defines a cavity that extends radially through the nozzle. The cavity is at least partially defined by a cavity wall. The cavity wall at least partially defines a pocket in fluid communication with the cavity. A cooling passage is defined by one of the inner side wall or the outer side wall. The cooling passage is in fluid communication with the cavity via the pocket.

An engine component assembly for impingement cooling. The engine component assembly includes an engine first component having a cooled surface. The engine first component having a flow path on one side of the cooled surface. A second component is a disposed adjacent to the engine first component between the flow path and the engine first component, and has a plurality of openings forming an array through the second component. The cooling flow path passes through the plurality of openings to cool the cooled surface. The second component having a surface facing the cooled surface of the engine first component. A plurality of discrete cooling features that have at least one wall that has a curved cross-section extend from the second component surface into a gap between and toward the cooled surface of the engine first component and defining an array.

A turbomachine blade assembly including a turbomachine blade (1), in particular for a gas turbine, and at least one tuning element container including a housing (10) attached to the turbomachine blade and an insert (20) disposed in a recess (11) of this housing. A wall (20; 21) of the insert spaces apart two first cavities (31), which each accommodate at least one tuning element (40) provided for impacting contact with the housing (10) and the insert (20).