A turbine blade includes: an airfoil portion extending in a blade height direction and having a pressure surface and a suction surface each of which extends between a leading and trailing edges; a shroud portion disposed on a blade tip side of the airfoil portion; a fillet portion formed by a curved surface and connected to an end portion of the shroud portion on a side of the airfoil portion; at least one first cooling hole extending along the blade height direction within the airfoil portion; at least one cooling cavity disposed at least partially within the shroud portion and communicating with the at least one first cooling hole; and a second cooling hole connected to the at least one cooling cavity and opening to a surface of the shroud portion. The airfoil portion has a reference airfoil in which a maximum blade thickness is minimum at a reference position.

A turbine blade includes: an airfoil portion extending in a blade height direction and having a pressure surface and a suction surface each of which extends between a leading and trailing edges; a shroud portion disposed on a blade tip side of the airfoil portion; a fillet portion formed by a curved surface and connected to an end portion of the shroud portion on a side of the airfoil portion; at least one first cooling hole extending along the blade height direction within the airfoil portion; at least one cooling cavity disposed at least partially within the shroud portion and communicating with the at least one first cooling hole; and a second cooling hole connected to the at least one cooling cavity and opening to a surface of the shroud portion. The airfoil portion has a reference airfoil in which a maximum blade thickness is minimum at a reference position.

A nozzle guide vane for a gas turbine engine having a combined side wall thickness value which varies within a cavity region so as to provide a point with a maximum value of combined side wall thickness, which is advantageous for capturing debris travelling through the engine core.

A nozzle guide vane for a gas turbine engine having a combined side wall thickness value which varies within a cavity region so as to provide a point with a maximum value of combined side wall thickness, which is advantageous for capturing debris travelling through the engine core.

A gas turbine engine assembly according to an example of the present disclosure includes, among other things, an endwall having a first material composition, an airfoil extending in a radial direction from the endwall, and a cupped contour of a second material composition that is formed on the endwall to define a cooling chamber, the first material composition different than the second material composition. A method of forming an endwall is also disclosed.

A cleaning process (blasted-particles cleaning process) includes performing, a plural number of consecutive cycles, an ultrasonic cleaning treatment including immersing a turbine rotor blade in a water basin and conducting an ultrasonic wave into the water basin to clean the turbine rotor blade, and a pressurized-water cleaning treatment including spraying pressurized water into an internal cooling flow channel after the ultrasonic cleaning treatment is performed. The cleaning process is performed after a bonding coat layer removing process of removing a bonding coat layer (first coating layer) by chemical treatment, and a cleaning process of cleaning the turbine blade by blast treatment. Heat tinging process is then performed.

A cleaning process (blasted-particles cleaning process) includes performing, a plural number of consecutive cycles, an ultrasonic cleaning treatment including immersing a turbine rotor blade in a water basin and conducting an ultrasonic wave into the water basin to clean the turbine rotor blade, and a pressurized-water cleaning treatment including spraying pressurized water into an internal cooling flow channel after the ultrasonic cleaning treatment is performed. The cleaning process is performed after a bonding coat layer removing process of removing a bonding coat layer (first coating layer) by chemical treatment, and a cleaning process of cleaning the turbine blade by blast treatment. Heat tinging process is then performed.

A turbine rotor includes a base and a plurality of blades. The base and the blades curve such that radially outward portions of the base and the blades extend in a direction with a greater component in a radial direction than in an axial direction. Radially central portions of the base and the blade extend in a direction with the two components being closer. Radially inner sections of the base and the blades extend in a direction with a greater component in the axial direction than in a radial direction. There is a cooling channel arrangement in the turbine rotor. The cooling channel arrangement includes impingement cooling for a nose and serpentine passages for cooling sections of the platform circumferentially intermediate the blades, and distinct serpentine passages for cooling the plurality of blades. A turbomachine and method are also disclosed.

An airfoil vane includes an airfoil section which includes an outer wall that defines an internal cavity. A baffle is situated in the internal cavity. The baffle includes a leading edge portion and a trailing edge portion which is bonded to the leading edge portion at a joint. The leading edge portion and the trailing edge portion define an internal cavity therewithin. Both the leading edge portion and the trailing edge portion include a plurality of cooling holes which are configured to provide cooling air to the airfoil outer wall. The trailing edge portion is formed of sheet metal and the leading edge portion is formed of non-sheet-metal. A method of making a baffle for a vane arc segment and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

An airfoil vane includes an airfoil section which includes an outer wall that defines an internal cavity. A baffle is situated in the internal cavity. The baffle includes a leading edge portion and a trailing edge portion which is bonded to the leading edge portion at a joint. The leading edge portion and the trailing edge portion define an internal cavity therewithin. Both the leading edge portion and the trailing edge portion include a plurality of cooling holes which are configured to provide cooling air to the airfoil outer wall. The trailing edge portion is formed of sheet metal and the leading edge portion is formed of non-sheet-metal. A method of making a baffle for a vane arc segment and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.