A turbine blade that allows an improvement in torque and power, and a turbine and gas turbine including the same are provided. The turbine blade includes an airfoil having a suction side and a pressure side, a platform coupled to a bottom of the airfoil, and a root protruding downward from the platform and coupled to a rotor disk, wherein the airfoil includes a cooling passage formed therein and a discharge hole connected to an upper portion of the cooling passage to discharge cooling air, and the discharge hole is inclined toward a tip of the turbine blade while extending from an inside to an outside thereof.

An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

The present embodiments set forth a blade including an airfoil including an outer tip having a floor, a leading edge and a trailing edge, a concave pressure sidewall and a convex suction sidewall extending axially between corresponding leading and trailing edges and radially between the floor and the tip cap. The airfoil further includes a tip cap extending from the floor of the outer tip and coextensive with the pressure sidewall and suction sidewall and around the leading edge and trailing edge. The tip cap includes a squealer tip configuration including a suction side tip cap portion and a pressure side tip cap portion. The suction side tip cap portion and pressure side tip cap portion extend unequal distances above the floor providing for cooling fluid flow out of the tip cap.

An airfoil of a gas turbine engine includes a leading edge and an opposed trailing edge defining a chord between the leading edge and the trailing edge, wherein the chord has a chord length. A concave surface is between the leading edge and the trailing edge, which includes a first portion proximal the leading edge of the airfoil and a second portion proximal the trailing edge of the airfoil, wherein the first portion of the concave surface includes about 10% to about 50% of the chord length. An erosion-resistant ceramic, cermet or intermetallic coating is on the second portion of the concave surface, which includes a coating leading edge pattern. The first portion of the concave surface is free of the erosion-resistant coating.

The present invention provides a gas turbine nozzle capable of reducing stress related to thermal elongation caused by a rise in gas turbine nozzle temperature and thus reducing stress produced when thermal deformation occurs in the gas turbine nozzle. The gas turbine nozzle according to the present invention includes nozzles formed integrally through an inner perimeter end wall and an outer perimeter end wall. The inner perimeter end wall has an upstream connection portion and a downstream connection portion. The upstream connection portion extends radially inward to be connected to an inner perimeter diaphragm. The downstream connection portion is located downstream from the upstream connection portion and extends radially inward to be connected to the inner perimeter diaphragm. The inner perimeter end wall has a thin-walled portion in a rear edge portion of the inner perimeter end wall, the thin-walled portion corresponding to a reduced wall thickness portion of the rear edge portion of the inner perimeter end wall.

Provided are a method for designing a vane, which can reduce peaks of secondary flow losses appearing locally in secondary flow regions and a vane obtained by the designing. The method for designing a vane includes: a step of determining a base vane formed by stacking profiles having airfoil shapes in a spanwise direction along a stacking line which is configured as a smooth curved line having no inflection point or a straight line; and a step of changing the stacking line of the base vane to a smooth wavy curved line which waves in an axial direction of a fan, a compressor or a turbine and has no elbows.

A turbomachine blade includes a root intended to be mounted in a recessed pocket of a turbomachine rotor disc, the root having two side faces extending radially and longitudinally, each of the side faces including at least one seating intended to be inserted against a side wall of the recessed pocket, along an axis of longitudinal direction and to be in contact with the side walls of the recessed pocket, a radially inner face intended to face the bottom of the recessed pocket when the root is mounted in the recessed pocket, the radially inner face connecting the two side faces, the radially inner face including at least one curved concave surface and two curved convex surfaces, the curved concave surface extending from one end of each of the two curved convex surfaces.

The present embodiments set forth a blade including an airfoil including an outer tip having a floor, a leading edge and a trailing edge, a concave pressure sidewall and a convex suction sidewall extending axially between corresponding leading and trailing edges and radially between the floor and the tip cap. The airfoil further includes a tip cap extending from the floor of the outer tip and coextensive with the pressure sidewall and suction sidewall and around the leading edge and trailing edge. The tip cap includes a squealer tip configuration including a suction side tip cap portion and a pressure side tip cap portion. The suction side tip cap portion and pressure side tip cap portion extend unequal distances above the floor providing for cooling fluid flow out of the tip cap.

A combustor includes a liner having an outlet end to pass combustion gas and a liner flange protruding outward from the outlet end; a transition piece to discharge combustion gas from the liner to a turbine, the transition piece having an inlet end for coupling to the outlet end of the liner and a transition piece flange protruding outward from the inlet end to face the liner flange; and a first elastic support installed on the liner flange to protrude toward the transition piece flange. A force applied from the transition piece elastically deforms an elastic arch of the first elastic support, which includes a movable support that is spaced apart from the liner flange if the force applied from the transition piece does not primarily deform the elastic arch. An auxiliary elastic support installed inside the first elastic support elastically deforms if the force secondarily deforms the elastic arch.

Aspects of the disclosure generally relate to a turbine center frame for a turbine engine through which a flow path extends. The turbine center frame can include an inner wall radially spaced from an outer wall, with the inner and outer walls extending between an inlet and an outlet, and with the outlet downstream of the inlet with respect to the flow path. A set of circumferentially-spaced airfoils can extend between the inner wall and the outer wall.