An airfoil cooling system may include an airfoil that defines a first cooling chamber and a second cooling chamber. The first cooling chamber may be configured to receive a first cooling airflow from a first fluid source and the second cooling chamber may be configured to receive a second cooling airflow from a second fluid source that is different from the first fluid source. The airfoil may be a vane of a gas turbine engine. For example, the vane may be a first vane of a plurality of vanes of a turbine section of the gas turbine engine. The first cooling chamber may be a leading edge chamber of the first vane and the second cooling chamber may be an aft chamber of the first vane that is positioned aft of the leading edge chamber.

Various embodiments of the disclosure include turbine blades and systems employing such blades. Various embodiments include a turbine blade having: an airfoil having an airfoil shape having a nominal profile substantially in accordance with at least a portion of Cartesian coordinate values of X, Y and Z set forth in TABLE I. The Cartesian coordinate values are non-dimensional values of from 0% to 100% convertible to distances by multiplying the values by a height of the airfoil expressed in units of distance. The X and Y values are connected by smooth continuing arcs to define airfoil profile sections at each distance Z along at least a portion of the airfoil, the profile sections at the Z distances being joined smoothly with one another to form the nominal profile.

A gas turbine combustor and a gas turbine; wherein inclined surfaces are provided on inner surfaces of side walls neighboring in a circumferential direction at downstream end portions of transition pieces of combustors, the inclined surfaces being configured to increase a passage area of the transition pieces, a ratio (S/P) is from 0 to 0.2, where (P) is a pitch dimension of first stage vanes, and (S) is a circumferential dimension from an intermediate point between neighboring transition pieces to an upstream end of a first stage vane closest in the circumferential direction; and a ratio (L/P) is from 0.3 to 0.55, where (P) is the pitch dimension, and (L) is an axial dimension from a downstream end of the transition piece to the upstream end of the first stage vane.

A gas turbine engine includes a compressor section defining a compressor exit temperature, T3. The gas turbine engine also includes a combustion section and a turbine section, with the turbine section defining a turbine inlet temperature, T4. A ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, during operation of the gas turbine engine at a rated speed is less than or equal to 1.85.

A turbine article includes a substrate with a geometric surface having a multiple of divots recessed into the substrate, and a ceramic topcoat disposed over the geometric surface, the topcoat including at least a first layer having a first hardness and a second layer having a second hardness, the first hardness different than the second hardness.

A combustor assembly for a gas turbine engine includes a TOBI module that includes a TOBI housing. The TOBI housing has a slot. A vane includes a dovetail removably received within the slot. The TOBI housing includes an axially extending TOBI nozzle array. The TOBI housing includes a plenum. A cooling passage fluidly connects the plenum to the vane. The TOBI housing includes passages configured to provide cooling fluid to the vane. The TOBI housing includes a first passageway further connecting the TOBI nozzle to the plenum.

The present invention discloses a novel apparatus and methods for providing a flow of cooling air to one or more turbine nozzles or turbine blade outer air seals. The flow of cooling air is provided by an external source and regulated in order to improve turbine nozzle and air seal cooling efficiency and component life.

A component assembly for a gas turbine engine having a combustor and defining a core air flowpath includes a first wall and a second wall. The first wall is configured for at least partially defining the core air flowpath at a location downstream of a combustion chamber defined by the combustor of the gas turbine engine. Additionally, the second wall is also configured for at least partially defining the core air flowpath at a location downstream of the combustion chamber, and is located opposite the core air flowpath of the first wall. The second wall includes a plurality of fins extending towards the first wall along the radial direction and spaced along the circumferential direction.

A blade outer air seal retention assembly may comprise a forward retention ring and an aft retention ring coupled to the forward retention ring. The forward retention ring may comprise a first shiplap flange extending aft from the forward retention ring. The aft retention ring may comprise a second shiplap flange extending forward from the aft retention ring.

The present disclosure is directed to a system for mitigating rotor bow at a turbine engine. The system includes a casing circumferentially surrounding a rotor assembly in which a heat pipe is attached to the casing and extended circumferentially around the rotor assembly.