Abstract: A turbine airfoil includes an internal cooling circuit between opposite pressure and suction sidewalls. The pressure sidewall includes a row of outlet slots for the circuit separated by converging partitions. Each slot includes an inlet, throat, and diverging outlet. Each of the partitions includes a step in a converging sidewall thereof for increasing divergence of the slots to diffuse coolant discharged therethrough.

Abstract: A multiple impingement cooled structure is provided having two or more stages of impingement cooling wherein the stages are arranged so as to have substantially constant cooling effectiveness.

Abstract: A flowpath wall is provided for bounding hot combustion gases in a gas turbine engine. The wall includes opposite outer and inner surfaces and a plurality of cooling holes extending therebetween. A thermal barrier coating is bonded to the outer surface and covers blind the holes thereat.

Abstract: A method for fabricating a nozzle for a gas turbine engine facilitates extending a useful life of the nozzles. The nozzle includes an airfoil. The method includes forming the airfoil to include a suction side and a pressure side connected at a leading edge and a trailing edge such that a cooling cavity and a cooling circuit are defined within the airfoil, wherein the suction side and the pressure side extend radially between a tip and a root. The method also includes forming a plurality of cooling slots within the airfoil that extend from the cooling circuit towards the airfoil trailing edge, and forming a control vane within the cooling circuit to facilitate maintaining a substantially constant cooling effectiveness within the cooling circuit.

Abstract: A flowpath wall is provided for bounding hot combustion gases in a gas turbine engine. The wall includes opposite outer and inner surfaces and a plurality of cooling holes extending therebetween. A thermal barrier coating is bonded to the outer surface and covers blind the holes thereat.

Abstract: A method for fabricating a rotor blade for a gas turbine engine facilitates reducing operating temperatures of a tip portion of the rotor blade. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip plate, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. A second tip wall is formed to extend from the rotor blade tip plate along the second sidewall.

Abstract: A method for fabricating a nozzle for a gas turbine engine facilitates extending a useful life of the nozzles. The nozzle includes an airfoil. The method includes forming the airfoil to include a suction side and a pressure side connected at a leading edge and a trailing edge such that a cooling cavity and a cooling circuit are defined within the airfoil, wherein the suction side and the pressure side extend radially between a tip and a root. The method also includes forming a plurality of cooling slots within the airfoil that extend from the cooling circuit towards the airfoil trailing edge, and forming a control vane within the cooling circuit to facilitate maintaining a substantially constant cooling effectiveness within the cooling circuit.

Abstract: A method for fabricating a rotor blade for a gas turbine engine facilitates reducing operating temperatures of a tip portion of the rotor blade. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip plate, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. A second tip wall is formed to extend from the rotor blade tip plate along the second sidewall.

Abstract: A multiple impingement cooled structure is provided having two or more stages of impingement cooling wherein the stages are arranged so as to have substantially constant cooling effectiveness.