General Electric Company

Mark Douglas Gledhill

Robert Francis Manning

Jason David Shapiro

Daniel Edward Demers

Tyler Frederick Hooper

Bhanu Mahasamudram Reddy

Magdi Naim Azer

David Allen Flodman

Gaoqiu Zhu

Robert Edward Athans

Huan Qi

Gene C F Tsai

Aureen Cyr Currin

Sudhir Kumar Tewari

James Earl Kopriva

Douglas Bryan Ballantyne

Donald Brett DeSander

Joseph Giancarlo Sabato

A method of fabricating a rotor blade is provided. The method includes forming at least one passageway within the rotor blade, wherein the passageway extends substantially radially from a root of the rotor blade to a tip of the rotor blade, and coupling a shroud to the tip of the rotor blade. The shroud includes at least one substantially radially-outward extending wall that at least partially defines an outer plenum that is radially outward from at least the shroud, wherein the outer plenum is in flow communication with the passageway.

Rotor blade and method of fabricating the same

A turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert.

Hybrid impingement cooled airfoil

A turbine blade includes an airfoil having a leading edge flow chamber disposed behind a leading edge in front of a pair of side channels. The leading edge channel receives coolant from the pressure side channel in isolation from the suction side channel.

Trichannel airfoil leading edge cooling

A turbine blade assembly for a gas turbine engine that includes a turbine platform including contoured edges. The turbine blade assembly includes a platform, a turbine airfoil extending radially outward from the platform, a shank that extends radially inward from the platform, and a dovetail extending from the shank. The platform includes a leading edge and a trailing edge separated by a pressure edge and an opposite suction edge, the pressure edge including a plurality of arcs extending between the leading edge and the trailing edge, the suction edge including a plurality of arcs extending between the leading and trailing edges.

Method and apparatus for turbine blade contoured platform

A gas turbine engine nozzle segment includes at least two circumferentially adjacent vanes joined together along an interface between the vanes. Each of the vanes includes a hollow airfoil disposed between radially inner and outer band panels and a cooling air intake duct leading to a hollow interior of the airfoil for directing cooling air into the hollow interior. The intake duct has a duct wall protruding radially outward from the outer band panel and at least one impingement cooling hole disposed through the intake duct wall and circumferentially and radially inwardly angled. A flange portion extends circumferentially along the outer band panel and an opening in the flange portion is in fluid communication with the intake duct. The hollow airfoil, radially inner and outer band panels, intake duct wall, and flange portion are integrally formed and, preferably, integrally cast such that the vane is a single piece integrally cast vane.

Turbine nozzle segment band cooling

Growing community of inventors

South Hamilton, MA, United States of America

Mark Douglas Gledhill