General Electric Company

Robert Alexander Nicoll

John W Vdoviak

Jason David Shapiro

Daniel Edward Demers

Tyler Frederick Hooper

Robert Patrick Tameo

Robert E Gladden

Dennis C Evans

Douglas Patrick Probasco

A method for assembling a gas turbine engine includes coupling a rotor assembly including a plurality of rotor blades about a rotatable main shaft of the gas turbine engine. The main shaft is aligned in an axial direction of the gas turbine engine. A shroud assembly is coupled to the gas turbine engine. The shroud assembly includes a plurality of shroud segments circumferentially coupled about the rotor assembly such that a shroud spacing gap is formed in the axial direction between adjacent shroud segments. A cooling fluid source is coupled to each shroud segment such that cooling fluid is channeled through each shroud segment into a corresponding shroud spacing gap to facilitate positive purge flow through the shroud spacing gap.

Turbine shroud assembly and method for assembling a gas turbine engine

The present invention provides coolable exhaust nozzle flap and seal assemblies having liners that are spaced inward and apart from their respective flap and seal cool walls, to form a cooling passage therebetween that converges in the axial downstream flow direction. One embodiment provides single wall multi-hole film cooled nozzle flap and seal liners particularly for use individually or in combination on convergent and divergent hollow flaps and seals for variable exhaust nozzles. Flap and seal coolable assemblies may be unitary structures that are integrally fabricated or formed such as by casting.

Cooling apparatus for aircraft gas turbine engine exhaust nozzles

The present invention provides a cooling apparatus with two spaced apart walls, one of which is a coolable wall, having a cooling flowpath between them wherein the cooling flowpath is converging in the direction of the cooling flow. In one embodiment the coolable wall is an afterburner cooling liner that includes a film cooling means on the hot side of the liner. Another embodiment provides a corrugated or wavy wall afterburner cooling liner wherein each wave includes a forward facing surface, from crest to trough of the wave in the direction of the hot gas flow, and an aft facing surface, from trough to crest. A high density pattern of multi-hole film cooling holes is disposed on the hotter of the two surfaces and a lower density pattern on the cooler one.

Multi-hole film cooled afterburner combustor liner

Multi-hole film cooled afterburner cumbustor liner

The turbine blades of a high pressure turbine in a gas turbine jet engine are cooled by cooling air injected into cooling channels in the blade using angled nozzles supplied with high pressure air from the engine compressor. The nozzles project the high pressure air across a lower-pressure gap into slots which lead to internal channels in the turbine disk without requiring the use of seals to produce a pressure differential to encourage the flow of cooling air. Maximum air transfer into the internal channels in the turbine blades occurs when the air impacting slots in the turbine wheel is travelling at an angle with respect to the slots which is approximately parallel to the axes of the internal channels. A relatively narrow range of angular misalignment for aiming the nozzles is permitted. The pressure of the air from the plenum may be reduced by employing two or more flow channels in series before projecting the air across the gap to the turbine wheel.

Growing community of inventors

Beverly, MA, United States of America

Robert Alexander Nicoll