Westinghouse Electric Corporation

Siemens Westinghouse Power Corporation

William E North

Kent Goran Hultgren

Raymond Scott Nordlund

Zachary Sinnott

Mark T Kennedy

Augustine C McClay

John P Donlan

Thomas Walter Zagar

Gerard McQuiggan

Steven Douglas Ward

Robert Kenmer Scott

Frank A Pisz

Thomas M Szewczuk

Leslie R Southall

Gary S Van Heusden

Graham Mark Upton

James M Allen

Anthony J Malandra

Robert A Dorris

David T Entenmann

Stephen Humphrey Robbins

Frank K Gabriel

Christopher D Dishman

Paul C Holden

Paul H Davis

A method for providing high cooling effectiveness over the entire length of a cooling path (,) by injecting supplemental coolant into the path (,) at one or more selected downstream locations (,). Optimal selection of the injection location (,) and the ratio of injected flow to main flow will provide a cooling design with superior temperature uniformity and reduced coolant consumption relative to non-supplemented cooling path designs.

Method of cooling a combustion turbine

A turbine blade with a cooling air flow path specifically directed toward cooling the platform portion of the blade root. Cooling air passages are formed in the blade root platform just below its radially outward facing surface on an overhanging portion of the platform opposite the convex surface of the blade airfoil. Each of these passage extends radially outward from an inlet that receives a flow of cooling air, and then extends through the platform. Cavities are formed in a radially inward facing surface of an over hanging portion of the platform opposite the concave surface of the blade airfoil. An impingement plate directs cooling air as jets into these cavities. A passage is connected to the cavities and directs this cooling air through the overhanging portion of the platform opposite the concave surface.

Gas turbine blade platform cooling concept

Method and apparatus are disclosed for providing a closed-loop air cooling system for a turbine engine. The method and apparatus provide for bleeding pressurized air from a gas turbine engine compressor for use in cooling the turbine components. The compressed air is cascaded through the various stages of the turbine. At each stage a portion of the compressed air is returned to the compressor where useful work is recovered.

Closed-loop air cooling system for a turbine engine

A gas turbine that has turbine vanes that can be cooled with either steam or air includes a compressor that compresses air, a combustor enclosed within a shell and in flow communication with the compressor, a turbine in flow communication with the combustor and receiving hot gas from the combustor and turbine vanes. The turbine vanes having an airfoil, a shroud and a cooling circuit that cools the airfoil and the shrouds with either steam or air.

Cooling system for gas turbine vane

An apparatus for cooling the trailing edge portion of a gas turbine vane. Two radially extending passages connected to the outer shroud direct cooling fluid to a plenum formed about mid-span adjacent the trailing edge. Two arrays of cooling fluid passages extend from the plenum. One array extends radially outward toward the outer shroud while the other array extends radially inward toward the inner shroud. The plenum distributes the cooling fluid to the two arrays of passages so that it flows radially inward and outward to manifolds formed in the inner and outer shrouds. The manifolds direct the spent cooling fluid to a discharge passage.

Apparatus for cooling a gas turbine airfoil

A gas turbine vane having an inner shroud that is cooled by a portion of the cooling air directed to a cavity between two adjacent rows of discs. A portion of the cooling air in the cavity flows through impingement plates and impinges on the inner surface of the inner shroud. Another portion of the cooling air flows through a passage in the leading edge of the inner shroud that has a pin fin array for enhanced cooling. The impingement plates form chambers that collect both the impingement air and the pin fin passage air and direct it through holes in the trailing edge of the inner shroud for cooling of the trailing edge. Longitudinal passages along the side of the inner shroud direct the cooling air from the pin fin passage tot the trailing edge.

Gas turbine vane with a cooled inner shroud

A gas turbine stationary vane having an airfoil portion and inner and outer shrouds. Five serpentine radially extending cooling air passages are formed in the vane airfoil. The first passage is disposed adjacent the leading edge of the airfoil and the second passage is disposed adjacent the trailing edge. A first portion of the cooling air enters the first passage, from which it flows sequentially to the second, third, fourth and fifth passages. Additional cooling air enters the third passage directly, thereby bypassing the first and second passages and preventing over heating of the cooling air by the time it reaches the fifth passage. A radial tube extends through the second passage and directs cooling air through the airfoil, with essentially no rise in temperature, to an interstage cavity for disc cooling. Fins project into each of the passages and serve to increase the effectiveness and flow rate of the cooling air. The fins in the first and fifth passages are angled so as to direct the cooling air toward the leading and trailing edges, respectively. In addition, the fins in the second through fifth passages are angled to retard flow separation as the cooling air turns 180.degree. from one passage to the next.

Gas turbine vane with enhanced cooling

A gas turbine blade having a shroud extending outwardly from the tip of the airfoil portion of the blade. The shroud is cooled by cooling air passages formed within it. A radial cooling air supply hole directs cooling air directly from the blade root through the airfoil and to the shroud. A plurality of cooling air passages extend from the supply hole and are disposed adjacent bearing surfaces along which the shroud contacts the shroud of an adjacent blade. One of these cooling air holes is formed in the portion of the shroud that projects from the convex surface of the airfoil and another one of the cooling air holes is formed in the portion of the shroud that projects from the concave surface of the airfoil. The cooling air holes extend to the edge of the shroud and discharge the cooling through an opening in the edge.

Gas turbine blade having a cooled shroud

A gas turbine stationary vane is provided having cooling air inserts and an inpingement plate attached to the outer shroud of the vane. The inserts extend only a short distance above their mounting surface on the outer shroud so that there is adequate access for properly welding the inserts to the outer shroud. Insert extensions are disposed in the end of each insert so as to form cooling air inlets which project above the impingement plate. The insert extensions are installed after the insert has been welded to the outer shroud so as to not interfere with access to the weld area. The insert extensions are attached to the inserts and the impingement plate by flexible seal collars, allowing the collars to form positive seals between the components.

Gas turbine vane cooling air insert

An apparatus and method are provided for cooling the rotating blades in the turbine section of a gas turbine. Cooling of the leading edge is accomplished by forming a radial passageway in the leading edge portion of the blade airfoil. Rows of holes along the leading edge of the blade airfoil allow cooling air to flow from the radial passageway to the surface of the leading edge, thereby cooling the leading edge portion of the airfoil. The flow area of the radial passageway reduces as it extends radially outward so that the velocity of the cooling air flowing through the passageway is maintained. A plenum formed in the root portion of the blade distributes cooling air to small diameter holes in the center and trailing edge portion of the airfoil.

Apparatus for cooling rotating blades in a gas turbine

A cooled turbine vane in a multistage gas turbine is cooled with a fluid and a major portion of the fluid is dissipated into the gas flow path. A minor portion is supplied to a seal housing and seal for cooling the seal housing and for providing sufficiently cool fluid to the seal. The vane has a hollow airfoil body fixedly attached to the seal housing via an inner shroud and to a blade ring via an outer shroud. An inlet inot the hollow interior of the airfoil body through the outer shroud is in flow communication with a source of coolant fluid. An outlet in the inner shroud is in flow communication with the seal and its housing. Ports in the airfoil body extending from its hollow interior are in flow communication with the gas flow path. In a preferred cooled turbine vane, the coolant fluid in the airfoil body flows through return bends with reduced stagnant zons and flow resistance for increasing the heat transfer across the surfaces of the airfoil body around the bends and for reducing the pressure drop in the bend areas. A bend at the inner shroud is adjacent the outlet and a bend adjacent the outer shroud is adjacent a port for flowing the coolant fluid out of the stagnant zones in the bends.

Cooled turbine vane

A gas turbine of the type having high pressure air supplied to the cavity formed by the inner shrouds of the turbine vanes is provided with film cooling of the shrouds. A manifold supplies high pressure cooling air to portions of the gaps between inner shrouds not otherwise supplied and intermittent reliefs in the strip seal between shrouds regulates the leakage of this air, over the outer surfaces of the shrouds.

Apparatus for film cooling of turbine van shrouds

The invention comprises a cooled turbine rotor blade having an improved blade tip structure. A groove is provided in the trailing edge end of the blade tip on those turbine blades whose trailing edge is too thin to support an extension of the blade walls to form a blade tip cavity which extends to the tip of the trailing edge of the blade. The groove protects adjoining exhaust apertures from closure by a blade tip smear.

Tip structure for a cooled turbine rotor blade

A hollow, airfoil shaped stator vane 10 is provided in its trailing edge portion with staggered rows of pin fins 20, 22 and 24 and with longitudinal ribs 26 which provide the required stiffness for the thin wall design of the vane, the ribs being provided with protuberances 36 in those locations where the turbulence inducing pin fins are displaced by the presence of the ribs.

Cooling arrangement for airfoil stator vane trailing edge

A hollow stator vane for a turbine is provided with separate forward and aft inserts 26 and 32 with the forward insert providing impingement cooling of the vane walls while the aft insert is closely fitted into the vane internal cavity to provide closely defined channel widths between the insert and facing walls of the vane, the aft insert also including means functioning during operation to ensure the maintenance of the closely defined channel widths, the coolant air admitted to the inserts functioning first for impingement cooling, then channel flow cooling along the aft insert, and then channel flow cooling of the trailing edge 18 through an exit slot.

Growing community of inventors

Winter Springs, FL, United States of America

William E North