General Electric Company

Martin C Hemsworth

Christopher Charles Glynn

Ian David Wilson

Wu-Yang Tseng

Sidney B Elston, Iii

Joey L Nelson

Robert J Corsmeier

Ambrose A Hauser

Dean Thomas Lenahan

Roger C Walker

Kevin Joseph Barb

Albert Myers

James C Przytulski

Gene David Palmer

Diether E Carreno

Susan Marie Hyde

Ming Cheng Li

Thomas Charles Mashey

Philip Michael Caruso

Frederick W Tegarden

Frederick Martin Miller

Charles H Gay, Jr

Ronald Richard Wesorick

Albert P Sterman

Nigel M Bond

The steam cooling circuit for a gas turbine includes a bore tube assembly supplying steam to circumferentially spaced radial tubes coupled to supply elbows for transitioning the radial steam flow in an axial direction along steam supply tubes adjacent the rim of the rotor. The supply tubes supply steam to circumferentially spaced manifold segments located on the aft side of the - spacer for supplying steam to the buckets of the first and second stages. Spent return steam from these buckets flows to a plurality of circumferentially spaced return manifold segments disposed on the forward face of the - spacer. Crossover tubes couple the steam supply from the steam supply manifold segments through the - spacer to the buckets of the first stage. Crossover tubes through the - spacer also return steam from the buckets of the second stage to the return manifold segments. Axially extending return tubes convey spent cooling steam from the return manifold segments to radial tubes via return elbows.

Steam cooling system for a gas turbine

In a compressor rotor assembly including a plurality of rotor disks affixed to a rotor body, and wherein at least one radially directed cooling tube is supported from a radially outer end within a rotor shaft, with an opposite, radially inner free end extending toward a longitudinal axis of the rotor, an improvement comprising a ring member loosely secured to the tube at the free end.

Cooling air tube vibration damper

A rotor cooling circuit includes a cooling medium supply reservoir for supplying cooling medium to a plurality of axial passages through the turbine wheel and spacer disc rims. A passageway communicates cooling medium from the reservoir for flow in series and in a serpentine manner in and about the cavities between the wheels and discs. The flows are reunited at a plenum for flowing the cooling medium into first and second stage buckets. Return passages through the rims of the wheels and spacer discs return spent cooling medium from the buckets to a return location external of the rotor. Transition elements between the buckets and spacer plates in the rotor rim accommodate any thermal mismatch between the adjoining buckets and spacer plates to prevent cooling medium leakage at the joints.

Closed or open circuit cooling of turbine rotor components

An assembly of a plurality of composite blades including blade roots carried by blade root receiving slots in the rotor wherein the slot has a slot wall with a radially outward portion which, when assembled, diverges from a spaced apart juxtaposed blade root pressure face radially outer surface in an amount which is a function of a predetermined amount of centrifugal loading on the blade during operation of the assembly, to allow at least a portion of the radially outer surface of the root pressure face to be in contact with the slot wall radially outward surface during operation.

Assembly of a composite blade root and a rotor

A rotor assembly for supporting rotor blades includes a plurality of axially adjoining discrete disks each having a rim, web, and hub. The rims include axial dovetail grooves for collectively supporting a respective blade dovetail therein. Each of the disks includes at least one axially extending annular arm joined integrally with the disk and fixedly joined to an adjacent one of the arms. The disks are fixedly joined together solely by the blade dovetails in the rim dovetail grooves and by the arms so that upon a crack failure of one of the disks, centrifugal load from the failed disk is transferred to an adjacent disk.

Multiple rotor disk assembly

Apparatus for preloading an airfoil blade attached to a rotor disk in a gas turbine engine. The dovetail root of a composite fan blade is preloaded to prevent rubbing and wear against the dovetail slot of a metal rotor disk during fan windmilling, with the aircraft engine off, due to wind and breezes at the airport. A resilient material is inserted into the lower portion of the slot below the engaged blade root. The disk has a radius which intersects the slot. The material is acted upon by a device which adjustably compresses the inserted material perpendicular to the radius of the disk. This causes the compressed material to exert a radially outward force against the bottom of the engaged root of the blade.

Apparatus for preloading an airfoil blade in a gas turbine engine

An aircraft propeller blade is constructed by forming two shells of composite material laminates and bonding the two shells to a metallic spar with foam filler pieces interposed between the shells at desired locations. The blade is then balanced radially and chordwise.

Method of making counterrotating aircraft propeller blades

A propulsor blade for an aircraft engine includes an airfoil section formed in the shape of a scimitar. A metallic blade spar is interposed between opposed surfaces of the blade and is bonded to the surfaces to establish structural integrity of the blade. The metallic blade spar includes a root end allowing attachment of the blade to the engine.

Counterrotating aircraft propulsor blades

A means for improving the aerodynamic efficiency of the compressor of an axial flow turbomachine is disclosed. The compressor includes a first airfoil relatively rotatable with respect to a radially disposed surface and a second airfoil, aft of the first airfoil, and fixed with respect to the surface. The surface bounds a flowpath for aft moving fluid. The surface has a circumferentially extending recess radially disposed relative to the airfoils. The recess has a generally aft facing wall, a generally axially directed wall, and a generally forward facing wall. The aft facing wall is oriented so as to provide a barrier to the forward flow of fluid in the clearance between airfoil and surface. The forward facing wall is oriented so as to provide an aerodynamically smooth transition from the recess into the flowpath.

Compressor casing recess

A turbine shroud includes a ceramic sealing layer secured to a metal substrate. In one form, the metal substrate includes a plurality of pegs extending therefrom. Intermediate bonding layers are disposed on the peg-metal substrate structure. A ceramic sealing layer of zirconium oxide with about 20 weight percentage magnesium oxide is disposed, e.g., plasma sprayed, on the intermediate bonding layers. The ceramic sealing layer includes an ordered pattern of very fine cracks therein which reduce the thermal stress in the ceramic sealing layer. A method of constructing a turbine shroud structure is also disclosed.

Metal-ceramic turbine shroud and method of making the same

A high temperature turbine shroud assembly for an aircraft gas turbine engine comprising a plurality of arcuate ceramic blocks arranged in a ring. The blocks are supported by a pair of annular support members which retain the blocks in the assembly. Spring-like fingers associated with the support members apply a preloaded force against a pair of slanted dovetail surfaces on the blocks to force the blocks against rigid stops formed upon the annular support members, thereby establishing their proper radial position within the assembly. The preload applied by the fingers accommodates the thermal expansion differences between the ceramic blocks and the metallic support members. In an alternative embodiment, the blocks are retained between cooperating rigid legs extending from the annular support members and wedged radially inwardly into position by a preloaded spring contacting the back side of the blocks. Heat transfer between the blocks and metallic support members is minimized by providing point or line contact between components, thus reducing the contact area available for heat conduction.

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Cincinnati, OH, United States of America

Martin C Hemsworth