Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Siemens Power Generation, Inc.

Mikro Systems, Inc.

Christian X Campbell

Jay Alan Morrison

John Joseph Marra

Allister William James

Jay Edgar Lane

Darryl S Eng

Ching-Pang Lee

Gary Brian Merrill

Anand A Kulkarni

Douglas A Keller

Harry A Albrecht

Daniel M Eshak

Samuel R Miller, Jr

Brian J Wessell

Ramesh Subramanian

Ahmed Kamel

Kok-Mun Tham

Yevgeniy P Shteyman

Anthony L Schiavo

Steven James Vance

Eric Joseph Schroeder

Malberto Gonzalez

Daniel George Thompson

Daniel O Davies

Thomas Barrett Jackson

Bonnie D Marini

Clinton A Mayer

Dimitrios Thomaidis

Steven C Butner

Christine Taut

Paul J Gear

Stephen John Messmann

Raymond G Snider

Andrew Whalley

Eric V Carelli

Harry Patat

Jamie Meeroff

Yevgeuiy Shteyman

Andrew S Lohaus

Phillip Behling

Daniel G Thompson

A gas turbine engine blade (), including a base portion () having a cast wall, and a tip portion () attached to the base portion and having a wall () formed by an additive manufacturing process. The tip portion wall may be formed to be solid and less than 2 mm in thickness, or it may be corrugated and be greater than 2 mm in thickness. Openings () defining the wall corrugations may be semi-circular, rectangular, trapezoidal, or elliptical in cross-sectional shape. The resulting blade has lower tip mass while retaining adequate mechanical properties. The tip portion may be formed to have a directionally solidified grain structure on a base portion having an equiaxed grain structure.

Gas turbine blade with corrugated tip wall

Platforms () span between turbine blades () on a disk (). Each platform may be individually mounted to the disk by a pin attachment (). Each platform () may have a rotationally rearward edge portion () that underlies a forward portion () of the adjacent platform (). This limits centrifugal bending of the rearward portion of the platform, and provides coolant sealing. The rotationally forward edge (A,B) of the platform overlies a seal element () on the pressure side () of the forwardly adjacent blade, and does not underlie a shelf on that blade. The pin attachment allows radial mounting of each platform onto the disk via tilting () of the platform during mounting to provide mounting clearance for the rotationally rearward edge portion (). This facilitates quick platform replacement without blade removal.

Turbine blade and non-integral platform with pin attachment

A vane array adapted to be coupled to a vane carrier within a gas turbine engine is provided comprising: a plurality of elongated airfoils comprising at least a first airfoil and a second airfoil located adjacent to one another; a U-ring; first connector structure for coupling a radially inner end section of each of the first and second airfoils to the U-ring; second connector structure for coupling a radially outer end section of each of the first and second airfoils to the vane carrier; a platform extending between the first and second airfoils; and platform connector structure for coupling the platform to one of the U-ring and the vane carrier.

Vane array with one or more non-integral platforms

A turbine blade is provided comprising: a root; an airfoil comprising an external wall extending radially from the root and having a radially outermost portion; and a damping structure. The external wall may comprise first and second side walls joined together to define an inner cavity of the airfoil. The damping structure may be positioned within the airfoil inner cavity and coupled to the airfoil so as to define a tuned mass damper.

Turbine blade with tuned damping structure

An outer rim seal arrangement (), including: an annular rim () centered about a longitudinal axis () of a rotor disc (), extending fore and having a fore-end (), an outward-facing surface (), and an inward-facing surface (); a lower angel wing () extending aft from a base of a turbine blade () and having an aft end () disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (); an upper angel wing () extending aft from the turbine blade base and having an aft end () disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (); and guide vanes () disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins () may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.

Aft outer rim seal arrangement

A cooling channel (B,-) cools inner surfaces () of exterior walls () of a component (). Interior side surfaces () of the channel converge to a waist (W), forming an hourglass shaped transverse profile (). The inner surfaces () may have fins () aligned with the coolant flow (). The fins may have a transverse profile (A,B) highest at mid-width of the inner surfaces (). Turbulators () may be provided on the side surfaces () of the channel, and may urge the coolant flow toward the inner surfaces (). Each turbulator () may have a peak () that defines the waist of the cooling channel. Each turbulator may have a convex upstream side (). These elements increase coolant flow in the corners (C) of the channel to more uniformly and efficiently cool the exterior walls ().

Component having cooling channel with hourglass cross section

Bi-casting a platform () onto an end portion () of a turbine airfoil () after forming a coating of a fugitive material () on the end portion. After bi-casting the platform, the coating is dissolved and removed to relieve differential thermal shrinkage stress between the airfoil and platform. The thickness of the coating is varied around the end portion in proportion to varying amounts of local differential process shrinkage. The coating may be sprayed (A,B) onto the end portion in opposite directions parallel to a chord line () of the airfoil or parallel to a mid-platform length () of the platform to form respective layers tapering in thickness from the leading () and trailing () edges along the suction side () of the airfoil.

Turbine airfoil to shroud attachment method

A triple hook ring segment including forward, midsection and aft mounting hooks for engagement with respective hangers formed on a ring segment carrier for supporting a ring segment panel, and defining a forward high pressure chamber and an aft low pressure chamber on opposing sides of the midsection mounting hook. An isolation plate is provided on the aft side of the midsection mounting hook to form an isolation chamber between the aft low pressure chamber and the ring segment panel. High pressure air is supplied to the forward chamber and flows to the isolation chamber through crossover passages in the midsection hook. The isolation chamber provides convection cooling air to an aft portion of the ring segment panel and enables a reduction of air pressure in the aft low pressure chamber to reduce leakage flow of cooling air from the ring segment.

Cooling system for three hook ring segment

A process is provided for forming an airfoil for a gas turbine engine involving: forming a casting of a gas turbine engine airfoil having a main wall and an interior cavity, the main wall having a wall thickness extending from an external surface of the outer wall to the interior cavity, an outer section of the main wall extending from a location between a base and a tip of the airfoil casting to the tip having a wall thickness greater than a final thickness. The process may further involve effecting movement, using a computer system, of a material removal apparatus and the casting relative to one another such that a layer of material is removed from the casting at one or more radial portions along the main wall of the casting.

Process for forming a long gas turbine engine blade having a main wall with a thin portion near a tip

A turbine airfoil () with an end portion () that tapers () toward the end () of the airfoil. A ridge () extends around the end portion. It has proximal () and distal () sides. A shroud platform () is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression () on the end portion () of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material () in the bi-casting stage. These gaps are designed in combination with the taper angle () to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle () may vary from lesser on the pressure side () to greater on the suction side () of the airfoil. A collar portion () of the platform provides sufficient contact area for connection stability.

Turbine airfoil to shround attachment

A turbine blade assembly in a turbine engine. The turbine blade assembly includes a turbine blade and a first snubber structure. The turbine blade includes an internal cooling passage containing cooling air. The first snubber structure extends outwardly from a sidewall of the turbine blade and includes a hollow interior portion that receives cooling air from the internal cooling passage of the turbine blade.

Cooled snubber structure for turbine blades

An airfoil () is provided with a tip () having an opening () to a center channel (). A damping element () is inserted within the opening of the center channel, to reduce an induced vibration of the airfoil. The mass of the damping element, a spring constant of the damping element within the center channel, and/or a mounting location () of the damping element within the center channel may be adjustably varied, to shift a resonance frequency of the airfoil outside a natural operating frequency of the airfoil.

Damping element for reducing the vibration of an airfoil

A turbine airfoil (B) with a thermal expansion control mechanism that increases the airfoil camber () under operational heating. The airfoil has four-wall geometry, including pressure side outer and inner walls (B), and suction side outer and inner walls (B). It has near-wall cooling channels (F,A,F,A) between the outer and inner walls. A cooling fluid flow pattern (C,W,H) in the airfoil causes the pressure side inner wall (B) to increase in curvature under operational heating. The pressure side inner wall (B) is thicker than walls (B) that oppose it in camber deformation, so it dominates them in collaboration with the suction side outer wall (), and the airfoil camber increases. This reduces and relocates a maximum stress area () from the suction side outer wall () to the suction side inner wall (B,) and the pressure side outer wall ().

Four-wall turbine airfoil with thermal strain control for reduced cycle fatigue

A turbine airfoil (E-H) extends from a shank (E-H). A platform (E-H) brackets or surrounds a first portion of the shank (E-H). Opposed teeth () extend laterally from the platform (E-H) to engage respective slots () in a disk. Opposed teeth () extend laterally from a second portion of the shank () that extends below the platform (E-H) to engage other slots () in the disk. Thus the platform (E-H) and the shank (E-H) independently support their own centrifugal loads via their respective teeth. The platform may be formed in two portions (E-H,E-H), that are bonded to each other at matching end-walls () and/or via pins (G) passing through the shank (E-H). Coolant channels () may pass through the shank beside the pins (G).

Modular turbine airfoil and platform assembly with independent root teeth

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Orlando, FL, United States of America

Christian X Campbell