General Electric Company

Siemens Energy, Inc.

Mikro Systems, Inc.

Chander Prakash

Ching-Pang Lee

David Glenn Cherry

Matthew D Montgomery

Robert John Beacock

Scott Andrew Burton

Joseph Machnaim

Steven Robert Brassfield

Aspi Rustom Wadia

Kok-Mun Tham

John Howard Starkweather

Reinhard Schilp

Scott Michael Carson

Ronald D Zerkle

Ronald Scott Bunker

Gary Brian Merrill

Paul Hadley Vitt

David Allen Little

Brian Keith

Monty Lee Shelton

Andrew Jeremiah Burns

Vincent Paul Laurello

Kevin Samuel Klasing

Leslie Eugene Leeke, Jr

Mark Edward Stegemiller

Craig Miller Kuhne

Hardev Singh

Gerard Anthony Rinck

Edward F Pietraszkiewicz

Melissa Harris

Stephen R Williamson

Andreas Heselhaus

Danning You

A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume.

Cooled conduit for conveying combustion gases

A resonance chamber () has an outer wall () with coolant inlet holes (A-C), an inner wall () with acoustic holes (), and side walls (A-C) between the inner and outer walls. A depression (A-C) in the outer wall has a bottom portion () that is close to the inner wall compared to peaks (A-C) of the outer wall. The coolant inlet holes may be positioned along the bottom portion of the depression and along a bottom portion of the side walls to direct coolant flows () toward impingement locations () on the inner wall that are out of alignment with the acoustic holes. This improves impingement cooling efficiency. The peaks (A-C) of the outer wall provide volume in the resonance chamber for a target resonance.

Damping resonator with impingement cooling

A wall structure () with layers (A, B, C, D, E) of non-random voids (A,B,B,B) that interconnect to form discretely defined tortuous passages between an interior () and an exterior surface () of the wall for transpiration cooling of the wall. A coolant flow () through the wall may be metered by restrictions in coolant outlets () and/or within the passages to minimize the coolant requirement. Pockets () may be formed on the exterior surface of the wall for thermal Insulation (). The layers may be formed by lamination, additive manufacturing, or casting. Layer geometries include alternating layers (A, B, C) with different overlapping void patterns (), 3-D lattice structures (), and offset waffle structures ().

Discreetly defined porous wall structure for transpirational cooling

A seal member for effecting a seal preventing fluid flow in an axial direction through an annular space formed between two relatively moving components including a rotatable shaft and a stator structure. The seal member includes a plurality of flexible seal strips. Each seal strip includes a planar plate extending radially through the annular space and having a radially outer end supported to the stator structure and a radially inner end defining a tip portion extending widthwise in the axial direction engaged in sliding contact with a peripheral surface of the rotatable shaft. At least one of the seal strips includes a plurality of perforations extending through the seal strip and located between a leading edge and a trailing edge of the seal strip for effecting an increased flexibility of the seal strip adjacent to the tip portion.

Seal including flexible seal strips

In a gas turbine engine, a flow directing member includes a platform supported on a rotor and includes a radially facing endwall and at least one axially facing axial surface extending radially inwardly from a junction with the endwall. The flow directing member further includes an airfoil extending radially outwardly from the endwall and a fluid flow directing feature. The fluid flow directing feature includes a groove extending axially into the axial surface. The groove has a radially inner groove end and a radially outer groove end, wherein the outer groove end defines an axially extending notch in the junction between the axial surface and the endwall and forms an opening in the endwall for directing a cooling fluid to the endwall.

Flow directing member for gas turbine engine

A film cooling apparatus with a cooling hole () in a component wall (). A first surface () of the wall is subject to a hot gas flow (). A second surface () receives a coolant gas (). The coolant flows through the hole, then downstream over the first surface (). One or more pairs of cooperating electrodes (---) generates and accelerates a plasma () that creates a body force acceleration () in the coolant flow that urges the coolant flow to turn around the entry edge () and/or the exit edge () of the cooling hole without separating from the adjacent surface (). The electrodes may have a geometry that spreads the coolant into a fan shape over the hot surface () of the component wall ().

Plasma actuator controlled film cooling

Plasma generators () in an endwall () of an airfoil () induce aerodynamic flows in directions () that modify streamlines () of the endwall boundary layer toward a streamline geometry () of a midspan region of the airfoil. This reduces vortices () generated by the momentum deficit of the boundary layer, increasing aerodynamic efficiency. The plasma generators may be arrayed around the leading edge as well as between two airfoils () in a gas turbine nozzle structure, and may be positioned at correction points () in streamlines caused by surface contouring () of the endwall. The plasma generators may be oriented to generate flow vectors () that combine with boundary layer flow vectors () to produce resultant flow vectors () in directions that reduce turbulence.

Plasma induced flow control of boundary layer at airfoil endwall

A stator vane that may be used in engine assemblies. The stator vane includes an airfoil having a first sidewall and a second sidewall that is coupled to the first sidewall at a leading edge and at a trailing edge. The airfoil extends radially from a root portion to a tip portion. Each of the leading and trailing edges includes at least one lean directional change and a plurality of sweep directional changes that are defined between the root portion and the tip portion.

Gas turbine engines including multi-curve stator vanes and methods of assembling the same

A stator vane that may be used in an engine assembly is provided. The stator vane includes an airfoil that has a first sidewall and a second sidewall, which connects to the first sidewall at a leading edge and at a trailing edge. The airfoil also includes a root portion and a tip portion. The first and second sidewalls both extend from the root portion to the tip portion. The airfoil root portion is formed with a negative lean, and the airfoil tip portion is formed with a positive lean.

Gas turbine engines including lean stator vanes and methods of assembling the same

A turbine assembly for a gas turbine engine. The turbine assembly includes at least one stator assembly including a radially inner band and at least one stator vane that extends radially outward from the inner band. The stator vane includes an airfoil having a root portion adjacent to the inner band and a tip portion. The airfoil also includes at least one lean directional change that is defined between the root portion and the tip portion. The turbine assembly also includes at least one turbine blade assembly that includes at least one rotor blade. The blade assembly is coupled in flow communication with the stator assembly such that an axial spacing is defined therebetween. The axial spacing defined adjacent to the at least one lean directional change is wider than the axial spacing defined adjacent to the root portion.

Turbine assembly for a gas turbine engine and method of manufacturing the same

An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil.

Turbine airfoil with curved squealer tip

An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil. At least a portion of the pressure-side tip wall is recessed from the pressure sidewall to define an outwardly facing tip shelf, such that the pressure-side tip wall and the tip shelf define a trough therebetween.

Turbine airfoil curved squealer tip with tip shelf

A turbine blade includes an airfoil having pressure and suction sidewalls extending between leading and trailing edges, and from root to tip. A dovetail is joined to the airfoil root at a platform. Three internal cooling circuits extend in span inside the airfoil, and each circuit includes a respective inlet channel commencing in axially adjacent alignment in the dovetail. The inlet channels twist together from the dovetail, through the platform, and into the airfoil behind the leading edge in transverse adjacent alignment across the sidewalls.

Triple circuit turbine blade

A turbine airfoil includes pressure and suction sidewalls extending in chord between leading and trailing edges and in span between a root and a tip. A septum is spaced between the sidewalls to define two cooling circuits on opposite sides of the septum which converge between the leading and trailing edges. An array of pins extends inwardly from the pressure sidewall at a discharge end of the circuits, and the pins decrease in length to conform with the converging circuit.

Converging pin cooled airfoil

A gas turbine engine blade includes pressure and suction sidewalls extending between leading and trailing edges and from root to tip. The pressure sidewall includes an inclined tip rib offset therein by a ramp defining a tip notch having compound inclinations.

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Cincinnati, OH, United States of America

Chander Prakash