Siemens Energy, Inc.

United Technologies Corporation

Doosan Heavy Industries & Construction Co., Ltd

Other

Matthew D Montgomery

Richard Colin Charron

John A Orosa

Chander Prakash

Ching-Pang Lee

Jay Alan Morrison

David James Wiebe

Jose L Rodriguez

Gary D Snyder

Jaewook Song

Jerry W Wood

Kok-Mun Tham

Paul Hadley Vitt

Alexander Ralph Beeck

Andrew Bradford Green

Raymond Scott Nordlund

William W Pankey

Clinton A Mayer

Bruce Loren Smith

Jody W Wilson

Ali Akturk

Daniel J Pierce

Marc Samit

Bernhard W Küsters

Krishna C Veluru

Benjamin G Hettinger

David L Wasdell

Sriram Srinivasan

William Brian Diggs

Peter D Silkowski

Ernie B Campbell

John Orosa

Caterine Meza

Stephen R Williamson

Melissa Harris

W Kerry Byrne

Rachel J Block

Compressor components, such as blades and vanes, having an airfoil portion with an uncoated, nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each Z distance in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil shape.

Airfoil profile

An inlet flow structure can comprise: a plurality of bellmouth elements, each of the plurality of bellmouth elements including a bellmouth and a diffuser; a plurality of supply plenums, each of the plurality of supply plenums surrounding a respective one of the plurality of bellmouth elements, wherein the plurality of supply plenums are divided by a wall which provides equivalent air flow to each bellmouth element. The inlet flow structure comprises an inlet plenum disposed on the plurality of supply plenums and an outlet plenum disposed on the plurality of bellmouth elements.

Inlet flow measurement structure for an industrial gas turbine

An arrangement to control vibrations in a gas turbine exhaust diffuser is provided. The arrangement includes a protrusion coupled to a turbine exhaust cylinder strut for controlling shock induced oscillations in a gas turbine diffuser. The controlled shock induced oscillations minimize pressure fluctuations in the gas turbine exhaust diffuser such that an unsteadiness of the fluid flow surrounding the turbine exhaust cylinder strut is reduced. A method to fluid flow induced vibrations in a gas turbine diffuser is also provided.

Turbine exhaust cylinder strut strip for shock induced oscillation control

A duct arrangement in a can annular gas turbine engine. The gas turbine engine has a gas delivery structure for delivering gases from a plurality of combustors to an annular chamber that extends circumferentially and is oriented concentric to a gas turbine engine longitudinal axis for delivering the gas flow to a first row of blades A gas flow path is formed by the duct arrangement between a respective combustor and the annular chamber for conveying gases from each combustor to the first row of turbine blades The duct arrangement includes at least one straight section having a centerline that is misaligned with a centerline of the combustor.

Gas flow path for a gas turbine engine

A gas turbine engine can-annular combustion arrangement (), including: an axial compressor () operable to rotate in a rotation direction (); a diffuser () configured to receive compressed air () from the axial compressor; a plenum () configured to receive the compressed air from the diffuser; a plurality of combustor cans () each having a combustor inlet () in fluid communication with the plenum, wherein each combustor can is tangentially oriented so that a respective combustor inlet is circumferentially offset from a respective combustor outlet in a direction opposite the rotation direction; and an airflow guiding arrangement () configured to impart circumferential motion to the compressed air in the plenum in the direction opposite the rotation direction.

Swirling midframe flow for gas turbine engine having advanced transitions

An industrial gas turbine engine (), including: a can annular combustion assembly (), having a plurality of discrete flow ducts configured to receive combustion gas from respective combustors () and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for delivery directly onto the first row () of turbine blades (); and a compressor diffuser () having a redirecting surface () configured to receive an axial flow of compressed air and redirect the axial flow of compressed air radially outward.

Gas turbine engine with radial diffuser and shortened mid section

An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge.

Airfoil assembly including vortex reducing at an airfoil leading edge

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

An exhaust diffuser system and method for a turbine engine includes an inner boundary and an outer boundary with a flow path defined therebetween. The inner boundary is defined at least in part by a hub structure that has an upstream end and a downstream end. The outer boundary may include a region in which the outer boundary extends radially inward toward the hub structure and may direct at least a portion of an exhaust flow in the diffuser toward the hub structure. The hub structure includes at least one jet exit located on the hub structure adjacent to the upstream end of the tail cone. The jet exit discharges a flow of gas substantially tangential to an outer surface of the tail cone to produce a Coanda effect and direct a portion of the exhaust flow in the diffuser toward the inner boundary.

Turbine exhaust diffuser with a gas jet producing a coanda effect flow control

An electrode () in the tip () of a turbine or compressor blade (), and a series of electrodes () in a shroud () that surrounds a rotation path () of the blade tip. As the blade tip reaches each shroud electrode, a controller () activates an electrical potential between them that generates a plasma-induced gas flow () directed toward the pressure side (PS) of the airfoil. The plasma creates a seal between the blade tip and the shroud, and induces a gas flow that opposes a leakage gas flow () from the pressure side to the suction side (SS) of the blade over the blade tip ().

Control of blade tip-to-shroud leakage in a turbine engine by directed plasma flow

A method for determining a preferred circumferential groove arrangement for a casing treatment of an axial flow compressor is disclosed. The method includes using the results from a three dimensional steady state computational fluid dynamic analysis to generate a flow field between a blade tip of a rotating blade and a compressor casing to determine the preferred circumferential groove arrangement. A stall margin for the axial flow compressor will be increased with the method.

Method for improving the stall margin of an axial flow compressor using a casing treatment

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

An arrangement () for conveying combustion gas from a plurality of can annular combustors to a turbine first stage blade section of a gas turbine engine, the arrangement () including a plurality of interconnected integrated exit piece (IEP) sections () defining an annular chamber () oriented concentric to a gas turbine engine longitudinal axis () upstream of the turbine first stage blade section. Each respective IEP () includes a first flow path section () receiving and fully bounding a first flow from a respective can annular combustor along a respective common axis () there between, and delivering a partially bounded first flow to a downstream adjacent IEP section (). Each respective IEP further includes a second flow path section () receiving a partially bounded second flow from an upstream adjacent IEP () and delivering at least part of the second flow to the turbine first stage blade section.

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Jupiter, FL, United States of America

Matthew D Montgomery