Siemens Energy, Inc.

Other

Richard Colin Charron

Matthew D Montgomery

Gary D Snyder

Daniel J Pierce

Jay Alan Morrison

David James Wiebe

David Allen Little

William W Pankey

Jody W Wilson

Raymond Scott Nordlund

Clinton A Mayer

Kenneth K Landis

John W Finneran

Jeffery W Samuelson

Benjamin G Hettinger

Ching-Pang Lee

Alexander Ralph Beeck

Jose L Rodriguez

Gerald A Myers

Hubertus Edward Paprotna

Walter Marussich

Bernhard W Küsters

Ernie B Campbell

Raymond S Nordlund

Canard anhedral controllable kite

Canard dihedral fishing kite

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

A transition duct assembly for a turbine engine includes a transition duct with an inner liner removably received therein. The duct is hollow with an inner peripheral surface and an outer peripheral surface. The duct can have an inlet end and an outlet end. The inner peripheral surface of the duct can be convergent along a majority of the length of the duct when moving from the inlet end to the outlet end thereof. The liner is hollow body with an inner peripheral surface and an outer peripheral surface. The outer peripheral surface of the liner can be correspondingly convergent to the inner peripheral surface of the duct. The inner peripheral surface of the liner can define an internal flow passage through the assembly. Such a construction permits the use of different materials for the liner and the duct and can allow the liner to be readily removed and replaced.

Combustor transition duct assembly with inner liner

A gas turbine engine ducting arrangement (), including: an annular chamber () configured to receive a plurality of discrete flows of combustion gases originating in respective can combustors and to deliver the discrete flows to a turbine inlet annulus, wherein the annular chamber includes an inner diameter () and an outer diameter (); an outer diameter mounting arrangement () configured to permit relative radial movement and to prevent relative axial and circumferential movement between the outer diameter and a turbine vane carrier (); and an inner diameter mounting arrangement () including a bracket () secured to the turbine vane carrier, wherein the bracket is configured to permit the inner diameter to move radially with the outer diameter and prevent axial deflection of the inner diameter with respect to the outer diameter.

Gas turbine structural mounting arrangement between combustion gas duct annular chamber and turbine vane carrier

A structural support system is provided in a can annular gas turbine engine having an arrangement including a plurality of integrated exit pieces (IEPs) forming an annular chamber for delivering gases from a plurality of combustors to a first row of turbine blades. A bracket structure is connected between an IEP and an inner support structure on the engine. The bracket structure includes an axial bracket member attached to an IEP and extending axially in a forward direction. A transverse bracket member has an end attached to the inner support structure and extends circumferentially to a connection with a forward end of the axial bracket member. The transverse bracket member provides a fixed radial position for the forward end of the axial bracket member and is flexible in the axial direction to permit axial movement of the axial bracket member.

Structural support bracket for gas flow path

A casing for a can annular gas turbine engine, including: a compressed air section () spanning between a last row of compressor blades () and a first row of turbine blades (), the compressed air section () having a plurality of openings () there through, wherein a single combustor/advanced duct assembly () extends through each opening (); and one top hat () associated with each opening () configured to enclose the associated combustor/advanced duct assembly () and seal the opening (). A volume enclosed by the compressed air section () is not greater than a volume of a frustum () defined at an upstream end () by an inner diameter of the casing at the last row of compressor blades () and at a downstream end () by an inner diameter of the casing at the first row of turbine blades ().

Casing for a gas turbine engine

A compressed air supply system for routing compressed air from a compressor to at least one combustor of a gas turbine engine is disclosed. The compressed air supply system may be formed from one or more plenums having an upstream end in fluid communication with an inner chamber of the compressor in which air is compressed and having a downstream end in fluid communication with the at least one combustor. Channeling compressed air through the plenum reduces damage to other components by confining the compressed air within the plenum. An upstream end of the plenum may be sealed to at least a portion of the compressor, and a downstream end of the plenum may be sealed to at least a portion of one or more combustors.

Compressed air plenum for a gas turbine engine

An exit piece () with an inlet throat () that conducts a combustion gas flow (A) in a path () from a combustor () to an annular chamber () that feeds the first blade section () of a gas turbine (). The exit piece further includes an outlet portion () that forms a circumferential segment of the annular chamber. The outlet portion interconnects with adjacent outlet portions by hinges (A,B,A,B). Each hinge may have a hinge axis (A,B) parallel to a centerline () of the turbine. Respective gas flows (A) are configured by an assembly () of the exit pieces to converge on the feed chamber () into a uniform helical flow that drives the first blade section with minimal circumferential variations in force.

Gas turbine combustor exit piece with hinged connections

An arrangement is provided for delivering gases from a plurality of combustors of a can-annular gas turbine combustion engine to a first row of turbine blades including a first row of turbine blades. The arrangement includes a gas path cylinder, a cone and an integrated exit piece (IEP) for each combustor. Each IEP comprises an inlet chamber for receiving a gas flow from a respective combustor, and includes a connection segment. The IEPs are connected together to define an annular chamber extending circumferentially and concentric to an engine longitudinal axis, for delivering the gas flow to the first row of blades. A radiused joint extends radially inward from a radially outer side of the inlet chamber to an outer boundary of the annular chamber, and a flared fillet extends radially inward from a radially inner side of the inlet chamber to an inner boundary of the annular chamber.

Assembly for directing combustion gas

A can annular industrial gas turbine engine, including: a single-piece rotor shaft spanning a compressor section (), a combustion section (), a turbine section (); and a combustion section casing () having a section () configured as a full hoop. When the combustion section casing is detached from the engine and moved to a maintenance position to allow access to an interior of the engine, a positioning jig () is used to support the compressor section casing () and turbine section casing ().

Full hoop casing for midframe of industrial gas turbine engine

A gas turbine combustor () disposed in a combustion air plenum () and a transition piece () for the combustor disposed in a separate cooling air plenum (). The combustion air plenum may receive combustion air () from a high-pressure compressor stage (A). The cooling air plenum may receive cooling air () from an intermediate-pressure compressor stage (B) at lower temperature and pressure than the combustion air. This cools the transition piece using less air than prior systems, thus making the gas turbine engine () more efficient and less expensive, because less expensive materials are needed and/or higher combustion temperatures are allowed. The cooling air may exit the cooling air plenum through holes () in a downstream portion () of the transition piece. An outer wall () on the transition piece may provide forced convection along the transition piece.

Gas turbine engine with high and intermediate temperature compressed air zones

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

A shaft cover support for a gas turbine engine is disclosed. The shaft cover support not only provides enhanced support to a shaft cover of the gas turbine engine, but also includes a cooling fluid chamber for passing fluids from a rotor air cooling supply conduit to an inner ring cooling manifold. Furthermore, the shaft cover support may include a cooling shield supply extending from the cooling fluid chamber between the radially outward inlet and the radially inward outlet on the radially extending region and in fluid communication with the cooling fluid chamber for providing cooling fluids to a transition section. The shaft cover support may also provide additional stiffness and reduce interference of the flow from the compressor. In addition, the shaft cover support accommodates a transition section extending between compressor and turbine sections of the gas turbine engine.

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West Palm Beach, FL, United States of America

Richard Colin Charron