Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Siemens Aktiengesellschaft

Siemens Power Generation, Inc.

David Allen Little

Richard Colin Charron

Abdullatif M Chehab

Reinhard Schilp

Hubertus Edward Paprotna

Ching-Pang Lee

David James Wiebe

Thomas E Lippert

Walter Ray Laster

Jiping Zhang

Chander Prakash

Jose L Rodriguez

Kevin Michael Light

Christopher W Ross

Raymond Scott Nordlund

Gerald A Myers

Clifford E Johnson

Jody W Wilson

Adam Matthew Foust

Gerard McQuiggan

Gary D Snyder

Zhengxiang Pu

David L Wasdell

Eduardo E Bancalari

Joseph Fadok

Nitin Chhabra

Gerry McQuiggan

Elizabeth R Colore

Raymond S Nordlund

Eduardo Bancalari

An arrangement () for delivering gases from combustors () to a first row of blades. The arrangement () includes at least an upstream flow path () including an aft first side wall () and a downstream flow path () including a forward second side wall (). A convergence junction trailing edge () is defined at a downstream terminal edge () of the first side wall (), and the second side wall () converges toward the first side wall () in the direction of the convergence junction trailing edge (). An impingement sheet structure () is located between and provides impingement cooling air to the first and second side walls (). Openings () provide a cooling air passage between the first and second side walls () and provide a flow of post impingement air into the gas path at the convergence junction trailing edge ().

Arrangement for a gas turbine combustion engine

A cooling arrangement (), including: a substrate having a pocket () defined by a rib side surface () and a bottom surface (); a feature () formed in the rib side surface; and an impingement plate (). In an installed configuration () a resilience of the impingement plate enables an interference between the impingement plate and the feature that locks the impingement plate in an installed position (). An elastic compression of the impingement plate from the installed configuration eliminates the interference.

Impingement cooling arrangement having a snap-in plate

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 midframe portion () of a gas turbine engine () is presented, and includes a compressor section () configured to discharge an air flow () directed in a radial direction from an outlet of the compressor section (). Additionally, the midframe portion () includes a manifold () to directly couple the air flow () from the compressor section () outlet to an inlet of a respective combustor head () of the midframe portion ().

Mid-section of a can-annular gas turbine engine with a radial air flow discharged from the compressor section

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

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 midframe portion () of a gas turbine engine () is presented and includes a compressor section with a last stage blade to orient an air flow () at a first angle (). The midframe portion () further includes a turbine section with a first stage blade to receive the air flow () oriented at a second angle (). The midframe portion () further includes a manifold () to directly couple the air flow () from the compressor section to a combustor head () upstream of the turbine section. The combustor head () introduces an offset angle in the air flow () from the first angle () to the second angle () to discharge the air flow () from the combustor head () at the second angle (). While introducing the offset angle, the combustor head () at least maintains or augments the first angle ().

Mid-section of a can-annular gas turbine engine with an improved rotation of air flow from the compressor to the turbine

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

A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.

Compressor airfoil tip clearance optimization system

A turbine engine including an intermediate space defined between outer and inner portions of the turbine engine. A flow energizer is provided including a flow body located within the intermediate space and including an inlet port, an outlet port and a flow passage extending within the flow body between the inlet and outlet ports. The inlet port receives a flow of a first medium located within the intermediate space and the flow body injects an energizing flow of a second medium to a portion of the first medium within the flow body to create an energized flow of a mixed medium from the outlet portion, the energized flow of mixed medium creates a flow of the first medium adjacent to the flow body within the intermediate space.

Temperature control within a cavity of a turbine engine

A strut cover for use in a gas turbine engine having structure defining an annular flow path for receiving exhaust gas from a turbine section of the engine. The strut cover is located downstream from a last row of blades of the turbine section and extends radially through the flow path between inner and outer walls. The strut cover includes an upstream section and a downstream section. The upstream section defines a leading edge for the strut cover and is supported on a pivot axis for pivotal movement about the pivot axis. The downstream section defines a trailing edge for the strut cover and includes an upstream end positioned adjacent to a downstream end of the upstream section. The downstream section is stationary relative to the inner and outer walls to define a predetermined flow angle for directing exhaust gases flowing from the upstream section and passing through the diffuser.

Movable strut cover for exhaust diffuser

The present invention comprises a gas turbine engine and a process for operating a gas turbine engine. A fluid structure receives compressed air from a compressor and extends toward a stationary blade ring in a turbine to discharge the compressed air directly against a surface of the blade ring such that the compressed air impinges on the blade ring surface. The compressed air then passes through at least one opening in the stationary blade ring and into cooling passages of a corresponding row of vanes.

Gas turbine engine with structure for directing compressed air on a blade ring

A fuel shroud assembly () into which fuel () is injected for mixing with an air stream () in a fuel manifold. The shroud assembly () comprises a plurality of parallel fuel scoops () each receiving the injected fuel (). The fuel stream () flows through each scoop (), exiting at an open scoop end (A). The air stream () flows between scoops, creating a shear region proximate each scoop end (A) where the fuel exits. The shear causes mixing of the air () and the fuel () streams, wherein the degree of mixing is not dependent on the momentum ratio of the air () or fuel () streams.

Apparatus for mixing fuel and air in a combustion system

A turbine engine assembly for a generator including a turbine engine having a compressor section, a combustor section and a turbine section. An air bleed line is in communication with the combustor section for receiving combustor shell air from the combustor section and conveying the combustor shell air as bleed air to a plurality of stages of the turbine section. Bleed air is controlled to flow through the air bleed line when an operating load of the turbine engine assembly is less than a base load of the engine to bypass air exiting the compressor section around a combustor in the combustor section and effect a flow of high pressure combustor shell air to the stages of the turbine section.

Cooling of turbine components using combustor shell air

A seal assembly that limits gas leakage from a hot gas path to one or more disc cavities in a turbine engine. The seal assembly includes a seal apparatus that limits gas leakage from the hot gas path to a respective one of the disc cavities. The seal apparatus comprises a plurality of blade members rotatable with a blade structure. The blade members are associated with the blade structure and extend toward adjacent stationary components. Each blade member includes a leading edge and a trailing edge, the leading edge of each blade member being located circumferentially in front of the blade member's corresponding trailing edge in a direction of rotation of the turbine rotor. The blade members are arranged such that a space having a component in a circumferential direction is defined between adjacent circumferentially spaced blade members.

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Chuluota, FL, United States of America

David Allen Little