Siemens Aktiengesellschaft

Siemens Energy, Inc.

Siemens Energy Global Gmbh & Co. Kg

Siemens Aktiengesell Schaft

Kok-Mun Tham

Ching-Pang Lee

Vincent Paul Laurello

Gm Salam Azad

Eric Joseph Schroeder

Nicholas F Martin, Jr

Manjit Shivanand

Abdullatif M Chehab

Zhihong Gao

David G Sansom

Neil Hitchman

Li Shing Wong

Ramesh Subramanian

Paul Hadley Vitt

John Joseph Marra

Chander Prakash

Steven Koester

Samuel R Miller, Jr

Sin Chien Siw

Brian Harry Terpos

Yan Yin

Erik Johnson

Jamie Meeroff

Barrry L Allmon

Dustan M Simko

Eric Chen

Gary Brian Merrill

Christian X Campbell

Peter Schröder

Matthew D Montgomery

Todd A Ebert

Kevin Michael Light

Christopher W Ross

Cora Schillig

Kevin Kampka

Ross H Peterson

Weidong Cai

Christopher K Rawlings

Farzad Taremi

Dimitrios Zois

Jonathan E Shipper

Robert W Sunshine

Andrew Lohaus

Shantanu P Mhetras

Gary D Lock

Santiago R Salazar

Christian Keune

Patrick M Pilapil

Roger Matthews

Dustin Jaime Muller

David A Kemp

Ernie B Campbell

Nicolas Benjamin Wappler

Andrew S Lohaus

Steven Coppess

Kahwai G Muriithi

John M Owen

Joana Verheyen

Carl M Sangan

Javan Albright

John A Fussner

James McCoy

Bijay K Sultanian

Stephen R Williamson

Melissa Harris

A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to an aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from the blade root for cooling the turbine blade.

Cover plate with flow inducer and method for cooling turbine blades

A blade () for a turbine engine that includes an internal cooling system () formed from at least one cavity () positioned within a generally elongated airfoil (). A squealer tip () and at least one densified oxide dispersion strengthened layer () extend radially from a radially outer tip cap () of the blade (), the tip cap () having a tip cap upper surface ().

Turbine blade with squealer tip and densified oxide dispersion strengthened layer

The invention relates to a method for starting up a gas turbine engine of a combined cycle power plant. The method includes applying load to the gas turbine engine and increasing the load until a predetermined combustor firing temperature is reached, while keeping the adjustable inlet guide vanes in a start position adapted to reduce the mass flow of air into the compressor; further increasing the load of the gas turbine engine while opening the adjustable inlet guide vanes and keeping the predetermined combustor firing temperature constant until the inlet guide vanes reach an end position adapted to increase the mass flow of air into the compressor; further increasing the load of the gas turbine engine while keeping the adjustable inlet guide vanes in the end position until a predetermined load of the gas turbine engine is reached.

Method for starting up a gas turbine engine of a combined cycle power plant

A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system () for the downstream turbine assembly.

Compressor bleed cooling system for mid-frame torque discs downstream from a compressor assembly in a gas turbine engine

A turbine component including a shrouded airfoil with a flow conditioner configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner is positioned on a shroud base radially adjacent to the tip of the airfoil and includes a ramped radially outer surface positioned further radially inward than a radially outer surface of the shroud base. The ramped radially outer surface extends from a first edge to a second edge in a direction generally from the suction side to the pressure side of the airfoil, such that the first edge is positioned further radially inward than the second edge. Multiple coolant ejection holes are positioned on the ramped radially outer surface. The coolant ejection holes are connected fluidically to an interior of the airfoil.

Shrouded turbine blade

A gas turbine engine having a rotor centering cooling system for cooling struts within an exhaust diffuser and turbine case to reduce tip rub during hot restarts is disclosed. In particular, the rotor centering cooling system may be positioned within struts in the exhaust diffuser downstream from a turbine assembly for limiting thermal gradients between top and bottom struts to prevent the exhaust bearing body from becoming off-center during steady state operation as a result of the top struts becoming hotter than the bottom struts. The rotor centering cooling system may reduce the temperature at the exhaust diffuser and turbine case, thereby reducing the thermal gradient between the top and bottom struts and top and bottom of the turbine case. As such, the exhaust bearing body remains centered, thereby preventing a tighter blade tip clearance at the top of the turbine assembly than at the bottom of the assembly.

Gas turbine engine with rotor centering cooling system in an exhaust diffuser

A compressor () configured for use in a gas turbine engine () and having a rotor assembly () with a pumping system () positioned on a rotor drum () to counteract reverse leakage flow at a gap () formed between one or more stator vane tips () and a radially outer surface () of the rotor drum (). The pumping system () may be from pumping components () positioned radially inward of one or more stator vane tips () to reduce, if not completely eliminate, reverse leakage flow at the stator vane tips (). In at least one embodiment, the pumping component () may be formed from one or more cutouts () in the outer surface () of the rotor drum (). In another embodiment, the pumping component () may be formed from at least one pumping fin () extending from the radially outer surface () of the rotor drum (). In at least one embodiment, rows () of pumping components () may be aligned with rows () of stator vanes () within the compressor ().

Compressor usable within a gas turbine engine

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with ridges projecting from the abradable surface, separated by grooves. The ridges have one or both sidewalls inclined against the opposing turbine blade tip rotational direction for redirecting and/or dissipating blade tip gap leakage airflow energy. In some embodiments the ridge tip and/or groove base have inclined profiles for redirecting airflow leakage away from the blade tip gap. In some embodiments, the inclined ridge tip profile provides a progressive wear zone that increases abradable surface area as the inclined ridge is abraded by the rotating blade tip.

Turbine abradable layer with inclined angle surface ridge or groove pattern

Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.

Turbine shroud with abradable layer having dimpled forward zone

Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.

Turbine shroud with abradable layer having ridges with holes

A shrouded turbine airfoil () with a leakage flow conditioner () configured to direct leakage flow to be aligned with main hot gas flow is disclosed. The leakage flow conditioner () may be positioned on a radially outer surface () of an outer shroud base () of the outer shroud () on a tip () of an airfoil (). The leakage flow conditioner () may include a radially outer surface () that is positioned further radially inward than the radially outer surface () of the outer shroud base () creating a radially outward extending wall surface () that serves to redirect leakage flow. In at least one embodiment, the radially outward extending wall surface () may be aligned with a pressure side () of the shrouded turbine airfoil () to increase the efficiency of a turbine engine by redirecting leakage flow to be aligned with main hot gas flow to reduce aerodynamic loss upon re-introduction to the main gas flow.

Shrouded turbine airfoil with leakage flow conditioner

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Turbine abradable layer with compound angle, asymmetric surface area ridge and groove pattern

Gas turbine engine blade squealer tips incorporate cooling slots formed in the suction side rail downstream of the leading edge for directing cooling gas flow along an inside edge of the squealer tip pressure side rail. Some embodiments incorporate a tip fin on the suction side rail proximal a cooling slot. Segmented suction side rail embodiments abrade opposing turbine casing abradable surfaces prior to potential contact with the pressure side rail, reducing likelihood of pressure side rail friction heating. During turbine engine operation cooler pressure side rails reduce likelihood of squealer tip erosion.

Segmented turbine blade squealer tip and cooling method

An abradable turbine component, a method of creating a turbine component with an abradable mesh structure, and a gas turbine engine are provided. The abradable turbine component includes a turbine component surface for coupling to a turbine casing, and a deposited abradable mesh structure coupled to the turbine component surface. The abradable mesh structure includes interlacing strands of material, each strand including a height relative to the turbine component surface. At least two of the plurality of interlacing strands include a height different from each other. The method includes applying a bond coat layer followed by a thermal barrier coating layer. An abradable mesh structure is deposited on top of thermal barrier coating wherein the abradable mesh structure includes interlacing strands of material wherein at least two of the interlacing strands include a height different from each other. A gas turbine engine including the abradable turbine component is also provided.

Additive manufacturing of abradable mesh structure on ring segment surface

A vane carrier temperature control system for use in a gas turbine engine includes a first cooling air source, a second cooling air source, and an air temperature control system. The first cooling air source supplies a first portion of vane carrier cooling air extracted from a compressor section of the engine to a first section of a vane carrier that supports a plurality of rows of vanes within a turbine section of the engine. The second cooling air source supplies a second portion of vane carrier cooling air extracted from the compressor section to a second section of the vane carrier spaced from the first section in an axial direction defined by a direction of hot working gas flow through the turbine section. The air temperature control system controls a temperature of at least one of the first and second portions of vane carrier cooling air.

Vane carrier temperature control system in a gas turbine engine

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Oviedo, FL, United States of America

Kok-Mun Tham