Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Siemens Power Generation, Inc.

Other

Mikro Systems, Inc.

Jay Alan Morrison

Gary Brian Merrill

Jay Edgar Lane

Ching-Pang Lee

Steven James Vance

Harry A Albrecht

Christian X Campbell

Yevgeniy P Shteyman

Daniel George Thompson

Douglas A Keller

Malberto Gonzalez

Anthony L Schiavo

Steven C Butner

Allister William James

Michael Anthony Burke

Richard Colin Charron

David Charles Radonovich

John Joseph Marra

Matthew D Montgomery

Thomas Barrett Jackson

David James Wiebe

Alexander Ralph Beeck

Alfred Paul Matheny

Jose L Rodriguez

Marc Tertilt

John V Stewart

Raymond Scott Nordlund

Daniel M Eshak

Bonnie D Marini

Chris Campbell

Walter Marussich

Stefan Hoffmann

Raymond G Snider

Humberto A Zuniga

Robert William Kreutzer

Daniel J Pierce

Andrew Szweda

Brian J Wessell

Kenneth K Landis

Jody W Wilson

Christine Taut

Scott M Widrig

Brede J Kolsrud

Bernhard W Küsters

Johan Hsu

Daniel Hofmann

Steve James Vance

Ernie B Campbell

Eric V Carelli

Michael E Crawford

Dougas A Keller

Vijay Parthasarathy

Yevgeuiy Shteyman

Evan McNeil Ludeman

Michael A Burke

Daniel G Thompson

Yevgenly Shteyman

Mark Tertilt

A multiple laser sight system for an archery bow or the like configured so that the multiple laser systems can be calibrated together and having features such that the user can use one laser system during the day and one laser system during the light. The laser sight is further configured to not interfere with the optional use of conventional sighting pins and the use of evening infrared systems, like the prior art use of night vision goggles.

Bow sight apparatus having multiple lasers

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 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 layer () may be bi-cast onto ligaments () extending from a porous cooling construction (). The material of the structural layer may be optimized for high-temperature strength, while the material of the porous construction may be optimized for high thermal conductivity. A fugitive material () such as wax may be formed on the ligaments of the porous construction. A second fugitive material () such as ceramic may fill the remaining part of the porous construction. An investment casting shell () may be disposed around the porous construction and the fugitive materials. The first fugitive material may then be replaced with the material of the structural layer (), and the second fugitive material may be removed to provide coolant paths (). A second structural layer () may be bi-cast onto further ligaments () on a second side of the porous construction.

Bi-cast layered wall with a porous element for component cooling

A cooling arrangement in a gas turbine system (). The arrangement includes a plurality of flow network units () to transfer heat to cooling fluid, at least one unit including first (), second (), and third () flow sections between openings () in a first wall () and an opening in a second wall () to pass cooling fluid through the walls. The first section includes first flow paths, between the openings in the first wall and the second section, extending to the second section. The third section includes third flow paths, between the second section and the opening in the second wall, to effect flow of cooling fluid. The second section includes one or more cooling fluid flow paths between the first section and the third section. The number of flow paths in the second section is fewer than the number of first flow paths and fewer than the number of third flow paths.

Cooling module design and method for cooling components of a gas turbine system

A gas turbine component () with a layer of ceramic material () defining a wear surface (), in which an array of cross-shaped depressions (A,B) formed in the wear surface () define a continuous labyrinth of orthogonal walls () of the ceramic material, and reduce the area of the wear surface () by about 50%. Within a representative area () of the wear surface (), the depressions (A,B) provide a ratio of a lineal sum of depression perimeters () divided by the representative area () of the wear surface of least 0.9 per unit of measurement for improved abradability characteristics of the wear surface ().

Patterned reduction of surface area for abradability

A cooling system for a transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have a multi-panel outer wall formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that at least one cooling chamber is formed between the inner and intermediate panels. The transition duct may also include an outer panel. The inner, intermediate and outer panels may include one or more metering holes for passing cooling fluids between cooling chambers for cooling the panels. The intermediate and outer panels may be secured with an attachment system coupling the panels to the inner panel such that the intermediate and outer panels may move in-plane.

Turbine transition component formed from a two section, air-cooled multi-layer outer panel for use in a gas turbine engine

An apparatus includes a metal shell () surrounding a body () that is made of a ceramic matrix composite (CMC) material () The metal shell defines a space () adapted to contain the body and includes at least one protrusion () adapted to contact the body. A compliant porous element () is adapted to fit in the space between the metal shell and the body. A preload spring () is provided in an urging orientation with the body wherein the preload spring is positioned against a first region () of the body and is adapted to urge the body toward one of the protrusions positioned against a second region () generally opposite the first region, and also to preload the compliant porous element. One of the protrusions may be a hard stop, and in the preload, one of the protrusions may be loaded.

Apparatus comprising a CMC-comprising body and compliant porous element preloaded within an outer metal shell

Turbine transition component formed from an air-cooled multi-layer outer panel for use in a gas turbine engine

A ducting arrangement () for a can annular gas turbine engine, including: a duct () disposed between a combustor () and a first row of turbine blades and defining a hot gas path () therein, the duct () having raised geometric features () incorporated into an outer surface (); and a flow sleeve () defining a cooling flow path () between an inner surface () of the flow sleeve () and the duct outer surface (). After a cooling fluid () traverses a relatively upstream raised geometric feature (), the inner surface () of the flow sleeve () is effective to direct the cooling fluid () toward a landing () separating the relatively upstream raised geometric feature () from a relatively downstream raised geometric feature ().

Ducting arrangement for cooling a gas turbine structure

A method for joining a first CMC part () having an outer joining portion (), and a second CMC part () having an inner joining portion (). The second CMC part () is heat-cured to a stage of shrinkage more complete than that of the first CMC part () prior to joining. The two CMC parts () are joined in a mating interface that captures the inner joining portion () within the outer joining portion (). The assembled parts () are then fired together, resulting in differential shrinkage that compresses the outer joining portion () onto the inner joining portion (), providing a tightly pre-stressed joint. Optionally, a refractory adhesive () may be used in the joint. Shrinkage of the outer joining portion () avoids shrinkage cracks in the adhesive ().

Method of joining ceramic parts and articles so formed

A cooling arrangement () having: a duct () configured to receive hot gases () from a combustor; and a flow sleeve () surrounding the duct and defining a cooling plenum () there between, wherein the flow sleeve is configured to form impingement cooling jets () emanating from dimples () in the flow sleeve effective to predominately cool the duct in an impingement cooling zone (), and wherein the flow sleeve defines a convection cooling zone () effective to cool the duct solely via a cross-flow (), the cross-flow comprising cooling fluid () exhausting from the impingement cooling zone. In the impingement cooling zone an undimpled portion () of the flow sleeve tapers away from the duct as the undimpled portion nears the convection cooling zone. The flow sleeve is configured to effect a greater velocity of the cross-flow in the convection cooling zone than in the impingement cooling zone.

Regeneratively cooled transition duct with transversely buffered impingement nozzles

A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having an inner surface, an outer surface, a region, an inlet, and an outlet. The inner surface defines an inner volume of the conduit. The region extends between the inner and outer surfaces and includes cooling fluid structure defining a plurality of cooling passageways. The inlet extends inwardly from the outer surface and provides fluid communication between the inlet and the passageways. The outlet extends from the passageways to the inner surface to provide fluid communication between the passageways and the inner volume. At least one first cooling passageway intersects with at least one second cooling passageway such that cooling fluid flowing through the first cooling passageway interacts with cooling fluid flowing through the second cooling passageway.

Mesh cooled conduit for conveying combustion gases

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Oviedo, FL, United States of America

Jay Alan Morrison