Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Anthony L Schiavo

Jay Alan Morrison

David Charles Radonovich

Bonnie D Marini

Gary Brian Merrill

Christian X Campbell

Timothy A Fox

Douglas A Keller

Bernd Prade

Steven James Vance

Reinhard Schilp

Malberto Gonzalez

Werner Krebs

Thomas Barrett Jackson

Thomas Walter Zagar

Stefan Mazzola

Danning You

Benjamin G Hettinger

James Bertoncello

John M Crane

Todd Willis

Kuangwei Huang

Malberto F Gonzalez

A gas turbine engine has transition ducts. Between two adjacent transition ducts there are transition side grooves that extend in a radial direction. A side seal is radially inserted into the transition side grooves. The side seal resiliently engages the transition side grooves to accommodate a thermo-mechanical stress that develops between the two adjacent transition ducts.

Transition system side seal for gas turbine engines

A gas turbine engine has a transition duct assembly. Between two adjacent transition ducts () there is a transition side groove () that is formed by transition side rails (). In this transition side groove () is inserted a side seal (-) that engages transition side grooves () that are formed in the transition side rails ().

A seal assembly between a transition seal structure associated with a downstream end of a transition duct and a vane seal structure associated with an upstream end of a vane structure in a first row vane assembly of a gas turbine engine includes a seal structure. The seal structure includes inner and outer seal members, each having a radially extending first leg and an axially extending second leg that provide each seal member with an L or V-shape. The seal members are arranged in a nested relationship with one of the seal members being positioned between the first and second legs of the other seal member. The first and/or second legs of the inner and outer seal members is/are received in a corresponding slot defined at least in part by one of the transition seal structure and the vane seal structure.

Seal assembly between a transition duct and the first row vane assembly for use in turbine engines

A ducting arrangement () for a combustion turbine engine is provided. The arrangement includes a ceramic liner () defining a hot gas path throughout a length of the ducting arrangement. A cooling sleeve () is disposed circumferentially outwardly onto the ceramic liner along the length. A metallic support frame () is disposed circumferentially outwardly onto the cooling sleeve along the length. The cooling sleeve may be structured with structural features along the length for biasing against the ceramic liner and the metallic support frame to resiliently accept mechanical and thermal growth induced loading that develops between the ceramic liner and the metallic support frame during operating conditions of the combustion turbine engine.

Ducting arrangement with a ceramic liner for delivering hot-temperature gases in a combustion turbine engine

The present disclosure provides a gas turbine combustor including a combustion structure () having a combustor liner () and a flow sleeve (). The combustor liner () includes inner and outer surfaces () and defines a combustion zone (). The gas turbine combustor further includes a plurality of hollow airfoil-shaped structures () affixed to the combustor liner () and extending radially outwardly into an airflow space () defined radially between the flow sleeve () and the combustor liner (). Each hollow structure () includes at least one metering tube () providing acoustic communication between the combustion zone () and the hollow structure (). The metering tubes () are detachably coupled to the combustor liner () for permitting interchanging of the metering tube () with at least one additional metering tube having at least one different dimension to effect a change in an acoustic characteristic of the hollow structure ().

Resonators with interchangeable metering tubes for gas turbine engines

A metal vane core or strut () is formed integrally with an outer backing plate (). An inner backing plate () is formed separately. A spring () with holes () is installed in a peripheral spring chamber () on the strut. Inner and outer CMC shroud covers () are formed, cured, then attached to facing surfaces of the inner and outer backing plates (). A CMC vane airfoil () is formed, cured, and slid over the strut (). The spring () urges continuous contact between the strut () and airfoil (), eliminating vibrations while allowing differential expansion. The inner end () of the strut is fastened to the inner backing plate (). A cooling channel () in the strut is connected by holes () along the leading edge of the strut to peripheral cooling paths () around the strut. Coolant flows through and around the strut, including through the spring holes.

CMC vane assembly apparatus and method

A gas turbine CMC shroud plate (A) with a vane-receiving opening () that matches a cross-section profile of a turbine vane airfoil (). The shroud plate (A) has first and second curved circumferential sides (A,A) that generally follow the curves of respective first and second curved sides () of the vane-receiving opening. Walls (A,A,A,A,) extend perpendicularly from the shroud plate forming a cross-bracing structure for the shroud plate. A vane () may be attached to the shroud plate by pins () or by hoop-tension rings () that clamp tabs () of the shroud plate against bosses () of the vane. A circular array () of shroud plates (A) may be assembled to form a vane shroud ring in which adjacent shroud plates are separated by compressible ceramic seals ().

Gas turbine vane platform element

A ceramic matrix composite (CMC) airfoil assembled from a pressure side wall () and a suction side wall () joined by interlocking joints () at the leading and trailing edges () of the airfoil to produce a tapered thin trailing edge. The trailing edge () is thinner than a combined thicknesses of the airfoil walls (). One or both of the interlocking joints () may be formed to allow only a single direction of assembly, as exemplified by a dovetail joint. Each joint () includes keys (F,F,F,F) on one side and respective keyways (K,K,K,K) on the other side. Each keyway may have a ramp () that eliminates indents in the airfoil outer surface that would otherwise result from the joint.

Interlocked CMC airfoil

A configuration of seals disposed around and between a plurality of ring segments () arrayed annularly about the periphery of moving blades in a gas turbine engine. The seals function to retain coolant in the plenum () within each of the ring segments (). The seals are disposed atop a substrate (A), which forms the top of the plenum (). The first seal () is made of a piece of sheet material and seals the gap between adjacent ring segments. This seal has an edge (A) thereof creased for mating with a similar seal on an adjacent ring segment. A second seal (), which is also made of sheet material, seals the ends of the plenum () of the ring segments (). Lastly, a third seal (), which is also made of a piece of sheet material, seals the sides of the second seal ().

Ring segment coolant seal configuration

A gas turbine ring segment () for use in gas turbine engines made from a ceramic matrix composite (CMC) material is disclosed. The ring segment includes a stacked multiplicity of CMC thin-sheet lamellae () each comprising a peripheral surface collectively defining a cross-section profile of the ring segment. The lamellae collectively define a channel () formed in the center thereof for receiving a bow-tie member (). The bow-tie member is disposed in the channel for holding together the stacked lamellae in a through thickness direction, and the in-plane strength of the bow-tie member is perpendicular to the in-plane strength of the lamellae. A stem portion () of the assembly may be further secured with a wrap () of CMC ribbon.

Stacked lamellae ceramic gas turbine ring segment component

A gas turbine transition duct is provided comprising a generally tubular main body having first and second ends, the first end being adapted to be positioned adjacent to a combustor unit and the second end being adapted to be positioned adjacent to a turbine and a collar coupled to the main body second end. The collar may have upper, lower and side portions. At least one of the side portions may be provided with a notched section.

Gas turbine transition duct apparatus

Protective insert plates () installed on a working gas face () of a turbine vane platform () provide replaceable portions of a turbine shroud for improved maintenance. The plates act as tiles, and may be formed of ceramic materials for thermal protection. Two cages () in the vane platform slidably receive the two insert plates () from opposite circumferential sides () of the platform. The plates slide into the cages up to the pressure and suction sides () of the vane airfoil (). The plates may have proximal edges () shaped to fit the respective pressure and suction sides of the vane airfoil. A retainer platemay be attached to a flange () on the cooled face () of each platform, and may contact each insert plate with a locking device () to prevent sliding of the plate in the cage.

Circumferential shroud inserts for a gas turbine vane platform

Aspects of the invention are directed to a ring seal segment having a base portion and a frame portion. The base portion and the frame portion are a unitary structure. The frame portion is made of four unitary side walls: a forward wall, an aft wall and a pair of transverse side walls. The side walls can be arranged in a rectangular configuration. At least a part of the base portion can be coated with a thermal insulating material. A ring seal segment according to aspects of the invention is well suited to withstand the expected operational loads in the turbine section. The transverse side walls can provide bending strength to the base and can strengthen the forward and aft side walls. Transition regions between the side walls and the base portion are strengthened by three perpendicular planes that cooperate to provide structural strength to the ring seal segment.

Ceramic matrix composite turbine engine components with unitary stiffening frame

A ceramic hybrid structure () that includes a wavy ceramic matrix composite (CMC) wall () bonded with a ceramic insulating layer () having a distal surface () that may define a hot gas passage () or otherwise be in proximity to a source of elevated temperature. In various embodiments, the waves () of the CMC wall () may conform to the following parameters: a thickness () between 1 and 10 millimeters; an amplitude () between one and 2.5 times the thickness; and a period () between one and 20 times the amplitude. The uninsulated backside surface () of the CMC wall () provides a desired stiffness and strength and enhanced cooling surface area. In various embodiments the amplitude (), excluding the thickness (), may be at least 2 mm.

Wavy CMC wall hybrid ceramic apparatus

A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a bolt or a plurality of bolts designed to prevent delamination of the ceramic plates when in use by keeping the ceramic plates in compression.

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Oviedo, FL, United States of America

Anthony L Schiavo