Siemens Energy, Inc.

Siemens Aktiengesellschaft

Siemens Westinghouse Power Corporation

Siemens Power Generation, Inc.

Samer P Wasif

Clifford E Johnson

Robert James Bland

Walter Ray Laster

Bernd Prade

John Carl Glessner

Stephen A Ramier

Muzaffer Sutcu

Andreas Koch

Domenico Gambacorta

Stefan Dahlke

Udo Schmitz

Karsten Jordan

Joachim Lepers

J Scott Markovitz

David Kargetta

Stefan Dahlk

Bertram Janus

Juergen Meisl

Domenlco Gambacorta

Stephen A Raimer

Scott J Markovitz

A can annular combustion arrangement () for a gas turbine engine (), including: a combustor can () having an inlet () at a head end (); an annular chamber () surrounding the combustor can () for delivering a flow of compressed air from a plenum () to the inlet (); a fuel injector () configured to inject a flow of fuel into the flow of compressed air; and a flow tripping device () surrounding the combustor can () and disposed in the annular chamber () downstream of the fuel injector ().

Can annular combustion arrangement with flow tripping device

A fuel gas cooling system for a combustion basket spring clip seal support is disclosed. The fuel gas cooling system may be formed from one or more fuel gas supply channels terminating proximate to a spring clip at the intersection between a combustor basket and a transition section such that fuel gas may be supplied to the hot gas path proximate to the intersection between the combustor basket and the transition section. The fuel gas supply channel may create an intermediate fuel gas burn at this intersection, which may reduce the firing temperature at the fuel nozzles and reduce NOx emissions.

Fuel gas cooling system for combustion basket spring clip seal support

A method of influencing combustion dynamics, including measuring a combustion dynamics parameter, and controlling a diluent flow () delivered to a fuel flow () upstream of a pilot burner fuel outlet () in response to the measured combustion dynamics parameter.

Utilizing a diluent to lower combustion instabilities in a gas turbine engine

Embodiments of the present invention provide resonators () that have lateral walls () disposed at non-square angles relative to the liner's longitudinal (and flow-based) axis () such that a film cooling of substantial portions of an intervening strip () is provided from apertures (A,B,) in a resonator box () adjacent and upstream from the intervening strip (). This film cooling also cools weld seams () along the lateral walls () of the resonator boxes (). In various embodiments the lateral wall angles are such that film cooling may be provided to include the most of the downstream portions of the intervening strips (). These downstream portions are closer to the combustion heat source and therefore expected to be in greater need of cooling.

Non-rectangular resonator devices providing enhanced liner cooling for combustion chamber

A combustor assembly is provided comprising a combustor casing; a liner coupled to the combustor casing; a burner assembly coupled to the combustor casing; and at least one resonator assembly. The resonator assembly comprises a resonator outer plate having at least one opening, a resonator side wall coupled to the resonator outer plate, and a resonator inner plate defined by a portion of the liner. The resonator inner plate is coupled to the resonator side wall and has at least one slot formed therein having an aspect ratio of at least 4:1.

Combustor assembly including one or more resonator assemblies and process for forming same

A gas sleeve () for a combustor () of gas turbine engine () attaches to a support housing () of the combustor () to convey a fuel gas and to fit within a fuel rocket (). The gas sleeve () comprises a plurality of apertures () formed to provide impingement cooling. The apertures () comprise a tilt angle directed toward a structure in need of impingement cooling, for instance a weld joint () that attaches the fuel rocket () to the support housing (). The apertures () additionally may comprise a rotational angle effective to create a rotationally swirling flow of the portion of fuel gas that passes through the apertures (). A method of operation using this structure also is provided.

Extended life fuel nozzle

One or more Helmholtz-type resonators () is/are provided at the junction () of a combustor () and a combustion chamber () of a gas turbine engine (). In one embodiment, adjacent Helmholtz-type resonators (), which may be separated by respective baffles (), have different volumes that help provide for damping different undesired combustion-generated acoustic pressure waves. In some embodiments, a structural member () may be provided between adjacent Helmholtz-type resonators () at the junction. At least one of the plurality of Helmholtz-type resonators comprises one or more inlet openings (), and one or more exit openings (). Embodiments (-) are described in which Helmholtz-type resonators provided at the junction are enlarged in size using various approaches. The positioning at the junction, upstream of the space () in which combustion occurs, and providing a plurality of differently sized resonators, provides for improved flexibility and resonator damping efficiencies.

Resonator device at junction of combustor and combustion chamber

Aspects of the invention are directed to an acoustic resonator with improved impingement cooling effectiveness. The resonator includes a plate with an inside face and an outside face. A plurality of passages extend through the plate. The resonator includes a side wall that extends from and about the plate. A plurality of cooling tubes are attached to the resonator plate such that an inner passage of each cooling tube is in fluid communication with a respective passage in the resonator plate. The resonator can be secured to a surface of a turbine engine combustor component to define a closed cavity. The ends of the cooling tubes are spaced from the surface. Thus, a coolant can enter the passages in the plate and can be directed to the surface so as to impingement cool the surface. The cooling tubes can minimize coolant loss by dispersion in the cavity.

Acoustic resonator with impingement cooling tubes

A gas turbine engine () includes a cylindrical basket () having an axis () and a single main burner assembly () disposed within the basket. A burner insert () is disposed in an annular space between the burner assembly and the basket. The insert includes a face perpendicular to the axis of the basket. A plurality of passageways () are formed in the basket, positioned proximate to and downstream of the burner insert for allowing passage of a portion of an oxidizer flow () into a combustion chamber (). A fluid flow path (), defined between a combustion chamber liner portion () of the basket and a casing () spaced radially outward from the combustion liner portion, discharges a fluid into a flow reversal region () proximate an inlet () of the burner assembly. A fuel outlet () is disposed in the flow reversal region.

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Oviedo, FL, United States of America

Samer P Wasif