Siemens Energy, Inc.

Siemens Aktiengesellschaft

David M Ritland

Timothy A Fox

David James Wiebe

Andreas Böttcher

John Carl Glessner

Stephen A Ramier

Reinhard Schilp

Tobias Krieger

Juan Enrique Portillo Bilbao

Jens Kleinfeld

Thomas Grieb

Daniel W Garan

Scott M Martin

Rajesh Rajaram

Erick J Deane

Krishna C Miduturi

Joseph Scott Markovitz

Mark B Rubio

Jaishree Sharma

Marcus Zurhorst

An acoustically dampened gas turbine engine () having a gas turbine engine combustor () with an acoustic damping resonator system () is disclosed. The acoustic damping resonator system () may be formed from one or more resonators () formed from a resonator housing () positioned within the gas turbine engine combustor () at an outer housing () forming a combustor basket () and extending circumferentially within the combustor (). In at least one embodiment, the resonator housing () may include resonator chambers () that may be welded in place within resonator chamber () receivers () but easily replaceable without exposing the resonator housing () to damage. In another embodiment, an inner surface () of the resonator chamber () may be offset radially outward from an inner surface () of the resonator housing (), thereby creating a flow-path discontinuity and reducing heating of the resonator chamber (). The acoustic damping resonator system () may mitigate dynamics thereby increasing an engine operating envelope and decreasing emissions.

Acoustic damping system for a combustor of a gas turbine engine

A combustor () including: a first premix main burner () comprising a first swirler airfoil section (); a second premix main burner () comprising a second swirler airfoil section (); and a supply air reversing region upstream of the premix burners (), (). The first swirler airfoil section () and the second swirler airfoil section () are effective to impart swirl to a first airflow and a second airflow characterized by a same swirl number as the airflows exit respective burners (), (). The combustor () is effective to generate a first airflow volume through the first premix main burner () that is different than a second airflow volume through the second premix main burner ().

Air biasing system in a gas turbine combustor

A combustor arrangement () including: a pilot burner () having a pilot cone (); a plurality of clockwise () main swirlers interposed among a plurality of counterclockwise () main swirlers and disposed concentrically about the pilot burner; and a base plate () transverse to the main swirlers. Inbound-zones () exist where adjacent portions () of adjacent flows () through main swirlers flow toward the pilot cone, and interposed between the inbound zones outbound zones () exist where adjacent portions of adjacent flows flow away from the pilot cone. The arrangement is configured to preferentially deliver more cooling fluid to the inbound zones than the outbound zones.

Asymmetric baseplate cooling with alternating swirl main burners

An annular fluid distribution device () for distributing fluid into a gaseous flow (), that includes: a first fluid distribution manifold () having a first fluid inlet and first fluid outlets (), wherein the first fluid outlets inject a first fluid into the gaseous flow; and a second fluid distribution manifold (), having a second fluid inlet and second fluid outlets (), wherein the second fluid outlets inject a second fluid into the gaseous flow. The second fluid manifold is isolated from the first fluid distribution manifold, and the first fluid outlets and the second fluid outlets are disposed on a common fluid outlet plane ().

Circumferential biasing and profiling of fuel injection in distribution ring

An assembly for a gas turbine is presented. The assembly includes a gas supply pipe passing through a bore in a flange of the gas turbine for supplying gas to a combustion chamber of the gas turbine and a sleeve surrounding the gas supply pipe, having a first end and a second end, wherein the first end is sealingly coupled to the gas supply pipe, and wherein the sleeve is adapted to be sealingly coupled to the flange at the second end such that the sleeve extends along a thickness of the flange.

Gas turbine assembly and method therefor

A combustion apparatus in a gas turbine engine comprises a combustor shell for receiving air, a fuel injection system associated with the combustor shell, a first fuel supply structure, and a shield structure. The fuel supply structure is in fluid communication with a source of fuel for delivering fuel from the source of fuel to the fuel injection system and comprises a first fuel supply elements including a first section extending along a first path having a component in an axial direction and a second section extending from the first section along a second path having a component in a circumferential direction. The shield structure is associated with at least a portion of the second section of the first fuel supply element.

Combustor apparatus in a gas turbine engine

A flow sleeve for use in a turbine engine including a compressor section, a combustion section, and a turbine section downstream from the combustion section. The flow sleeve includes a sleeve wall defining a pre-mixing passage of the combustion section and has a forward end proximate to a cover plate of the combustion section and an opposed aft end. The sleeve wall includes a first wall section extending axially from the sleeve wall forward end and a second wall section extending from the first wall section between the forward and aft ends of the sleeve wall toward the aft end of the sleeve wall. A cavity is formed between the first wall section and the second wall section.

Turbine engine flow sleeve

A can-annular gas turbine engine including a casing defining an annulus for receiving air flow from a compressor. A combustor basket is supported within the casing, and a swirler assembly is located within the basket and includes an air inlet and an outlet located in a combustion chamber. The swirler assembly is configured to discharge a fuel/air mixture from the outlet downstream into the combustion chamber. An acoustic filter is provided comprising a Herschel-Quincke tube having a first open end at the annulus and a second open end adjacent to the swirler assembly within the combustion chamber. Components of acoustic pressure oscillations travel along two paths of different lengths from one end of the Herschel-Quincke tube to the opposite end of the tube where out of phase components of the pressure oscillations interact to attenuate the pressure oscillations.

Attenuation of combustion dynamics using a Herschel-Quincke filter

A combustor for use in a turbine engine including a compressor section, a combustion section downstream from the compressor section, and a turbine section downstream from the combustion section. An inner annulus wall extends from a burner end of the combustor to an outlet end of the combustor. An outer annulus wall is disposed outwardly from the inner annulus wall and extends from the burner end of the combustor to the outlet end of the combustor. A passageway is formed between the inner annulus wall and the outer annulus wall and extends from a combustion zone to the outlet end of the combustor. A plurality of symmetrically distributed section walls extends between the inner annulus wall and the outer annulus wall from the burner end of the combustor toward the outlet end of the combustor. The section walls divide the combustion zone into a plurality of segments.

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Winter Park, FL, United States of America

David M Ritland