Siemens Energy, Inc.

Siemens Aktiengesellschaft

Mikro Systems, Inc.

Reinhard Schilp

Timothy A Fox

Walter Ray Laster

Ching-Pang Lee

David Allen Little

Chander Prakash

David James Wiebe

Anthony L Schiavo

Christopher W Ross

David M Ritland

Muzaffer Sutcu

Domenico Gambacorta

Steven Williams

Erick J Deane

Joseph Scott Markovitz

Benjamin G Hettinger

James Bertoncello

Kyle L Landry

Jaishree Sharma

Danning You

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

The present disclosure provides a gas turbine combustor liner () comprising an outer surface () and an inner surface (), a plurality of film cooling holes () through a thickness of the gas turbine combustor liner (), and a plurality of resonator boxes () affixed to the outer surface () of the gas turbine combustor liner (). The film cooling holes () extend circumferentially around the gas turbine combustor liner () and comprise a first set of holes () having a first axial row spacing X and a second set of holes () having a second axial row spacing X'. The second set of holes () is formed in the gas turbine combustor liner () in a downstream direction relative to the first set of holes (). The second axial row spacing X&#x2032; is greater than the first axial row spacing X.

Film cooling hole arrangement for acoustic resonators in gas turbine engines

A flow conditioning device for a can annular gas turbine engine, including a plurality of flow elements disposed in a compressed air flow path leading to a combustor, configured such that relative adjustment of at least one flow directing element with respect to an adjacent flow directing element during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path.

Device for improved air and fuel distribution to a combustor

A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume.

Cooled conduit for conveying combustion gases

A pilot fuel nozzle for a combustor includes an igniter forming a central body extending along a longitudinal center of the nozzle. A nozzle tip includes a plurality of circumferentially spaced fuel passages and a plurality of circumferentially spaced air passages extending to an outer side of the nozzle tip. The central body extends through a center of the nozzle tip for producing a spark to ignite a fuel/air mixture adjacent to the nozzle tip. A plurality of fuel tubes extend along the central body, each of the fuel tubes having an outlet end engaged on the nozzle tip for delivery of fuel from the nozzle tip into a combustion chamber of the combustor An outer sleeve surrounds the fuel tubes and defines an annular space in fluid communication with the air passages of the nozzle tip between the outer sleeve and the central body.

Nozzle with multi-tube fuel passageway for gas turbine engines

A resonance chamber () has an outer wall () with coolant inlet holes (A-C), an inner wall () with acoustic holes (), and side walls (A-C) between the inner and outer walls. A depression (A-C) in the outer wall has a bottom portion () that is close to the inner wall compared to peaks (A-C) of the outer wall. The coolant inlet holes may be positioned along the bottom portion of the depression and along a bottom portion of the side walls to direct coolant flows () toward impingement locations () on the inner wall that are out of alignment with the acoustic holes. This improves impingement cooling efficiency. The peaks (A-C) of the outer wall provide volume in the resonance chamber for a target resonance.

Damping resonator with impingement cooling

A cover (A-B) enclosing with clearance () an acoustic damping resonator () on a working gas path liner () of a gas turbine component (). The cover includes a coolant inlet chamber (B) with a top wall (B) that is closer to the liner than a top wall () of the resonator, and extends upstream from the resonator relative to the working gas flow (). Compressed air () surrounds the cover at a higher pressure than the working gas and flows () into and through the coolant inlet chamber, then through holes () in the resonator, then exits through holes () the liner into the working gas. The coolant inlet chamber directs the flow of compressed air over a weld () of the upstream wall () of the resonator to cool it. The cover may be formed as a box () or a sleeve ().

Cooling chamber for upstream weld of damping resonator on turbine component

A flow conditioning device () for a can annular gas turbine engine, including a plurality of flow elements () disposed in a compressed air flow path () leading to a combustor (), configured such that relative adjustment of at least one flow directing element () with respect to an adjacent flow directing element () during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path ().

A gas turbine combustor assembly including a combustor liner and a plurality of fuel nozzle assemblies arranged in an annular array extending within the combustor liner. The fuel nozzle assemblies each include fuel nozzle body integral with a swirler assembly, and the swirler assemblies each include a bellmouth structure to turn air radially inwardly for passage into the swirler assemblies. A radially outer removed portion of each of the bellmouth structures defines a periphery diameter spaced from an inner surface of the combustor liner, and an interface ring is provided extending between the combustor liner and the removed portions of the bellmouth structures at the periphery diameter.

Interface ring for gas turbine fuel nozzle assemblies

A midframe portion () of a gas turbine engine () is presented and includes a compressor section with a last stage blade to orient an air flow () at a first angle (). The midframe portion () further includes a turbine section with a first stage blade to receive the air flow () oriented at a second angle (). The midframe portion () further includes a manifold () to directly couple the air flow () from the compressor section to a combustor head () upstream of the turbine section. The combustor head () introduces an offset angle in the air flow () from the first angle () to the second angle () to discharge the air flow () from the combustor head () at the second angle (). While introducing the offset angle, the combustor head () at least maintains or augments the first angle ().

Mid-section of a can-annular gas turbine engine with an improved rotation of air flow from the compressor to the turbine

A gas turbine engine compressed air flow control arrangement, including: a combustion gas structure having an acceleration geometry () configured to receive combustion gas () from a can combustor and accelerate the combustion gas () to a speed appropriate for delivery onto a first row of turbine blades, the combustion gas structure defining a straight combustion flow path; a film cooling hole () disposed through the combustion gas structure at a location within or downstream of the acceleration geometry (); a sleeve () surrounding at least a portion of the combustion gas structure comprising the film cooling hole and defining a volume () between the combustion gas structure and the sleeve (); and an adjustable flow control system configured to adjust a flow volume between the plenum () and the volume ().

Air regulation for film cooling and emission control of combustion gas structure

An emissions control system for a gas turbine engine including a flow-directing structure () that delivers combustion gases () from a burner () to a turbine. The emissions control system includes: a conduit () configured to establish fluid communication between compressed air () and the combustion gases within the flow-directing structure (). The compressed air () is disposed at a location upstream of a combustor head-end and exhibits an intermediate static pressure less than a static pressure of the combustion gases within the combustor (). During operation of the gas turbine engine a pressure difference between the intermediate static pressure and a static pressure of the combustion gases within the flow-directing structure () is effective to generate a fluid flow through the conduit ().

Device to lower NOx in a gas turbine engine combustion system

A fuel injector for use in a gas turbine engine combustor assembly. The fuel injector includes a main body and a fuel supply structure. The main body has an inlet end and an outlet end and defines a longitudinal axis extending between the outlet and inlet ends. The main body comprises a plurality of air/fuel passages extending therethrough, each air/fuel passage including an inlet that receives air from a source of air and an outlet. The fuel supply structure communicates with and supplies fuel to the air/fuel passages for providing an air/fuel mixture within each air/fuel passage. The air/fuel mixtures exit the main body through respective air/fuel passage outlets.

Air/fuel supply system for use in a gas turbine engine

A combustor liner has a plurality of resonators formed thereon. Each resonator has a radially outer wall and at least one side wall. The outer wall can be free of holes. Each resonator has an inner cavity defined between the outer wall, the at least one side wall and the outer peripheral surface of the liner. The at least one side wall of each resonator surrounds a subset of a plurality of holes that extend substantially radially through the liner. A plurality of cooling passages extends generally longitudinally within the liner. Each cooling passage has an inlet in fluid communication with the exterior of the liner and an outlet in fluid communication with the inner cavity of a respective one of the resonators. A coolant, such as compressor air, can enter and flow along the cooling passages to thereby cool the liner and purge the inner cavity of the resonator.

Resonator system with enhanced combustor liner cooling

A fuel injector assembly includes a fuel injector that provides fuel to be mixed with air for ignition in a combustor. The fuel injector assembly includes a fuel shroud and a plurality of spanning members. The spanning members are spaced circumferentially on the fuel shroud between the fuel shroud and a base plate assembly, which supports the fuel injector assembly within the combustor. The spanning members space the fuel shroud from the base plate assembly such that an annular gap is formed therebetween. The gap permits a flow of purge air therethrough to be mixed with compressed air entering the fuel shroud and with fuel provided by the fuel injector. The spanning members comprise elongate members having upstream ends that are circumferentially offset from opposed downstream ends so that the spanning members provide tangential velocity to the purge air flowing through the annular gap.

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Orlando, FL, United States of America

Reinhard Schilp