System and procedure for the determination of an exhaust goal temperature for a gas turbine.

Die Erfindung betrifft ein System und ein Verfahren zum Bestimmen einer Sollabgastemperatur für eine Gasturbine (10). Das Verfahren beinhaltet die Bestimmung einer Sollabgastemperatur wenigstens zum Teil auf der Basis eines Verdichterdruckzustandes; der Bestimmung einer Temperaturanpassung an die Sollabgastemperatur wenigstens zum Teil auf der Basis einer Dampffeuchtigkeit; und einer Änderung der Sollabgastemperatur wenigstens zum Teil auf der Basis der Temperaturanpassung.

System and method for determining an exhaust gas-target temperature for a gas turbine.

Die Erfindung betrifft ein System und ein Verfahren zum Bestimmen einer Sollabgastemperatur für eine Gasturbine (10). Das Verfahren beinhaltet die Bestimmung einer Sollabgastemperatur wenigstens zum Teil auf der Basis eines Verdichterdruckzustandes; der Bestimmung einer Temperaturanpassung an die Sollabgastemperatur wenigstens zum Teil auf der Basis einer Dampffeuchtigkeit; und einer Änderung der Sollabgastemperatur wenigstens zum Teil auf der Basis der Temperaturanpassung.

Burner with a inlet guide vane system.

Die vorliegende Erfindung stellt einen Brenner (100) zur Vermischung eines verdichteten Luftstroms mit einem Brennstoffstrom und Zünden des Gemischs bereit. Der Brenner (100) enthält ein Brennerrohr, das sich von einer an seinem ersten Ende positionierten Endabdeckung zu einem Kappenelement an seinem gegenüberliegenden Ende erstreckt, wobei zwischen der Endabdeckung und dem Kappenelement ein Abstand für einen dadurch verlaufenden inneren Strömungspfad für den verdichteten Luftstrom angeordnet ist, wobei eine Anzahl von in dem Kappenelement positionierten und in Verbindung mit dem inneren Strömungspfad für den verdichteten Luftstrom stehenden Brennstoffdüsen angeordnet ist. Der Brenner enthält ein Einlassleitschaufelsystem (120), das um den inneren Strömungspfad herum positioniert ist, um eine zumindest teilweise verwirbelte Strömung (200) darin zu erzeugen.

Burner with radial inlet guide vanes.

Die vorliegende Anmeldung stellt einen Brenner (100) bereit. Der Brenner (100) enthält einen dadurch verlaufenden inneren Strömungspfad, eine Anzahl von Düsen in Verbindung mit dem inneren Strömungspfad und ein um den inneren Strömungspfad herum positioniertes Einlassleitschaufelsystem (120), um eine verwirbelte Strömung (200) darin zu erzeugen.

SYSTEM FOR IDENTIFICATION OF HOT AND COLD CHAMBERS IN A GAS TURBINE COMBUSTOR

A system for ranking combustion chambers in a gas turbine in order of chamber combustion temperature including: at least one fuel nozzle supplying fuel to the combustion chambers at a predetermined fuel rate abnormally near a lean blow out (LBO) condition of the chambers; at least one dynamic pressure sensor in each chamber sensing combustion dynamic pressures in said combustion chambers and generating dynamic pressure signals representative of the dynamic pressure in each of said combustion chambers, wherein the dynamic pressure signals provide information regarding the dynamic pressure at a plurality of frequencies; a signal band pass filter segmenting the signals into a plurality of predefined frequency bands comprising a lean blow out (LBO) precursor frequency band, a cold tone frequency band and a hot tone frequency band; a processor determining a value for each chamber representative of amplitudes of the signals in each of LBO precursor frequency band, the cold tone frequency band ...

Systeme und Verfahren zur Bestimmung einer Abgas-Zieltemperatur für eine Gasturbine

Ausführungsformen können Systeme und Verfahren zum Bestimmen einer Sollabgastemperatur für eine Gasturbine bereitstellen. In einer Ausführungsform der Beschreibung wird ein Verfahren zum Bestimmen einer Sollabgastemperatur für eine Gasturbine beschrieben. Das Verfahren kann die Bestimmung einer Sollabgastemperatur wenigstens zum Teil auf der Basis eines Verdichterdruckzustandes; der Bestimmung einer Temperaturanpassung an die Sollabgastemperatur wenigstens zum Teil auf der Basis einer Dampffeuchtigkeit; und einer Änderung der Sollabgastemperatur wenigstens zum Teil auf der Basis der Temperaturanpassung beinhalten. Embodiments may provide systems and methods for determining a desired exhaust temperature for a gas turbine. In one embodiment of the description, a method for determining a target exhaust gas temperature for a gas turbine is described. The method may include determining a desired exhaust temperature based at least in part on a compressor pressure condition; determining a temperature adjustment to the target exhaust gas temperature based at least in part on a vapor humidity; and changing the target exhaust temperature based at least in part on the temperature adjustment.

Method and apparatus for identification of hot and cold chambers in a gas turbine combustor

A method for identifying combustion characteristics of a plurality combustion chambers in a gas turbine, the method including: supplying fuel to the combustion chambers at a predetermined fuel rate; sensing combustion dynamic pressure in said plurality of combustion chambers and generating dynamic pressure signals representative of the dynamic pressure in each of said combustion chambers, wherein the dynamic pressure signals provide information regarding the dynamic pressure at a plurality of frequencies; segmenting the signals into a plurality of predefined frequency bands; determining a characteristic value for each of the segmented signals, and identifying an order of combustion chambers based on the characteristic value.

Independent pilot fuel control in secondary fuel nozzle

Disclosed herein is a fuel nozzle. The fuel nozzle includes a first fuel introduction location, a second fuel introduction location, and fuel passages. The first fuel introduction location is located radially about the fuel nozzle and is connected with a fuel passage. The second fuel introduction location is located at an end of the fuel nozzle and is connected with another fuel passage such that the fuel passage connected to the first fuel introduction location is separate from the fuel passage connected to the second fuel introduction location.

FUEL NOZZLE, GAS TURBINE COMBUSTOR, AND METHOD FOR INDEPENDENT PILOT FUEL CONTROL IN SECONDARY FUEL NOZZLE

PROBLEM TO BE SOLVED: To provide a fuel nozzle with optimized operability by reducing the discharge. SOLUTION: Disclosed herein is a fuel nozzle 36. The fuel nozzle 36 includes a first fuel introduction location, a second fuel introduction location, and fuel passages 50. The first fuel introduction location is located radially about the fuel nozzle 36 and is connected with the fuel passage 50. The second fuel introduction location is located at an end of the fuel nozzle 36 and is connected with another fuel passage 50 such that the fuel passage 50 connected to the first fuel introduction location is separate from the fuel passage 50 connected to the second fuel introduction location. COPYRIGHT: (C)2007,JPO&INPIT ...

Radial inlet guide vanes for a combustor

A combustor may include an interior flow path therethrough, a number of fuel nozzles in communication with the interior flow path, and an inlet guide vane system positioned about the interior flow path to create a swirled flow therein. The inlet guide vane system may include a number of windows positioned circumferentially around the fuel nozzles. The inlet guide vane system may also include a number of inlet guide vanes positioned circumferentially around the fuel nozzles and adjacent to the windows to create a swirled flow within the interior flow path.

Method and apparatus for identification of hot and cold chambers in a gas turbine combustor

A method for identifying combustion characteristics of a plurality combustion chambers in a gas turbine, the method including: supplying fuel to the combustion chambers at a predetermined fuel rate; sensing combustion dynamic pressure in said plurality of combustion chambers and generating dynamic pressure signals representative of the dynamic pressure in each of said combustion chambers, wherein the dynamic pressure signals provide information regarding the dynamic pressure at a plurality of frequencies; segmenting the signals into a plurality of predefined frequency bands; determining a characteristic value for each of the segmented signals, and identifying an order of combustion chambers based on the characteristic value.

Systems and methods for detecting fuel leaks in gas turbine engines

Embodiments can provide systems and methods for detecting fuel leaks in gas turbine engines. According to one embodiment, there is disclosed a method for detecting a fuel leak in a gas turbine engine. The method may include adjusting a control valve to correspond with a desired fuel flow. The method may also include determining an actual fuel flow based at least in part on an upstream pressure in a fuel manifold and one or more gas turbine engine parameters. The method may also include comparing the desired fuel flow with the actual fuel flow. Moreover, the method may include determining a difference between the desired fuel flow and the actual fuel flow, wherein the difference indicates a fuel leak.

Independent pilot fuel control in secondary fuel nozzle

Disclosed herein is a fuel nozzle. The fuel nozzle includes a first fuel introduction location, a second fuel introduction location, and fuel passages. The first fuel introduction location is located radially about the fuel nozzle and is connected with a fuel passage. The second fuel introduction location is located at an end of the fuel nozzle and is connected with another fuel passage such that the fuel passage connected to the first fuel introduction location is separate from the fuel passage connected to the second fuel introduction location.

RADIAL INLET GUIDE VANE FOR COMBUSTOR

PROBLEM TO BE SOLVED: To provide a combustor (100) providing a more uniform air flow distribution about the combustor and a combustor cap. SOLUTION: The combustor (100) includes an interior flow path (22) piercing the combustor, a plurality of nozzles (24) in communication with the interior flow path (22), and an inlet guide vane system (120) positioned about the interior flow path (22) to create a swirled flow (200) in the interior flow path. COPYRIGHT: (C)2011,JPO&INPIT ...

Radial inlet guide vanes for a combustor

The present application thus provides a combustor (100). The combustor (100) may include an interior flow path (22) therethrough, a number of nozzles (24) in communication with the interior flow path (22), and an inlet guide vane system (120) positioned about the interior flow path (22) to create a swirled flow (200) therein.

Systems and methods for determining a target exhaust temperature for a gas turbine

Embodiments of the invention can provide systems and methods for determining a target exhaust temperature for gas turbines. In one embodiment of the disclosure, there is disclosed a method for determining a target exhaust temperature for a gas turbine. The method can include determining a target exhaust temperature based at least in part on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based at least in part on steam humidity; and changing the target exhaust temperature based at least in part on the temperature adjustment.

Systems and Methods for Determining a Target Exhaust Temperature for a Gas Turbine

Embodiments of the invention can provide systems and methods for determining a target exhaust temperature for gas turbines. In one embodiment of the disclosure, there is disclosed a method for determining a target exhaust temperature for a gas turbine. The method can include determining a target exhaust temperature based at least in part on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based at least in part on steam humidity; and changing the target exhaust temperature based at least in part on the temperature adjustment. 1. A method for determining a target exhaust temperature for a gas turbine , comprising:determining a target exhaust temperature based at least in part on a compressor pressure condition;determining a temperature adjustment to the target exhaust temperature based at least in part on steam humidity; andchanging the target exhaust temperature based at least in part on the temperature adjustment.2. The method of claim 1 , wherein determining the temperature adjustment is further based on at least one of: CO or NOx emissions reference condition.3. The method of claim 1 , wherein determining the temperature adjustment is further based at least in part on turbine inlet specific humidity and steam humidity.4. The method of claim 1 , wherein determining the temperature adjustment is further based at least in part on delta inlet pressure loss claim 1 , current compressor condition claim 1 , and delta exhaust temperature output.5. The method of claim 1 , wherein determining the temperature adjustment is further based at least in part on delta back pressure claim 1 , current compressor condition claim 1 , and delta exhaust temperature output.6. The method of claim 1 , further comprising:generating a plurality of the target exhaust temperatures; andselecting one of the plurality of target exhaust temperatures to facilitate control of the gas turbine.7. The method of claim 1 , wherein the target exhaust ...

Independent pilot fuel control in secondary fuel nozzle

Disclosed herein is a fuel nozzle (36). The fuel nozzle (36) includes a first fuel introduction location, a second fuel introduction location, and fuel passages (50). The first fuel introduction location is located radially about the fuel nozzle (36) and is connected with a fuel passage (50). The second fuel introduction location is located at an end of the fuel nozzle (36) and is connected with another fuel passage (50) such that the fuel passage (50) connected to the first fuel introduction location is separate from the fuel passage (50) connected to the second fuel introduction location.

System for identification of hot and cold chambers in a gas turbine combustor

A system for ranking combustion chambers in a gas turbine in order of chamber combustion temperature including: at least one fuel nozzle supplying fuel to the combustion chambers at a predetermined fuel rate abnormally near a lean blow out (LBO) condition of the chambers; at least one dynamic pressure sensor in each chamber sensing combustion dynamic pressures in said combustion chambers and generating dynamic pressure signals representative of the dynamic pressure in each of said combustion chambers, wherein the dynamic pressure signals provide information regarding the dynamic pressure at a plurality of frequencies; a signal band pass filter segmenting the signals into a plurality of predefined frequency bands comprising a lean blow out (LBO) precursor frequency band, a cold tone frequency band and a hot tone frequency band; a processor determining a value for each chamber representative of amplitudes of the signals in each of LBO precursor frequency band, the cold tone frequency band ...

COMBUSTOR PROBE FOR GAS TURBINE

Embodiments of the present disclosure are directed towards a turbine combustor probe having a combustion dynamics monitoring probe configured to monitor combustion dynamics within a turbine combustor. The turbine combustor probe also has a gas sampling sleeve configured to collect a gas sample from an airflow path between a liner and a flow sleeve of the turbine combustor. 1. A system , comprising:a turbine combustor; and a combustion dynamics monitoring probe configured to monitor combustion dynamics within the turbine combustor; and', 'a gas sampling sleeve configured to collect a gas sample from an airflow path between a liner and a flow sleeve of the turbine combustor., 'a combustor probe, comprising2. The system of claim 1 , wherein the combustion dynamics monitoring probe comprises a wave guide.3. The system of claim 1 , wherein the gas sampling sleeve comprises a gas sample tube and a gas sample cavity operatively coupled to the gas sample tube.4. The system of claim 1 , wherein the combustion dynamics monitoring probe is disposed within the gas sampling sleeve forming an annular passage between the combustion dynamics monitoring probe and the gas sampling sleeve.5. The system of claim 1 , wherein the combustion dynamics monitoring probe and the gas sampling sleeve are concentric.6. The system of claim 3 , wherein the gas sample tube includes at least one aperture configured to receive the gas sample from the airflow path.7. The system of claim 3 , wherein the gas sample cavity includes at least one lead out port configured to output the gas sample.8. The system of claim 1 , wherein the combustor probe is disposed through a probe port of the turbine combustor claim 1 , such that the combustor probe at least partially enters an airflow passage between the liner and the flow sleeve of the turbine combustor.9. The system of claim 1 , wherein the probe port is located near a head end of the turbine combustor.10. The system of claim 1 , wherein the gas sampling ...

SYSTEMS AND METHODS FOR BULK TEMPERATURE VARIATION REDUCTION OF A GAS TURBINE THROUGH CAN-TO-CAN FUEL TEMPERATURE MODULATION

A gas turbine includes a plurality of combustion chambers; at least one fuel nozzle for each of the combustion chambers; at least one fuel line for each fuel nozzle in each of the combustion chambers; at least one heat exchanger for each fuel line configured to adjust a temperature of a fuel flow to each fuel nozzle; and a controller configured to control each of the heat exchangers to reduce temperature variations amongst the combustion chambers. 1. A gas turbine , comprising:a plurality of combustion chambers;at least one fuel nozzle for each of the combustion chambers;at least one fuel line for each fuel nozzle in each of the combustion chambers;at least one heat exchanger for each fuel line configured to adjust a temperature and an amount of a fuel flow to each fuel nozzle; anda controller configured to control each of the heat exchangers to reduce temperature variations amongst the combustion chambers.2. The gas turbine of claim 1 , further comprising a plurality of manifolds claim 1 , wherein a plurality of fuel lines extend from each manifold to the fuel nozzles in each of the combustion chambers.3. The gas turbine of claim 2 , wherein the plurality of manifolds have different fuel capacities.4. The gas turbine according to claim 2 , wherein at least one of the manifolds of the plurality of manifolds is configured to provide purge gas to at least one fuel nozzle of each combustion chamber.5. The gas turbine of claim 1 , further comprising a plurality of sensors for measuring a condition of exhaust gas from the plurality of combustion chambers.6. The gas turbine of claim 5 , wherein the plurality of sensors comprises a plurality of temperature sensors configured to sense temperatures at different regions of an exhaust outlet of the turbine claim 5 , and the controller correlates the sensed exhaust temperatures to fuel flow to individual combustion chambers and controls the heat exchangers to modify a profile of exhaust gas temperatures.7. The gas turbine of ...

Sector nozzle mounting systems

Systems are provided for mounting sector nozzles within gas turbine combustors. In one embodiment, a sector nozzle includes a nozzle portion configured to mix fuel and air to produce a fuel-air mixture and a shell coupled to the nozzle portion. The sector nozzle also includes a first longitudinal strut and a second longitudinal strut coupled to a first surface of the shell on opposite sides of a window within the first surface. A third longitudinal strut is coupled to a second surface of the shell, and the second surface is disposed opposite of the first surface.

Premixing apparatus for gas turbine system

A premixing apparatus for a gas turbine system includes non-swirl elements around a periphery of a face of a premixing apparatus and a swirl assembly located substantially at a center of the face. The non-swirl elements premix a premixture prior to the premixture being delivered to a combustor of the gas turbine system. The swirl assembly disturbs a flow of fluid prior to the fluid being delivered to the combustor. The premixture includes fuel and oxidant, and the fluid disturbed by the swirl assembly includes the oxidant or the premixture.

COMBUSTOR RECOVERY METHOD AND SYSTEM

A method is disclosed for controlling gas turbine operation in response to lean blowout of a combustion can. The gas turbine comprises a pair of combustion cans. The method includes sensing that a first combustion can is extinguished during a full load operation of the gas turbine, adjusting a fuel ratio between the fuel nozzles in each can, delivering a richer fuel mixture to the fuel nozzles nearest to the cross-fire tubes, generating a cross-fire from the second combustion can to the first combustion can, detecting a recovery of the turbine load, and adjusting the fuel ratio to the normal balanced fuel distribution between the fuel nozzles in each can. 1. A method of controlling gas turbine operation in response to lean blowout of a combustion can , wherein the gas turbine comprises at least two combustion cans , the method comprising:sensing that a first combustion can is extinguished during a full load operation of the gas turbine;adjusting a fuel ratio between a first fuel nozzle associated with the first combustion can and a second fuel nozzle associated with a second combustion can;delivering a richer fuel mixture to the second fuel nozzle, wherein the second fuel nozzle is a fuel nozzle nearest to a cross-fire tube;generating a cross-fire from the second combustion can to the first combustion can via the cross-fire tube;detecting a recovery of the turbine load; andadjusting the fuel ratio to the normal balanced fuel distribution between the respective fuel nozzles.2. The method of claim 1 , wherein the step of adjusting a fuel ratio further comprises:in response to sensing an upset in the operating load, initiating reignition.3. The method of claim 1 , wherein the step of adjusting a fuel ratio further comprises generating a command by a controller to adjust the fuel division between fuel nozzles.4. The method of claim 1 , wherein the step of generating a cross-fire further comprises:causing flame from the second combustion can to bridge across to the first ...

ANNULAR PREMIXED PILOT IN FUEL NOZZLE

A combustor for a gas turbine engine has a head end portion that carries at least one fuel/air nozzle. Each fuel/air nozzle includes a premixed pilot nozzle having premix conduits that are configured with concentric axes that direct the fuel/air mixture axially from the premixed pilot nozzle. The premixed pilot nozzle can include an annular channel disposed radially outwardly from the premix and including air jets that direct air radially outwardly from the premix conduits. 1. A fuel/air nozzle for a gas turbine engine , the fuel/air nozzle comprising:an axially elongating peripheral wall defining an outer envelope of the fuel/air nozzle, the peripheral wall having an outer surface and an inner surface facing opposite the outer surface and defining an axially elongating inner cavity;a hollow, axially elongating centerbody disposed within the inner cavity of the fuel/air nozzle and defining a central axis, the centerbody being defined by a centerbody wall defining an upstream end and a downstream end disposed axially opposite the upstream end, the centerbody wall being defined by an exterior surface and an interior surface facing opposite the exterior surface, the interior surface of the centerbody wall defining an axially elongating interior passage disposed concentrically about the central axis of the centerbody;an elongated, hollow fuel supply line extending axially through the interior passage of the centerbody, the fuel supply line having an upstream end disposed at the upstream end of the centerbody and configured for connection to a source of fuel, the fuel supply line having a downstream end disposed at the downstream end of the centerbody;a primary air flow channel being defined by an annular space between the exterior surface of the centerbody and the inner surface of the peripheral wall;a premixed pilot nozzle having an upstream end connected to the downstream end of the centerbody, the premixed pilot nozzle having a downstream end disposed axially ...

Systems and Methods for Detecting Fuel Leaks in Gas Turbine Engines

Embodiments can provide systems and methods for detecting fuel leaks in gas turbine engines. According to one embodiment, there is disclosed a method for detecting a fuel leak in a gas turbine engine. The method may include adjusting a control valve to correspond with a desired fuel flow. The method may also include determining an actual fuel flow based at least in part on an upstream pressure in a fuel manifold and one or more gas turbine engine parameters. The method may also include comparing the desired fuel flow with the actual fuel flow. Moreover, the method may include determining a difference between the desired fuel flow and the actual fuel flow, wherein the difference indicates a fuel leak.

TURBOMACHINE AND STAGED COMBUSTION SYSTEM OF A TURBOMACHINE

A turbomachine including a combustor in which fuel is combustible to produce a working fluid, a turbine section, which is receptive of the working fluid for power generation operations, a transition piece in which additional fuel is combustible, the transition piece being disposed to transport the working fluid from the combustor to the turbine section and a staged combustion system coupled to the combustor and the transition piece. The staged combustion system is configured to blend components of the fuel and the additional fuel in multiple modes. 1. A turbomachine , comprising:a combustor in which fuel is combustible to produce a working fluid;a turbine section, which is receptive of the working fluid for power generation operations;a transition piece in which additional fuel is combustible, the transition piece being disposed to transport the working fluid from the combustor to the turbine section; anda staged combustion system coupled to the combustor and the transition piece, the staged combustion system being configured to blend components of the fuel and the additional fuel in multiple modes.2. The turbomachine according to claim 1 , wherein the transition piece is curved.3. The turbomachine according to claim 1 , wherein a flowpath through the combustor is offset from a flowpath through the turbine section claim 1 , the transition piece comprising:a forward end aligned with the flowpath through the combustor; andan aft end aligned with the flowpath through the turbine section.4. The turbomachine according to claim 1 , wherein the staged combustion system comprises:head end injectors disposable to deliver the fuel to a head end of the combustor; andaxially staged injectors disposable to deliver the additional fuel to downstream sections of the combustor and the transition piece.5. The turbomachine according to claim 4 , wherein the axially staged injectors are arranged in a first stage and a second stage disposed downstream from the first stage.6. The ...

TURBOMACHINE WITH TRANSITION PIECE HAVING DILUTION HOLES AND FUEL INJECTION SYSTEM COUPLED TO TRANSITION PIECE

A turbomachine is provided and includes a combustor in which fuel and air are combusted, a turbine disposed for reception of products of combustion from the combustor, a transition piece fluidly interposed between the combustor and the turbine and including a body formed to define dilution holes configured to allow air to enter the combustor and for enabling steam injection toward a main flow of the products of the combustion proceeding from the combustor to the turbine and a fuel injection system supportively coupled to the transition piece and configured to inject fuel toward the main flow of the products of the combustion to thereby restore a flame temperature of the main flow of the products of the combustion reduced by steam injection enabled by the dilution holes. 1. A turbomachine , comprising:a combustor in which fuel and air are combusted;a turbine disposed for reception of products of combustion from the combustor;a transition piece fluidly interposed between the combustor and the turbine and including a body formed to define dilution holes configured to allow air to enter the combustor and for enabling steam injection toward a main flow of the products of the combustion proceeding from the combustor to the turbine; anda fuel injection system supportively coupled to the transition piece and configured to inject fuel toward the main flow of the products of the combustion to thereby restore a flame temperature of the main flow of the products of the combustion reduced by steam injection enabled by the dilution holes.2. The turbomachine according to claim 1 , wherein the body of the transition piece comprises:a forward portion, which is coaxial with an aft portion of the combustor;an aft portion, which is coupled to a forward portion of the turbine; anda central portion, which extends curvilinearly from the forward portion to the aft portion.3. The turbomachine according to claim 1 , wherein the fuel injection system comprises:at least one or more fuel ...

ANNULAR PREMIXED PILOT IN FUEL NOZZLE

A combustor for a gas turbine engine has a head end portion that carries at least one fuel/air nozzle. Each fuel/air nozzle includes a premixed pilot nozzle having premix conduits where each premix conduit has an upstream end defining an opening and having a central axis and downstream end defining an outlet and having a central axis where the central axis of the downstream end is non-parallel with a central axis of a center body of the fuel/air nozzle. The premixed pilot nozzle can include an annular channel disposed radially outwardly from the premix conduits and may include air jets that direct air radially outwardly from the premix conduits. 1. A fuel/air nozzle for a gas turbine engine , the fuel/air nozzle comprising:a premixed pilot nozzle having an upstream end axially spaced from a downstream end with respect to an axial center line of a center body of the fuel/air nozzle, wherein the upstream end is connected to a downstream end of the center body; andthe premixed pilot nozzle further defining a plurality of axially elongated, hollow premix conduits annularly arranged around a pilot fuel nozzle portion of the premixed pilot nozzle, each premix conduit having an upstream end defining an entrance opening that communicates fluidly with an interior passage of the center body, each premix conduit having at least one fuel hole in fluid communication with the pilot fuel nozzle portion, each premix conduit including a downstream end defining an exit opening that allows fluid to discharge from the hollow premix conduit;wherein each downstream end of each premix conduit has a central axis that is not parallel to the central axis of the center body.2. The fuel/air nozzle as in claim 1 , wherein each central axis of at least one premix conduit is disposed at an acute angle with respect to the central axis of the center body.3. The fuel/air nozzle as in claim 1 , wherein the premixed pilot nozzle further defines an annular channel disposed radially outwardly from the ...

FUEL STAGING IN A GAS TURBINE ENGINE

A gas turbine system includes a turbine combustor. The turbine combustor includes one or more first fuel injectors coupled to a head end of the turbine combustor and configured to deliver a first portion of a fuel to a first axial position of a combustion section of the turbine combustor. The turbine combustor also includes one or more second fuel injectors coupled to the turbine combustor axially downstream of the head end and configured to deliver a second portion of the fuel to a second axial position of the combustion section of the turbine combustor, the second axial position being downstream of the first axial position. The gas turbine system also includes a controller configured to deliver the first and the second portions of the fuel such that combustion of the fuel and oxidant within the overall combustion section of the turbine combustor is at a defined equivalence ratio. 1. A gas turbine system comprising: one or more first fuel injectors coupled to a head end of the turbine combustor and configured to deliver a first portion of a fuel to a first axial position of a combustion section of the turbine combustor;', 'one or more second fuel injectors coupled to the turbine combustor axially downstream of the head end, wherein the one or more second fuel injectors are configured to deliver a second portion of the fuel to a second axial position of the combustion section of the turbine combustor, the second axial position being downstream of the first axial position;, 'a turbine combustor, comprisingan oxidant supply system configured to provide an oxidant to a head end of the combustion section of the turbine combustor;a fuel supply system configured to supply the fuel to the one or more first and second fuel injectors; anda controller configured to cause the one or more first fuel injectors to deliver the first portion of the fuel such that combustion of the fuel and oxidant at the first axial position is oxidant-rich, and to cause the one or more second fuel ...

Fuel staging in a gas turbine engine

A gas turbine system includes a turbine combustor. The turbine combustor includes one or more first fuel injectors coupled to a head end of the turbine combustor and configured to deliver a first portion of a fuel to a first axial position of a combustion section of the turbine combustor. The turbine combustor also includes one or more second fuel injectors coupled to the turbine combustor axially downstream of the head end and configured to deliver a second portion of the fuel to a second axial position of the combustion section of the turbine combustor, the second axial position being downstream of the first axial position. The gas turbine system also includes a controller configured to deliver the first and the second portions of the fuel such that combustion of the fuel and oxidant within the overall combustion section of the turbine combustor is at a defined equivalence ratio.

Combustor with Pre-Mixing Primary Fuel-Nozzle Assembly

A combustor includes a first combustion chamber, a pre-mixing primary fuel-nozzle assembly associated with the first combustion chamber, a second combustion chamber, and a secondary fuel-nozzle assembly associated with the second combustion chamber. The pre-mixing primary fuel-nozzle assembly includes a number of vanes configured to swirl airflow, each vane comprising a number of fuel injection holes configured to inject fuel into the airflow.

Inboard radial dump venturi for combustion chamber of a gas turbine

A double wall venturi chamber having a converging section, a diverging section and a cylindrical section wherein said chamber defines a venturi zone in which compressed air, fuel and combustion products flow downstream through converging section, diverging section and cylindrical section, and has a cooling gas passage between the walls of the venturi chamber, a least one cooling gas inlet in an outlet wall of the venturi chamber, and at least one cooling gas outlet in an inner wall of the venturi chamber, wherein said cooling gas outlet is in at least one of the diverging and the cylindrical section, and the outlet is downstream of the at least one cooling gas inlet and upstream of an axial end of the chamber.

Turbocharger with transition piece with dilution holes and fuel injection system, which is connected to the transition piece

Es ist eine Turbomaschine geschaffen, und diese enthält eine Brennkammer, in der Brennstoff und Luft verbrannt werden, eine Turbine, die zur Aufnahme von Verbrennungsprodukten aus der Brennkammer angeordnet ist, ein Übergangsstück, das strömungsmäßig zwischen der Brennkammer und der Turbine eingefügt ist und einen Körper enthält, der ausgebildet ist, um Verdünnungslöcher zu definieren, die konfiguriert sind, um Luft zu gestatten, in die Brennkammer einzutreten, und um eine Dampfinjektion in Richtung auf eine Hauptströmung der Produkte der Verbrennung, die von der Brennkammer zu der Turbine fortschreiten, zu ermöglichen, und ein Brennstoffeinspritzsystem, das unterstützend mit dem Übergangsstück verbunden und konfiguriert ist, um Brennstoff in Richtung auf die Hauptströmung der Produkte der Verbrennung zu injizieren, um dadurch eine Flammentemperatur der Hauptströmung der Produkte der Verbrennung, die durch eine durch die Verdünnungslöcher ermöglichte Dampfinjektion reduziert wurde, wiederherzustellen.

Fuel staging in a gas turbine engine

A gas turbine system includes a turbine combustor (160). The turbine combustor includes one or more first fuel injectors coupled to a head end (166) of the turbine combustor and configured to deliver a first portion of a fuel to a first axial position of a combustion section of the turbine combustor. The turbine combustor also includes one or more second fuel injectors (528) coupled to the turbine combustor axially downstream of the head end and configured to deliver a second portion of the fuel to a second axial position of the combustion section of the turbine combustor, the second axial position being downstream of the first axial position. The gas turbine system also includes a controller (100) configured to deliver the first and the second portions of the fuel such that combustion of the fuel and oxidant within the overall combustion section of the turbine combustor is at a defined equivalence ratio.

Independent pilot fuel control in secondary fuel nozzle

Disclosed herein is a fuel nozzle (36). The fuel nozzle (36) includes a first fuel introduction location, a second fuel introduction location, and fuel passages (50). The first fuel introduction location is located radially about the fuel nozzle (36) and is connected with a fuel passage (50). The second fuel introduction location is located at an end of the fuel nozzle (36) and is connected with another fuel passage (50) such that the fuel passage (50) connected to the first fuel introduction location is separate from the fuel passage (50) connected to the second fuel introduction location.

Method of mitigating undesired gas turbine transient response using event based actions

A method of managing transient events regularly seen during gas turbine 10 operation that may cause undesirable operation and hardware damage. During certain transient operations, a lag may be seen between reference exhaust temperature and actual turbine exhaust temperature. This lag can result in an under-fired condition within the combustion system of variable magnitude and duration. Either fuel split schedules or a control algorithm can be positioned during these transients to prevent combustion dynamics or loss of flame. Combustion dynamics are known to cause damage that may require hardware replacement. Once the transient has completed, normal control operation is resumed.

Radial Inlet Guide Vanes for a Combustor

The present application thus provides a combustor. The combustor may include an interior flow path therethrough, a number of nozzles in communication with the interior flow path, and an inlet guide vane system positioned about the interior flow path to create a swirled flow therein.

Turbine combustor transition piece having dilution holes

A transition piece body, having an inlet end for receiving combustion products from a turbine combustor and an outlet end for flowing the gaseous products into a first stage nozzle, has dilution holes in zones respectively adjacent the transition piece body inlet and outlet ends. The volume of dilution air flowing into the gas stream is substantially equal at the inlet and outlet ends of the transition piece. The locations and sizes of the openings are given in the respective X, Y, Z coordinates and hole diameters in Table I. The X and Y coordinates lie in the circular plane of the transition body at its inlet end and the Z coordinates extend in the direction of gas flow from the origin.

Fuel staging in a gas turbine engine

Apparatus and method for minimzing and/or eliminating dilution air leakage in a combustion liner assembly

A combustion liner assembly for a gas turbine includes an outer liner, the outer liner having a flange at a forward end. An inner liner is disposed within the outer liner. The inner liner has a first inner wall. A venturi includes a second inner wall, a venturi throat, and the first inner wall of the inner liner. A slip joint is connected to the second inner wall. The slip joint receives the flange of the outer liner. Alternatively, or additionally, the combustion liner assembly includes a slip joint between the inner or outer liner and an aft section.

APPARATUS/METHOD FOR COOLING COMBUSTION CHAMBER/VENTURI IN A LOW NOx COMBUSTOR

A dry low nitric oxides (NOx) emissions combustor includes a premixing chamber for mixing fuel and cooling gas and a combustion chamber positioned downstream of the premixing chamber for the combustion of pre-mixed fuel and cooling gas. The combustor further includes a venturi having generally annular walls including converging and diverging wall portions that define a constricted portion and positioned between the premixing chamber and the combustion chamber through which the premixed fuel and air pass to the combustion chamber. The walls defining a passage for cooling gas flow extending axially along the combustion chamber and having an exit for flowing cooling gas to the combustion chamber. A plurality of inlets at the converging and diverging wall portions ingest cooling gas into the passage to produce an impingement cooling effect. A plurality of tubulators disposed downstream of the inlets interact with the cooling gas to produce a turbulated cooling effect. The combustor may be effectively fired over a substantial temperature range to reduce the NOx emissions of the combustor.

Combustor assembly for use in a gas turbine engine and method of assembling same

A combustor assembly for use in a gas turbine engine and method of assembly is described. The combustor assembly includes a combustor liner having a slot that at least partially circumscribes the combustor liner. The slot is defined adjacent to a venturi throat region defined within the liner. The combustor assembly also includes a restrictor plate having at least one aperture defined therein. The restrictor plate is removably coupled within the combustor assembly such that the restrictor plate is inserted within the slot and extends at least partially across the venturi throat region.

Method for controlling fuel splits to gas turbine combustor

A method for determining a target exhaust temperature for a gas turbine including: determining a target exhaust temperature based on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based on at least one parameter of a group of parameters consisting of specific humidity, compressor inlet pressure loss and turbine exhaust back pressure; and adjusting the target exhaust temperature by applying the temperature adjustment.

Gas turbine combustor flame stabilizer

A gas turbine combustor is presented, which includes a combustion chamber that is positioned downstream of a premixing chamber. The premixing chamber includes at least one opening for ingesting air. At least one primary fuel nozzle is disposed to discharge fuel into the premixing chamber. The fuel discharged from the primary fuel nozzle mixes with the ingested air in the premixing chamber to provide a fuel air mix. A secondary fuel nozzle is disposed proximate the combustion chamber to discharge fuel at the combustion chamber. A stabilizer is disposed at the secondary fuel nozzle so as to be positioned in close proximity to a flame when fuel at the secondary fuel nozzle is ignited. The stabilizer is composed of a material having the ability to absorb heat from a heat flux generated within the combustor and maintaining a temperature sufficient to sustain ignition of the flame. A method of stabilizing a flame in a gas turbine combustor is also presented. The method including discharging fuel at a combustion chamber of the gas turbine combustor and positioning a stabilizer in close proximity to a flame when the fuel at a combustion chamber is ignited. The stabilizer absorbing heat from a heat flux generated within the combustor and maintaining a temperature sufficient to sustain ignition of the flame.

Combustion chamber with cooled venturi.

Eine Brennkammer (30) mit trockenen niedrigen Stickoxid-(NOx-) Emissionen schliesst eine Vormischkammer (36) zum Mischen von Kraftstoff (31) und Kühlgas und eine Verbrennungskammer (32), die stromabwärts von der Vormischkammer (36) angeordnet ist, für die Verbrennung des vorgemischten Kraftstoffs und Kühlgases, ein. Die Brennkammer (30) schliesst ferner ein Venturirohr (46) ein, das im Allgemeinen ringförmige Wände hat, die zusammenlaufende (41) und auseinanderlaufende (43) Wandabschnitte einschliessen, die einen eingeengten und zwischen der Vormischkammer (36) und der Verbrennungskammer (32) angeordneten Abschnitt definieren, durch den der vorgemischte Kraftstoff und die Luft zu der Verbrennungskammer (32) hindurchgehen. Die Wände definieren einen Durchgang (44) für einen Kühlgasstrom, der sich in Axialrichtung längs der Verbrennungskammer (32) erstreckt und einen Auslass (64) hat, um das Kühlgas zu der Verbrennungskammer (32) strömen zu lassen. Mehrere Einlässe (56) an den zusammenlaufenden (41) und auseinanderlaufenden (43) Wandabschnitten nehmen Kühlgas in den Durchgang (44) auf, um eine Aufprallkühlwirkung zu erzeugen. Mehrere stromabwärts von den Einlässen (56) angeordnete Turbulenzgeneratoren (62) treten in Wechselwirkung mit dem Kühlgas, um eine Turbulenzkühlwirkung zu erzeugen. Die Brennkammer (30) kann über einen beträchtlichen Temperaturbereich wirksam befeuert werden, um die NOx-Emissionen der Brennkammer (30) zu verringern. A dry low nitrogen oxide (NOx) emissions combustor (30) includes a premixing chamber (36) for mixing fuel (31) and cooling gas and a combustion chamber (32) located downstream of the premixing chamber (36) for Combustion of the premixed fuel and refrigerant gas, a. The combustor (30) further includes a venturi (46) having generally annular walls including converging (41) and diverging (43) wall sections defining a restricted and intermediate the premixing chamber (36) and the combustion chamber (32). define a portion through which ...

Flame stabilizer for a gas turbine burner

Es wird ein Gasturbinenbrenner (10) vorgestellt, der eine Brennkammer (12) enthält, die stromabwärts von einer Vormischkammer (18) angeordnet ist. Die Vormischkammer (18) enthält wenigstens eine Öffnung (50) zur Aufnahme von Luft. Wenigstens eine primäre Brennstoffdüse (14) ist angeordnet, um einen Brennstoff (54) in die Vormischkammer (18) hinein austreten zu lassen. Der aus der primären Brennstoffdüse (14) abgegebene Brennstoff (54) vermischt sich mit der aufgenommenen Luft in der Vormischkammer (18), um ein Brennstoff-Luft-Gemisch zu bilden. Eine sekundäre Brennstoffdüse (16) ist in der Nähe der Brennkammer (12) angeordnet, um den Brennstoff (54) an der Brennkammer (12) abzugeben. Ein Stabilisator (32) ist an der sekundären Brennstoffdüse (16) angeordnet, um in unmittelbarer Nähe zu einer Flamme positioniert zu sein, wenn der Brennstoff (54) an der sekundären Brennstoffdüse (16) gezündet wird. Der Stabilisator (32) ist aus einem Material gebildet, das die Fähigkeit aufweist, Wärme aus einem in dem Brenner (10) erzeugten Wärmestrom zu absorbieren und eine Temperatur beizubehalten, die ausreicht, um die Zündung der Flamme aufrechtzuerhalten. Es wird auch ein Verfahren zur Stabilisierung einer Flamme in einem Gasturbinenbrenner (10) vorgestellt. Das Verfahren enthält das Ausgeben von Brennstoff (54) an einer Brennkammer (12) des Gasturbinenbrenners (10) und das Positionieren eines Stabilisators (32) in unmittelbarer Nähe zu einer Flamme, wenn der Brennstoff (54) an der ... A gas turbine combustor (10) is presented which includes a combustor (12) located downstream of a premix chamber (18). The premixing chamber (18) contains at least one opening (50) for receiving air. At least one primary fuel nozzle (14) is arranged to allow a fuel (54) to exit into the premixing chamber (18). The fuel (54) discharged from the primary fuel nozzle (14) mixes with the intake air in the premixing chamber (18) to form a fuel-air mixture. A secondary fuel nozzle (16) is disposed in the ...

Method for controlling fuel splits to gas turbine combustor

A method for determining a target exhaust temperature for a gas turbine including: determining a target exhaust temperature based on a compressor pressure condition; determining a temperature adjustment to the target exhaust temperature based on at least one parameter of a group of parameters consisting of specific humidity, compressor inlet pressure loss and turbine exhaust back pressure; and adjusting the target exhaust temperature by applying the temperature adjustment.

Inboard radial dump venturi for combustion chamber of a gas turbine

A double wall venturi chamber (60) having a converging section (66), a diverging section (68) and a cylindrical section (59) wherein said chamber defines a venturi zone in which compressed air, fuel and combustion products flow downstream through the converging section, diverging section and cylindrical section, and has a cooling gas passage (64) between the walls of the venturi chamber, a least one cooling gas inlet (72) in an outlet wall of the venturi chamber, and at least one cooling gas outlet (74) in an inner wall of the venturi chamber, wherein said cooling gas outlet is in at least one of the diverging and the cylindrical section, and the outlet is downstream of the at least one cooling gas inlet and upstream of an axial end (76) of the chamber.

Premixing apparatus for gas turbine system

Radiale Einlassleitschaufeln für einen Brenner

Brenner (100), aufweisend:einen inneren Strömungspfad (22) durch diesen;mehrere sich axial erstreckende und radial beabstandete rohrförmige Düsen (24) in Verbindung mit dem inneren Strömungspfad (22); undein Einlassleitschaufelsystem (120), das um den inneren Strömungspfad (22) herum und stromaufwärts der mehreren rohrförmigen Düsen (24) positioniert ist, wobei das Einlassleitschaufelsystem (120) aufweist:mehrere Fenster (170), die in Umfangsrichtung um die mehreren rohrförmigen Düsen (24) herum und stromaufwärts von diesen angeordnet sind; undmehrere einstellbare Einlassleitschaufeln (130), die in Umfangsrichtung um die mehreren rohrförmigen Düsen (24) herum und stromaufwärts von diesen, benachbart zu den mehreren Fenstern (170) angeordnet sind;wobei die mehreren Fenster (170) und die mehreren einstellbaren Einlassleitschaufeln (130) eingerichtet sind, um einer Luftströmung (190) zu ermöglichen, aus einem äußeren Strömungspfad (30) durch die mehreren Fenster (170) in den inneren Strömungspfad (22) hinein zu strömen, und dabei die in den inneren Strömungspfad (22) eintretende Luftströmung (190) in eine Verwirbelung zu versetzen, um eine gleichmäßigere Strömungsverteilung an den mehreren rohrförmigen Düsen (24) stromaufwärts von Einlässen der mehreren rohrförmigen Düsen (24) zu schaffen.

Drallwinkel einer Sekundärbrennstoffdüse für eine Turbomaschinen-Brennkammer

Eine Brennkammer (10) enthält einen Primärbrennraum (12) und einen Sekundärbrennraum (14), eine oder mehrere Primärdüsen (24), die in dem Primärbrennraum (12) angeordnet sind und dem Primärbrennraum (12) Brennstoff zuführen, eine Mittenkörperbaugruppe (30), ein Venturi-Rohr (16), das stromabwärts von der Mittenkörperbaugruppe (30) angeordnet ist, und eine Sekundärbrennstoffdüse (40), die in der Mittenkörperbaugruppe (30) untergebracht ist und sich zu dem Venturi-Rohr (16) hin erstreckt und dem Sekundärbrennraum (14) Brennstoff zuführt. Die Sekundärbrennstoffdüse (40) enthält einen Brennstoffkanal (44) und einen Luftkanal (46) und eine Dralleinrichtung (10), die um den Brennstoffkanal (44) herum positioniert ist und einen oder mehrere Leitflügel (105) aufweist, die radial in den Luftkanal (46) ragen, wobei jeder Leitflügel (105) eine Hinterkante (110) besitzt, die in einem Drallwinkel in Bezug auf eine Längsachse der Sekundärbrennstoffdüse (40) angeordnet ist, wobei der Drallwinkel größer als 45° ist.