Patent application number | Description | Published |
20090019855 | Low emissions gas turbine combustor - A gas turbine combustor including: a primary combustion chamber; a secondary combustion chamber downstream of the primary combustion chamber; a venturi having a venturi throat; a transition piece; a cap assembly attached to the primary combustion chamber, and an external turbulator member in operable communication with the cap assembly, wherein the primary combustion chamber includes a mixing hole arrangement for improving homogeneity of an air and fuel mixture in the combustor; the venturi throat is disposed within a predetermined distance upstream from the downstream end of the primary combustion chamber; the transition piece is composed of a duct body, with a plurality of dilution holes formed in the duct body; and the external turbulator member includes a step positioned at the second end of the centerbody, the step defining a radial distance about the second end of the centerbody. | 01-22-2009 |
20100018181 | CENTERBODY CAP FOR A TURBOMACHINE COMBUSTOR AND METHOD - A turbomachine includes a combustor assembly, a cap assembly attached to the combustor assembly, a centerbody within the cap assembly, a wall of the centerbody having a first end, a second end and an intermediate portion, and an external turbulator member in operable communication with the cap assembly. The external turbulator member is spaced from the wall to form a passage defined by a gap between the wall of the centerbody and the external turbulator. The external turbulator member includes a step positioned at the second end of the centerbody. The step defines a radial distance about the second end of the centerbody. The external turbulator member is formed having a step-to-gap ratio relative to the centerbody in a range of about 0.8 to about 1.2. | 01-28-2010 |
20100018211 | Gas turbine transition piece having dilution holes - A gas turbine transition piece includes a duct body having a forward end and an aft end, the duct body defining an enclosure for confining a flow of combustion products from a combustor to a turbine first stage nozzle. A plurality of dilution holes are formed in the duct body, located at selected X, Y, Z coordinates measured from a zero reference point at a center of an exit plane of the transition piece. | 01-28-2010 |
20100043441 | METHOD AND APPARATUS FOR ASSEMBLING GAS TURBINE ENGINES - A method for assembling a gas turbine engine includes coupling a transition piece between a combustor liner and a nozzle assembly. The method also includes extending a first portion of a flow sleeve from the transition piece about at least a portion of the combustor liner. The method further includes coupling a second portion of the flow sleeve to the first portion of the flow sleeve such that the flow sleeve second portion extends from the flow sleeve first portion and at least partially about at least a portion of the transition piece. The flow sleeve second portion includes a scoop that cooperates with the transition piece to at least partially define a unitary cooling air passage that includes a unitary scoop-shaped opening. The scoop is oriented to introduce a substantially uniform cooling air flow to the transition piece. | 02-25-2010 |
20100205970 | Systems, Methods, and Apparatus Providing a Secondary Fuel Nozzle Assembly - Provided are systems, methods, and apparatus providing secondary fuel nozzle assemblies. For example, a secondary fuel nozzle assembly can include a central portion having a proximal end and distal end, and defining a central passage therethrough, which can include at least one coiled tube extending through the central passage from the proximal end to the distal end; a flange having at least one main secondary fuel orifice in fluid communication with the central passage at the proximal end and at least one pilot orifice in fluid communication with the at least one coiled tube at the proximal end; and a tip portion having a proximal end and distal end, and defining at least one passage therethrough, which can be in fluid communication with the distal end of the at least one coiled tube and at least one orifice formed in the distal end of the tip portion. | 08-19-2010 |
20110041507 | Integral Liner and Venturi for Eliminating Air Leakage - A combustion liner assembly for a gas turbine combustor includes a plurality of fuel nozzles disposed circumferentially about a central axis of the combustor, and a venturi section disposed downstream of the fuel nozzles and connected to a head end of the liner assembly. The venturi section defines an annular throat area downstream of the fuel nozzles. A liner sleeve is connected to and commences at a downstream end of the venturi section. At least a portion of the venturi section serves as a liner upstream of the liner sleeve. | 02-24-2011 |
20120204571 | COMBUSTOR AND METHOD FOR INTRODUCING A SECONDARY FLUID INTO A FUEL NOZZLE - A combustor is disclosed that includes a baffle plate and a fuel nozzle extending through the baffle plate. The combustor may also include a shroud extending from the baffle plate and surrounding at least a portion of the fuel nozzle. A passage may be defined between the shroud and an outer surface of the fuel nozzle for receiving a first fluid. Additionally, the passage may be sealed from a second fluid flowing adjacent to the shroud. | 08-16-2012 |
20120210717 | APPARATUS FOR INJECTING FLUID INTO A COMBUSTION CHAMBER OF A COMBUSTOR - A combustor is disclosed having a combustion liner defining a combustion chamber. The combustor may also include a liner cap disposed upstream of the combustion chamber. The liner cap may include a first plate and a second plate. Additionally, the combustor may include a fluid conduit extending between the first and second plates. The fluid conduit may be configured to receive fluid flowing adjacent to the first plate and inject the fluid into the combustion chamber. | 08-23-2012 |
20120234011 | GAS TURBINE COMBUSTOR HAVING A FUEL NOZZLE FOR FLAME ANCHORING - A combustor includes an end cover having a nozzle. The nozzle has a front end face and a central axis. The nozzle includes a plurality of fuel passages and a plurality of oxidizer passages. The fuel passages are configured for fuel exiting the fuel passage. The fuel passages are positioned to direct fuel in a first direction, where the first direction is angled inwardly towards the center axis. The oxidizer passages are configured for having oxidizer exit the oxidizer passages. The oxidizer passages are positioned to direct oxidizer in a second direction, where the second direction is angled outwardly away from the center axis. The plurality of fuel passages and the plurality of oxidizer passages are positioned in relation to one another such that fuel is in a cross-flow arrangement with oxidizer to create a burning zone in the combustor. | 09-20-2012 |
20130025253 | REDUCTION OF CO AND O2 EMISSIONS IN OXYFUEL HYDROCARBON COMBUSTION SYSTEMS USING OH RADICAL FORMATION WITH HYDROGEN FUEL STAGING AND DILUENT ADDITION - A method for reducing the amount of carbon monoxide and oxygen emissions in an oxyfuel hydrocarbon combustion system, comprising the steps of feeding defined amounts of hydrocarbon fuel and an oxidizer (e.g., air) to one or more combustors in the engine and igniting the mixture to form a first combustor exhaust stream; determining the amount of carbon monoxide present at the head end of a combustor in the initial combustor exhaust stream; identifying one or more target locations within the combustor at a point downstream from the first exhaust stream for injecting free hydrogen and a supplemental oxidizer; injecting hydrogen and the supplemental oxidizer into the combustor at specified downstream locations based on the amount of detected carbon monoxide; and injecting a diluent (e.g., CO | 01-31-2013 |
20130167548 | METHOD AND APPARATUS FOR OPERATING A GAS TURBINE ENGINE - A gas turbine includes a compressor, a combustor downstream from the compressor and a heat transfer system, wherein the heat transfer system receives a compressed working fluid from the compressor. A fluid coupling between the heat transfer system and the combustor, wherein the fluid coupling receives the compressed working fluid from the heat transfer system. A conditioner in fluid communication with the compressor and a fluid coupling between the heat transfer system and the conditioner, wherein the fluid coupling receives a cooling media from the heat transfer system. A method for operating the gas turbine includes flowing a compressed working fluid from the compressor to the heat transfer system, transferring heat energy from the compressed working fluid to the heat transfer system, flowing the compressed working fluid from the heat transfer system to a combustor, and flowing a cooling media from the heat transfer system to a compressor inlet. | 07-04-2013 |
20140260309 | SYSTEM AND METHOD FOR FUEL BLENDING AND CONTROL IN GAS TURBINES - A system includes a gas turbine engine having a combustor, and a fuel blending system. The fuel blending system further includes a first fuel supply configured to supply a first fuel, a second fuel supply configured to supply a second fuel, a first fuel circuit, a second fuel circuit, and a controller. The first fuel circuit may be configured to blend the first fuel and the second fuel to form a first fuel mixture. The second fuel circuit may be configured to blend the first fuel and the second fuel to form a second fuel mixture. The controller may be configured to regulate blending of the first fuel mixture and the second fuel mixture based on a measured composition of the first fuel. | 09-18-2014 |
20140283523 | SYSTEM AND METHOD FOR CONTROLLED FUEL BLENDING IN GAS TURBINES - A system includes a gas turbine engine having a combustor, and a fuel blending system. The fuel blending system further includes a first fuel supply configured to supply a first fuel, a second fuel supply configured to supply a second fuel, a first fuel circuit, a second fuel circuit, and a controller. The first fuel circuit may be configured to blend the first fuel and the second fuel to form a first to form a first fuel mixture. The second fuel circuit may be configured to blend the first fuel and the second fuel to form a second fuel mixture. The controller may be configured to regulate blending of the first fuel mixture and the second fuel mixture based on a measured operating parameter of the combustor. | 09-25-2014 |
20140318134 | BACKUP FUEL SUPPLY FOR A GAS TURBINE - A system for supplying a gaseous fuel to a gas turbine includes a liquefied fuel source for supplying a liquefied fuel to a liquid fuel pump that is disposed downstream from the liquefied fuel source. The liquid fuel pump is sufficient to raise the pressure of the liquefied fuel to a substantially supercritical pressure. A supercritical liquefied fuel vaporizer is disposed downstream from the liquid fuel pump. A heat recovery system is in thermal communication with the liquefied fuel. The heat recovery system is positioned between the liquid fuel source and the supercritical liquefied fuel vaporizer. | 10-30-2014 |
20150047360 | SYSTEM FOR INJECTING A LIQUID FUEL INTO A COMBUSTION GAS FLOW FIELD - A system for injecting a liquid fuel into a combustion gas flow field includes an annular liner that defines a combustion gas flow path. The annular liner includes an inner wall, an outer wall and a fuel injector opening that extends through the inner wall and the outer wall. The system further includes a gas fuel injector that is coaxially aligned with the fuel injector opening. The gas fuel injector includes an upstream end and a downstream end. The downstream end terminates substantially adjacent to the inner wall. A dilution air passage is at least partially defined by the gas fuel injector. A liquid fuel injector extends partially through the dilution air passage. The liquid fuel injector includes an injection end that terminates upstream from the inner wall. | 02-19-2015 |
Patent application number | Description | Published |
20090260340 | Combustor of a Turbine, a Method of Retro-Fitting a Combustor of a Turbine and a Method of Building a Combustor of a Turbine - A combustor of a turbine, a method of retro-fitting a combustor of a turbine and a method of building a combustor of a turbine are each provided. The combustor includes a combustion chamber along which a dilution breach is defined, a casing perimetrically surrounding the combustion chamber so as to define an airflow between the casing and the combustion chamber, the airflow being configured to supply dilution air to the combustion chamber via the dilution breach, and an easy to adjust dilution airflow tuning part disposed on the casing and in communication with the airflow. The dilution airflow tuning part is configured to increase and/or decrease an available amount of the dilution air to be supplied to the combustion chamber. | 10-22-2009 |
20090282833 | Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface - A combustor liner includes a forward end and an aft end, the aft end having a reduced diameter portion and a cooling and dilution sleeve overlying the reduced diameter portion thereby establishing a cooling plenum therebetween. A plurality of cooling and dilution air entry holes are formed in the cooling and dilution sleeve and a plurality of cooling and dilution air exit holes formed adjacent an aft edge of the liner such that, in use, cooling and dilution air flows through the cooling and dilution air entry holes, and through the plenum, exiting the cooling and dilution air exit holes, thereby cooling and dilution tuning the aft end of the combustor liner without having to remove the transition piece. | 11-19-2009 |
20110162373 | FUEL NOZZLE FOR A TURBINE ENGINE WITH A PASSIVE PURGE AIR PASSAGEWAY - A secondary fuel nozzle for a turbine includes a passive purge air passageway which provides purge air to the secondary nozzle at all times that the nozzle is in operation. The passive purge air passageway draws in air from a location adjacent an upstream end of the nozzle. Because of a pressure differential between air located at the downstream end of the nozzle and air located at the upstream end of the nozzle, purge air will run through the passive purge air passageway at all times the nozzle is in operation. There is no need for a supply of compressed purge air. | 07-07-2011 |
20110247340 | APPARATUS AND METHOD FOR MINIMIZING 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. | 10-13-2011 |
20120023956 | POWER PLANT AND METHOD OF OPERATION - A power plant and method of operation is provided. The power plant comprises at least one main air compressor, an oxidizer unit configured to deliver a compressed oxygen-rich gas flow to at least one gas turbine assembly. Each assembly comprises a turbine combustor for mixing the compressed oxygen-rich gas flow with a recirculated gas flow and a fuel stream to burn a combustible mixture and form the recirculated gas flow. The assembly also comprises a recirculation loop for recirculating the recirculated gas flow from a turbine to a turbine compressor. The assembly further comprises a recirculated gas flow extraction path for extracting a portion of the recirculated gas flow from the assembly and delivering this to a gas separation system. The gas separation system separates the portion of the recirculated gas flow into a nitrogen portion and a carbon dioxide portion. | 02-02-2012 |
20120036863 | METHOD, APPARATUS AND SYSTEM FOR DELIVERY OF WIDE RANGE OF TURBINE FUELS FOR COMBUSTION - In operating a gas turbine, there can be a difference between the desired heating value of the fuel and the actual needs of the fuel for sustainable combustion during various stages of the turbine operation. In one aspect, combustible lean limit operation of the gas turbine free of lean blow out is enabled by adjusting fuel-air-ratio of the fuel and fuel-air mixture properties, based on the operation requirements of the turbine and flammability of the fuel components. | 02-16-2012 |
20120102914 | SYSTEMS, METHODS, AND APPARATUS FOR COMPENSATING FUEL COMPOSITION VARIATIONS IN A GAS TURBINE - Certain embodiments of the invention may include systems and methods for compensating fuel composition variations in a gas turbine. According to an example embodiment of the invention, a method is provided for compensating for fuel composition variations in a turbine. The method can include: monitoring at least one fuel parameter associated with a turbine combustor; monitoring one or more combustion dynamics characteristics associated with the turbine combustor; monitoring one or more performance and emissions characteristics associated with the turbine; estimating fuel composition based at least in part on the at least one fuel parameter, the one or more combustion dynamics characteristics, and the one or more performance and emissions characteristics, and adjusting at least one fuel parameter based at least in part on the estimated fuel composition. | 05-03-2012 |
20120102967 | Method and system for preventing combustion instabilities during transient operations - A method and system for preventing or reducing the risk of combustion instabilities in a gas turbine includes utilizing a turbine controller computer processor to compare predetermined and stored stable combustion characteristics, including rate of change of the characteristics, with actual operating combustion characteristics. If the actual operating combustion characteristics are divergent from stable combustion characteristics then the controller modifies one or more gas turbine operating parameters which most rapidly stabilize the operation of the gas turbine. | 05-03-2012 |
20120266604 | FUEL NOZZLE AND METHOD FOR OPERATING A COMBUSTOR - A fuel nozzle and a method for operating a combustor are disclosed. The method includes flowing a fuel and an oxidizer through a fuel nozzle, the fuel nozzle comprising an inner tube, an intermediate tube, and an outer tube each configured for flowing one of the fuel or the oxidizer therethrough. At least one of the inner tube, the intermediate tube, or the outer tube includes a plurality of swirler vanes. The method further includes imparting a swirl to the fuel and the oxidizer in the fuel nozzle, and exhausting the fuel and the oxidizer from the fuel nozzle into a combustion zone. | 10-25-2012 |
20130074504 | SYSTEM FOR INJECTING FUEL IN A GAS TURBINE ENGINE - A system including a controller. The controller may receive a signal indicative of an emissions level of a turbine. The controller may also generate at least one control signal. This control signal may control a split of diluent between a first and a second mixing chamber, whereby the first and the second mixing chambers are located in a fuel nozzle in the turbine. | 03-28-2013 |
20130098048 | DIFFUSION NOZZLES FOR LOW-OXYGEN FUEL NOZZLE ASSEMBLY AND METHOD - A fuel nozzle assembly has been conceived for a combustor in a gas turbine including a first passage and fourth passage connectable to a source of gaseous fuel, a second passage connectable to a source of a gaseous oxidizer, and a third passage coupled to a source of a diluent gas, wherein the first passage is a center passage and is configured to discharge gaseous fuel from nozzles at a discharge end of the center passage, the second passage is configured to discharge the gaseous oxidizer through nozzles adjacent to the nozzles for the center passage, the third passage discharges a diluent gas through nozzles adjacent to the nozzles for the second passage, and the fourth passage is configured to discharges the gaseous fuel downstream of the discharge location for the first, second and third passages. | 04-25-2013 |
20130263605 | Diffusion Combustor Fuel Nozzle - The present application thus provides a fuel nozzle for use with one or more flows of fuel and a flow of air in a combustor. The fuel nozzle may include one or more gas fuel passages for the one or more of flows of fuel, a swirler with one or more air chambers therein surrounding the gas fuel passages, and a collar with one or more curtain slots surrounding the swirler. The flow of air is divided between a swirler flow through the air chambers and a curtain flow through the curtain slots. | 10-10-2013 |
20130276450 | COMBUSTOR APPARATUS FOR STOICHIOMETRIC COMBUSTION - Gas turbine combustor with a specific fuel and oxidizer flow arrangement which provides high combustion efficiency for stoichiometric diffusion combustion in gas turbine applications operating with oxygen deficient working fluids. | 10-24-2013 |
20140090353 | 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. | 04-03-2014 |
20140260302 | DIFFUSION COMBUSTOR FUEL NOZZLE FOR LIMITING NOx EMISSIONS - The present application and the resultant patent provide a diffusion combustor fuel nozzle for a gas turbine engine. The fuel nozzle may include one or more gas fuel passages for one or more flows of gas fuel, a swirler surrounding the one or more gas fuel passages and positioned about a downstream face of the fuel nozzle, a number of swirler gas fuel ports defined in the swirler, and a number of downstream face gas fuel ports defined in the downstream face of the fuel nozzle. The swirler may include a number of swirl vanes and a number of air chambers defined between adjacent swirl vanes. The present application and the resultant patent further provide a method of operating a diffusion combustor fuel nozzle of a gas turbine engine. | 09-18-2014 |