General Electric Company

Exxonmobil Upstream Research Company

Simons, Derrick Walter

Derrick Walter Simons

Jonathan Dwight Berry

Gregory Allen Boardman

Willy Steve Ziminsky

Douglas Edward Dean

Carey Edward Romoser

Bryan Wesley Romig

Timothy Andrew Healy

Minesh Ashok Shah

Krishna Kumar Venkataraman

Predrag Popovic

Ajai Singh

Abhijit Prabhakar Kulkarni

Arthur James Fossum

Michael John Hughes

Wei A Chen

Abdul Rafey Khan

William David York

Baifang Zuo

Jason Thurman Stewart

Jeffrey Scott LeBegue

Jong Ho Uhm

Jason Dean Fuller

Ilya Aleksandrovich Slobodyanskiy

Kevin Wood Wilkes

Chunyang Wu

Alberto Jose Negroni

Jonathan Hale Kegley

Igor Petrovich Sidko

Robert Wade Clifford

Daniel Doyle Vandale

Kara Johnston Edwards

Thomas Charles Amond, Iii

Ajay Kumar Gupta

Venkat Narra

Christopher T Meadows

Aja Kumar Gupta

Wei Chen

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.

Fuel staging in a gas turbine engine

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.

Premixing apparatus for gas turbine system

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

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.

Turbomachine and staged combustion system of a turbomachine

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.

Turbomachine with transition piece having dilution holes and fuel injection system coupled to transition piece

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.

Systems and methods for detecting fuel leaks in gas turbine engines

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.

Combustor probe for gas turbine

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.

Combustor recovery method and system

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.

Sector nozzle mounting systems

A method of assembling a turbine engine includes defining a first chamber and defining a second chamber. The method also includes forming at least one venturi device oriented with a predetermined venturi step angle greater than approximately 48&#xb0;. The method further includes coupling the first chamber in flow communication with the second chamber via the venturi device therebetween.

Method and apparatus for assembling turbine engines

A method of managing transient events regularly seen during gas turbine 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.

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

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.

Radial inlet guide vanes for a combustor

A gas turbine comprises a plurality of target exhaust temperature determination modules, the plurality of target exhaust temperature modules comprising a nitrogen oxide (NOx) compliance module configured to determine an exhaust temperature at which an exhaust of the gas turbine complies with a maximum permitted level of NOx; at least one bias module, the at least one bias module configured to apply a bias to an output of at least one of the plurality of target exhaust temperature determination modules; and a controller configured to operate the gas turbine to produce the exhaust temperature determined by the NOx compliance module.

NOX compliant peak for gas turbine

Disclosed is a mixing hole arrangement for improving homogeneity of an air and fuel mixture in a combustor, the mixing hole arrangement comprising a plurality of mixing holes defined by a liner, wherein at least one of the plurality of mixing holes is a mixing hole that is at least one of sized and positioned to impede penetration of a fluid flow into a primary mixing zone located in a head end of the combustor.

Mixing hole arrangement and method for improving homogeneity of an air and fuel mixture in a combustor

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.

Growing community of inventors

Greer, SC, United States of America

Derrick Walter Simons