Abstract: A method for use in a gas turbine engine. The method includes the steps of: configuring a downstream injection system within the interior flowpath that includes two injection stages, a first stage and a second stage, wherein the first stage and the second stage are each axially spaced from the other; and circumferentially positioning the injectors of the first stage and the second stage based on: a) a characteristic of an anticipated combustion flow occurring just upstream of the first stage during a mode of operation; and b) the characteristic of an anticipated combustion flow just downstream of the second stage given an anticipated effect of the air and fuel injection from the first stage and the second stage.

Abstract: A gas turbine that includes: a combustor coupled to a turbine that together define an interior flowpath, the interior flowpath extending aftward about a longitudinal axis from a primary air and fuel injection system that defines a forward end, through an interface at which the combustor connects to the turbine, and through a row of stator blades in the turbine that defines an aft end; and a downstream injection system that includes two injection stages, a first stage and a second stage, that are axially spaced along the longitudinal axis of the interior flowpath. The first stage and the second stage each includes multiple injectors configured to inject an air and fuel mixture into the interior flowpath.

Abstract: A method of tuning a gas turbine includes receiving a first plurality of operating parameters as the gas turbine engine is operated at a first operating state. Further, the method includes operating the gas turbine engine at a second operating state to measure a second plurality of operating parameters at the second operating state. In addition, the method includes operating the gas turbine engine at a third operating state to measure a third plurality of operating parameters at the third operating state, wherein the first, second, and third operating states are different from each other. Additionally, the method includes generating a correction factor based on the first, second, and third plurality of operating parameters. The method also includes adjusting the operation of the gas turbine engine based on the correction factor.

Abstract: Optical flame holding and flashback detection systems and methods are provided. Exemplary embodiments include a combustor including a combustor housing defining a combustion chamber having combustion zones, flame detectors disposed on the combustor housing and in optical communication with the combustion chamber, wherein each of the flame detectors is configured to detect an optical property related to one or more of the combustion zones.

Abstract: Embodiments of methods and apparatus for providing compressor extraction cooling are provided. According to one example embodiment, a method is disclosed for controlling compressor extraction cooling. The method can include providing a cooling medium. The method can include extracting air from a compressor associated with a gas turbine. The method can also include introducing the cooling medium to the compressor extraction air, wherein the compressor extraction air is cooled by the cooling medium prior to or during introduction to the turbine section. Furthermore, method can include selectively controlling at least one of the compressor extraction air or the cooling medium based at least in part on a characteristic associated with the gas turbine.

Abstract: A gas turbine engine is provided and includes a late lean injection (LLI) compatible combustor having a first interior in which a first fuel supplied thereto by a fuel circuit is combustible, a turbine, a transition zone, including a second interior in which a second fuel supplied thereto by the fuel circuit and the products of the combustion of the first fuel are combustible, the transition zone being disposed to fluidly couple the combustor and the turbine to one another, and a plurality of fuel injectors, which are structurally supported by the transition zone and coupled to the fuel circuit, and which are configured to supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages.

Abstract: A gas turbine engine is provided and includes a fuel circuit, including multiple fuel circuit branches, a combustor having a first interior in which a first fuel supplied thereto by any one of the multiple fuel circuit branches is combustible, a turbine, a transition zone, including a second interior in which a second fuel supplied thereto by any one of the multiple fuel circuit branches, the second fuel including gas receivable by the fuel circuit from an external source, and the products of the combustion of the first fuel are combustible, the transition zone being disposed to fluidly couple the combustor and the turbine to one another, and a plurality of fuel injectors which supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages.

Abstract: A turbomachine includes a plurality of injection nozzles arranged in a can-annular array and a transition piece including at least one wall that defines a combustion flow passage. A dilution orifice is formed in the at least one wall of the transition piece. The dilution orifice guides dilution gases to the combustion flow passage. A heat shield member is mounted to the at least one wall of the transition piece in the combustion flow passage. The heat shield member includes a body having a first surface and an opposing second surface through which extends a dilution passage. The dilution passage is off-set from the dilution orifice. The heat shield member is spaced from the at least one wall of the transition piece defining a flow region between the at least one wall and the second surface.

Abstract: A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel supplied thereto by a fuel circuit is combustible, a turbine, a transition zone, including a second interior in which a second fuel supplied thereto by the fuel circuit and the products of the combustion of the first fuel are combustible, a plurality of fuel injectors, which are structurally supported by the transition zone and coupled to the fuel circuit, and which are configured to supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages, and a control system coupled to the fuel circuit and configured to control relative amounts of the first and second fuels supplied by the fuel circuit to the first and second interiors.

Abstract: A method for controlling a gas turbine combustion system includes supplying an ultra low calorific fuel to a combustor of the combustion system through a first fuel circuit, controlling a supply of the ultra low calorific fuel through a second fuel circuit as required to control the volumetric flow of the ultra low calorific fuel through the combustor, and combusting the ultra low calorific fuel in the combustor.

Abstract: A combustor assembly for a turbine engine includes a combustor liner and a flow sleeve which surrounds the combustor liner. Compressed air flows through an annular space located between an outer surface of the combustor liner and an inner surface of the flow sleeve. A plurality of cooling holes are formed through the flow sleeve to allow compressed air to flow from a position outside the flow sleeve, through the cooling holes, and into the annular space. The height of the annular space may vary along the length of the combustor assembly. Thus, the flow sleeve may have reduced diameter portions which result in the height of the annular space being smaller in certain locations than at other locations along the length of the combustor assembly.

Abstract: A turbine power generation system with enhanced cooling provided by a stream of carbon dioxide from a carbon dioxide source and a method of using a stream of carbon dioxide to cool hot gas path components. The turbine power generating system includes a compressor, a combustor, a turbine, a generator, and at least one shaft linking the compressor and turbine and generator together such that mechanical energy produced from the turbine is used to drive the compressor and the generator. Carbon dioxide that is sequestered from the exhaust of the turbine may be stored and injected back into the turbine to cool hot gas path components of the turbine.

Abstract: Certain embodiments of the invention may include systems and methods for providing optical interrogation sensors for combustion control. According to an example embodiment of the invention, a method for controlling combustion parameters associated with a gas turbine combustor is provided. The method can include providing an optical path through the gas turbine combustor, propagating light along the optical path, measuring absorption of the light within the gas turbine combustor, and controlling at least one of the combustion parameters based at least in part on the measured absorption.

Abstract: A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel is combustible, a turbine into which products of at least the combustion of the first fuel are receivable, a transition zone, including a second interior in which a second fuel and the products of the combustion of the first fuel are combustible, a plurality of fuel injectors which are configured to supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages, a compressor, by which air is supplied to the first and second interiors for the combustion therein, and a control system configured to control relative amounts of the air to the first and second interiors and relative amounts of the first and second fuels supplied to the first and second interiors.

Abstract: A combustor within which a combustion zone is defined is provided and includes an annular liner having a first mixing hole defined therein at a first axial position, a flow sleeve, having a second mixing hole defined therein at a second axial position, the flow sleeve surrounding the liner to form a first flow space at an exterior of the liner, a port, coupled to the flow sleeve at the second axial position, which is configured to remove air from the first flow space via the second mixing hole, and a shield, having a third mixing hole defined therein at the second axial position, the shield being disposed to shield the liner and to form a second flow space within the liner, which is communicable with the combustion zone via the third mixing hole and with the first flow space via the first mixing hole.

Abstract: A combustor for a gas turbine engine and related method is provided in which a plurality of combustor cans are selectively adapted with corresponding resonators. The resonators may, for example, be attached to every can in the consecutive arrangement of combustor cans, every other can, every third can or the like, and may be tuned to the same or first, second, third, etc. frequencies of operation. Such selective tuning is configured to suppress one or more of out-of-phase and in-phase dynamic interaction of streams discharged from adjacent combustor cans by changing the frequencies of pressure oscillation instabilities across the arrangement of consecutive cans.

Abstract: A nozzle is provided and includes an outer annulus defined by an exterior wall and an interior wall and including air inlets through which air flows to a fuel mixing zone and a combustion zone, an inner annulus disposed within the interior wall and including a fuel volume into which fuel is fed to a distal end thereof, which is adjacent to and isolated from the combustion zone and an airflow line, disposed between the fuel volume and the interior wall, through which air flows to the combustion zone with the airflow line and the combustion zone isolated from the fuel volume, and a fuse configured to melt during a flameholding incident and to form a breach through which fuel flows from the fuel volume, bypassing the fuel mixing zone, to a fuel burning zone downstream from the fuel mixing zone.

Abstract: A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel is combustible, a turbine into which products of at least the combustion of the first fuel are receivable, a transition zone, including a second interior in which a second fuel and the products of the combustion of the first fuel are combustible, a plurality of fuel injectors which are configured to supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages, a compressor, by which air is supplied to the first and second interiors for the combustion therein, and a control system configured to control relative amounts of the air to the first and second interiors and relative amounts of the first and second fuels supplied to the first and second interiors.

Abstract: A nozzle is provided and includes an outer annulus defined by an exterior wall and an interior wall and including air inlets through which air flows to a fuel mixing zone and a combustion zone, an inner annulus disposed within the interior wall and including a fuel volume into which fuel is fed to a distal end thereof, which is adjacent to and isolated from the combustion zone and an airflow line, disposed between the fuel volume and the interior wall, through which air flows to the combustion zone with the airflow line and the combustion zone isolated from the fuel volume, and a fuse configured to melt during a flameholding incident and to form a breach through which fuel flows from the fuel volume, bypassing the fuel mixing zone, to a fuel burning zone downstream from the fuel mixing zone.

Abstract: A fuel nozzle assembly for a combustor of a gas turbine including: a nozzle body having a front and an inner tube defining a fuel passage extending through the nozzle body, wherein the front proximate to a combustion section of the combustor; an outer casing around the inner tube, wherein an air passage is defined between the outer casing and the inner tube; a gas conduit arranged in the air passage and having an outlet proximate to the front of the nozzle body, wherein fuel starts flowing through the expandable conduit only after a flashback condition occurs in the combustor, and a premix fuel passage and port discharging fuel to a premix section of the combustor, wherein the gas conduit has an inlet open to the premix fuel passage.