Abstract: A mold assembly for use in forming a component having an internal passage defined therein is provided. The component is formed from a component material. The mold assembly includes a mold that defines a mold cavity therein. The mold assembly also includes a lattice structure selectively positioned at least partially within the mold cavity. The lattice structure is formed from a first material that has a selectively altered composition in at least one region of the lattice structure. A channel is defined through the lattice structure, and a core is positioned in the channel such that at least a portion of the core extends within the mold cavity and defines the internal passage when the component is formed in the mold assembly.

Abstract: A mold assembly for use in forming a component having an internal passage defined therein is provided. The component is formed from a component material. The mold assembly includes a mold that defines a mold cavity therein. The mold assembly also includes a lattice structure selectively positioned at least partially within the mold cavity. The lattice structure is formed from a first material that has a selectively altered composition in at least one region of the lattice structure. A channel is defined through the lattice structure, and a core is positioned in the channel such that at least a portion of the core extends within the mold cavity and defines the internal passage when the component is formed in the mold assembly.

Abstract: A method of recovering heat energy from a cooling medium used to cool hot gas path components in a turbine engine includes cooling one or more hot gas path components with the cooling medium; supplying spent cooling medium used to cool the one or more hot gas path components to a heat exchanger; supplying air (e.g., compressor discharge air) to the heat exchanger so as to be in heat exchange relationship with the spent cooling medium and thereby add heat to the compressor discharge air; and supplying the air heated in the heat exchanger to at least one combustor.

Abstract: A method of recovering heat energy from a cooling medium used to cool hot gas path components in a turbine engine includes cooling one or more hot gas path components with the cooling medium; supplying spent cooling medium used to cool the one or more hot gas path components to a heat exchanger; supplying air (e.g., compressor discharge air) to the heat exchanger so as to be in heat exchange relationship with the spent cooling medium and thereby add heat to the compressor discharge air; and supplying the air heated in the heat exchanger to at least one combustor.

Abstract: Certain embodiments of the invention may include systems, methods, and apparatus for detecting failure in gas turbine hardware. According to an example embodiment of the invention, a method for detecting a failure in a gas turbine is provided. The method can include monitoring a parameter associated with the turbine, wherein the monitored parameter comprises at least one turbine bucket temperature, detecting an event associated with operation of the turbine, wherein the event is based at least in part on the monitored parameter, and initiating shutdown of the turbine upon detection of the event wherein the monitored parameter is above a predetermined value for at least a predetermined time duration.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the flow sleeve and into the distribution manifold provides fluid communication through the flow sleeve to the plurality of fuel injectors.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the flow sleeve and into the distribution manifold provides fluid communication through the flow sleeve to the plurality of fuel injectors.

Abstract: A system for controlling oxygen emissions from a gas turbine includes a combustor, a plenum downstream from the combustor, and a turbine downstream from the plenum that produces an exhaust. A compressor downstream from the turbine receives the exhaust and produces a compressed exhaust. A compressed exhaust plenum downstream from the compressor provides fluid communication to the combustor and the plenum. A method for controlling oxygen emissions from a gas turbine includes flowing an exhaust from a turbine and increasing the pressure of the exhaust to produce a compressed exhaust. The method further includes flowing a first portion of the compressed exhaust to an inlet of a combustor and flowing a second portion of the compressed exhaust to a plenum between the turbine and the combustor.

Abstract: A combustor for a gas turbine includes a plurality of nozzles provided in an array; a baffle plate configured to provide a desired air flow distribution to the array of nozzles; and a casing comprising a plurality of holes in an outer surface. The casing extends from a headend of the combustor to the baffle plate. A method of distributing an air flow in a combustor of a gas turbine includes providing an air flow to the outer surface of the casing; directing the air flow around the baffle plate; and distributing the air flow through the baffle plate to the array of nozzles.

Abstract: A nozzle includes an inlet, an outlet, and an axial centerline. A shroud surrounding the axial centerline extends from the inlet to the outlet and defines a circumference. The circumference proximate the inlet is greater than the circumference at a first point downstream of the inlet, and the circumference at the first point downstream of the inlet is less than the circumference at a second point downstream of the first point. A method for supplying a fuel through a nozzle directs a first airflow along a first path and a second airflow along a second path separate from the first path. The method further includes injecting the fuel into at least one of the first path or the second path and accelerating at least one of the first airflow or the second airflow.

Abstract: A power generation system is provided and includes a compressor to increase a pressure of inlet ambient air and a turbine downstream from the compressor having a first stage and additional stages downstream from the first stage, the first stage structurally incorporating a combustor in which the compressed ambient air is mixed with fuel and in which the mixture is combusted to generate high temperature exhaust gas from which mechanical energy is derived in the first and additional stages.

Abstract: A power generation system is provided and includes a compressor to increase a pressure of inlet ambient air and a turbine downstream from the compressor having a first stage and additional stages downstream from the first stage, the first stage structurally incorporating a combustor in which the compressed ambient air is mixed with fuel and in which the mixture is combusted to generate high temperature exhaust gas from which mechanical energy is derived in the first and additional stages.

Abstract: A combustor cap is cooperable with an impingement plate in a turbine fuel nozzle and includes a plurality of cap segments independently securable to the impingement plate. The plurality of cap segments are radially and tangentially movable relative to the impingement plate.

Abstract: Certain embodiments of the invention may include systems, methods, and apparatus for detecting failure in gas turbine hardware. According to an example embodiment of the invention, a method for detecting a failure in a gas turbine is provided. The method can include monitoring a parameter associated with the turbine, wherein the monitored parameter comprises at least one turbine bucket temperature, detecting an event associated with operation of the turbine, wherein the event is based at least in part on the monitored parameter, and initiating shutdown of the turbine upon detection of the event wherein the monitored parameter is above a predetermined value for at least a predetermined time duration.

Abstract: Disclosed is a method and system for determining composition of a fuel entering a combustor. The method includes determining a temperature of the fuel entering the combustor, calculating a first estimated total fuel flow utilizing fuel properties and fuel nozzle effective area (Ae), and calculating a second estimated total fuel flow utilizing an aero-thermal cycle model analysis. The first estimated total fuel flow is compared to the second estimated total fuel flow and a lower heating value of the fuel is determined from a difference between the first estimated total fuel flow and the second estimated total fuel flow. A method and system for controlling a gas turbine includes calculating effects of the fuel composition on performance of the gas turbine and comparing one or more performance parameters to one or more parameter limits. One or more machine controls of the gas turbine are changed based on the results of the comparison.

Abstract: A combustor for a gas turbine includes a plurality of nozzles provided in an array; a baffle plate configured to provide a desired air flow distribution to the array of nozzles; and a casing comprising a plurality of holes in an outer surface. The casing extends from a headend of the combustor to the baffle plate. A method of distributing an air flow in a combustor of a gas turbine includes providing an air flow to the outer surface of the casing; directing the air flow around the baffle plate; and distributing the air flow through the baffle plate to the array of nozzles.

Abstract: A nozzle includes an inlet, an outlet, and an axial centerline. A shroud surrounding the axial centerline extends from the inlet to the outlet and defines a circumference. The circumference proximate the inlet is greater than the circumference at a first point downstream of the inlet, and the circumference at the first point downstream of the inlet is less than the circumference at a second point downstream of the first point. A method for supplying a fuel through a nozzle directs a first airflow along a first path and a second airflow along a second path separate from the first path. The method further includes injecting the fuel into at least one of the first path or the second path and accelerating at least one of the first airflow or the second airflow.

Abstract: A fuel/air premixer for use in a burner in a combustion system of an industrial gas turbine includes an air inlet, a fixed nozzle geometry, and an annular mixing passage. The fuel/air premixer mixes fuel and air in the annular mixing passage for injection into a combustor reaction zone. A plurality of fuel sources are connected with the fixed nozzle geometry, and each of the fuel sources is cooperable with the fixed nozzle geometry to effect multiple fuel flow variations including variations in fuel type, fuel blend, volumetric flow, and pressure ratios.

Abstract: A method facilitates assembling a gas turbine engine including a combustor assembly and a nozzle assembly. The method comprises providing a transition piece including a first end, a second end, and a body extending therebetween, where the body includes an inner surface, an opposite outer surface, coupling the first end of the transition piece to the combustor assembly, and coupling the second end of the transition piece to the nozzle assembly such that a turbulator extending helically over the outer surface of the transition piece extends from the transition piece first end to the transition piece second end to facilitate inducing turbulence to cooling air supplied to the combustor assembly.

Abstract: A combustor cap is cooperable with an impingement plate in a turbine fuel nozzle and includes a plurality of cap segments independently securable to the impingement plate. The plurality of cap segments are radially and tangentially movable relative to the impingement plate.