Abstract: 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.

Abstract: A method for operating a gas turbine engine including a combustor assembly includes channeling a first fluid through a plurality of first nozzles into the combustor and igniting the first fluid downstream from the first nozzles. The method also includes increasing the operating speed of the engine and attaining a first predetermined percentage of a baseload by channeling the first fluid only through the first nozzles and then channeling a second fluid through a second nozzle into the combustor. The method also includes igniting the second fluid within the combustor downstream from the second nozzle. The method further includes channeling the second fluid to the first nozzles when the engine attains a second predetermined percentage of the baseload. The second predetermined percentage of the baseload is greater than the first predetermined percentage of the baseload. The method also includes terminating a flow of the first fluid through the first nozzles.

Abstract: 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.

Abstract: A system for a gas turbine engine includes a combustor assembly coupled to a plurality of first nozzles having outlet passageways coupled to a first fuel source and first passageways coupled to a second fuel source. The system also includes the combustor assembly coupled to a second nozzle that has second passageways coupled to the second fuel source. An associated control system, coupled to the combustor assembly, is configured to channel a first fuel from the first fuel source through the outlet passageways, and channel a second fuel from the second fuel source through the second passageways when the engine attains a predetermined load. The control system is further configured to channel the second fuel from the second fuel source through the first passageways and to reduce a flow of the first fuel entering the first nozzles when the engine is at a load greater than a first predetermined load.

Abstract: A system includes a gas turbine transition piece. The gas turbine transition piece includes a duct body having a forward end portion and an aft end portion. The duct body defines an enclosure for routing a flow of combustion products from a combustor to a turbine first stage nozzle, and the forward end portion includes a straight portion extending in a downstream direction of the flow. The gas turbine transition piece also includes a first set of dilution holes formed in the forward end portion and arranged in a first pattern configured to reduce emissions. In certain embodiments, the gas turbine transition piece includes a second set of dilution holes formed in the aft end portion and arranged in a second pattern configured to alter the thermal gradient of the combustion gases.

Abstract: A system for a gas turbine engine includes a combustor assembly coupled to a plurality of first nozzles having outlet passageways coupled to a first fuel source and first passageways coupled to a second fuel source. The system also includes the combustor assembly coupled to a second nozzle that has second passageways coupled to the second fuel source. An associated control system, coupled to the combustor assembly, is configured to channel a first fuel from the first fuel source through the outlet passageways, and channel a second fuel from the second fuel source through the second passageways when the engine attains a predetermined load. The control system is further configured to channel the second fuel from the second fuel source through the first passageways and to reduce a flow of the first fuel entering the first nozzles when the engine is at a load greater than a first predetermined load.

Abstract: A method for operating a gas turbine engine including a combustor assembly includes channeling a first fluid through a plurality of first nozzles into the combustor and igniting the first fluid downstream from the first nozzles. The method also includes increasing the operating speed of the engine and attaining a first predetermined percentage of a baseload by channeling the first fluid only through the first nozzles and then channeling a second fluid through a second nozzle into the combustor. The method also includes igniting the second fluid within the combustor downstream from the second nozzle. The method further includes channeling the second fluid to the first nozzles when the engine attains a second predetermined percentage of the baseload. The second predetermined percentage of the baseload is greater than the first predetermined percentage of the baseload. The method also includes terminating a flow of the first fluid through the first nozzles.

Abstract: A nozzle for channeling a fluid into a combustor assembly. The nozzle includes a body including a centerline, a first passageway extending through the body along the centerline, and a nozzle tip coupled to a downstream end of the first passageway. The nozzle tip has a radius extending from a center of the nozzle tip to an outer surface of the nozzle tip. The nozzle tip includes a downstream face, and a plurality of outlet passageways that each include an opening defined in the downstream face. Each opening includes respective X-, Y-, and Z-axes defined with respect to a tangent line, the radius, and the centerline. Each of the plurality of outlet passageways is at discharge angle of greater than 30° measured with respect to the respective Z-axis in a respective X-Z plane.

Abstract: A method for operating a gas turbine engine including a combustor assembly is provided. The method includes channeling a first fluid through a first nozzle into the combustor assembly, igniting the first fluid within the combustor assembly downstream from the first nozzle, channeling a second fluid through a second nozzle into the combustion assembly when the gas turbine engine attains a speed of greater than 85% of rated speed, igniting the second fluid within the combustor assembly downstream from the second nozzle, terminating a flow of the first fluid through the first nozzle, and channeling the second fluid to the first nozzle.

Abstract: 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.