Abstract: A combustor nozzle includes a plurality of mixing tubes through which air and fuel flow, a plurality of accommodation tubes each accommodating and supporting the plurality of mixing tubes therein, a first fuel tube coupled to each accommodation tube to supply a first fuel into the accommodation tube, a second fuel tube coupled to each accommodation tube to supply a second fuel into the accommodation tube, a dispersion chamber provided in each accommodation tube so as to be combined with the second fuel tube to receive the second fuel therein, a first fuel supply member supplying the first fuel received in each accommodation tube into each mixing tube, and a second fuel supply member supplying the second fuel received in the dispersion chamber into each mixing tube.

Abstract: A pilot fuel nozzle assembly includes a fuel nozzle, a swirler, and a vented pilot venturi. The vented pilot venturi has an annular wall with an oxidizer flow passage therein. An expansion flow surface portion of the venturi has a larger diameter at an outlet than at a throat of the venturi. A plurality of venturi oxidizer outlet ports extend through the expansion flow surface to the oxidizer flow passage within the annular wall to provide a flow of oxidizer through the venturi wall into a mixing cavity of the venturi and at an outlet end of the venturi. The oxidizer outlet ports are circumferentially spaced about a circumference of the expansion flow surface, and may be arranged in a plurality of rows. The oxidizer outlet ports may be angled with respect to the expansion flow surface and may angled circumferentially in a co-swirl direction with the swirler.

Abstract: A combustor liner for a gas turbine engine is disposed about an axis and defines a combustion chamber. The combustor liner includes a plurality of attachment elements circumferentially spaced about an outer wall of the combustor liner opposite the combustion chamber and a plurality of support elements connecting the plurality of attachment elements to combustor liner such that the plurality of attachment elements is radially spaced from the outer wall of the combustor liner. Each of the plurality of attachment elements includes an opening configured to slidingly receive a pin.

Abstract: A combustor for a gas turbine engine comprises: a head end section defining a head end plenum and containing a fuel nozzle assembly; and a liner extending downstream from the head end section to an aft frame and defining a combustion chamber therein. First injectors, disposed at a first axial location, direct a first fuel/air mixture through the liner. Second injectors, disposed at a different, second axial location, direct a second fuel/air mixture through the liner. Each of the head end section, the first injectors, and the second injectors receives a respective air supply from a compressor discharge plenum that at least partially surrounds the combustor. The respective air supplies are directed to only one of the fuel nozzle assembly, the first injectors, and the second injectors. The first injectors and the second injectors receive more than 50% of the air supply from the compressor discharge plenum.

Abstract: A combustor for a gas turbine engine includes a liner surrounding a fuel and air mixing body. A fuel supply passage communicates into an open fuel plenum downstream of the fuel supply passage. A wall of the mixing body has air openings to receive air flow, and communicate air into mixing passages. The mixing passages pass through the fuel plenum. Fuel openings in the mixing passages to allow fuel to flow from the mixing passage and mix with the air. There are passage sections extending downstream of the fuel plenum, such that the mixed air and fuel travels downstream of the fuel plenum and into a combustion chamber. A gas turbine engine is also disclosed.

Abstract: Proposed are a nozzle assembly, a combustor using the same, and a gas turbine. The nozzle assembly is configured to inject fuel and compressed air into a combustion chamber of a combustor of a gas turbine, and includes a hollow nozzle frame, a mixture supply tube having a plurality of tubes configured to supply a mixture of air and fuel to the combustion chamber, a fuel supply part configured to supply fuel to the mixture supply tube, an air supply part configured to supply air to the mixture supply tube, and a mixing part including an injection member formed in each of the tubes and configured to inject fuel supplied from the fuel supply part toward air passing through the tube, and a swirl vane formed on one side of the injection member and configured to mix fuel and air passing through the injection member.

Abstract: A gas turbine engine includes a bypass duct and a heat-exchanger assembly. The bypass duct is configured to direct air through a flow path. The heat-exchanger assembly is configured to receive a first portion of the air flowing through the flow path of the bypass duct and to divert a second portion of the air flowing through the flow path around the heat-exchanger assembly.

Abstract: An assembly for an aircraft powerplant includes a compressor rotor and an air system. The compressor rotor includes a plurality of compressor blades. The compressor blades are arranged circumferentially around the axis in a compressor blade array and are disposed in a flowpath. The air system includes an air source, an air injector and an air circuit fluidly coupling the air source to the air injector. The air system is configured to direct compressed air from the air source to the air injector through the air circuit. The air injector is configured to direct a stream of the compressed air into the flowpath and against one or more of the compressor blades to apply a rotational driving force onto the compressor rotor about an axis. The flowpath has a radial height at an injector location. The air injector projects a radial distance into the flowpath at the injector location.

Abstract: An exhaust nozzle for a gas turbine engine includes: an exhaust duct configured to receive an exhaust flow of gas from a combustor of the engine; a first flap rotatably coupled to the exhaust duct for rotation about a first axis; a first actuator configured to actuate the first flap about the first axis between a first inner and a first outer position; a second flap rotatably coupled to the exhaust duct for rotation about a second axis; and a second actuator configured to actuate the second flap about the second axis between a second inner and a second outer position. The first and second flaps at least in part define a passageway configured to convey the exhaust flow of gas to an exterior of the gas turbine engine. The first and second axes of rotation are coaxial. Also provides is a method of operating an exhaust nozzle.

Abstract: Combustors, gas turbines, and associated methods of operation are provided. A method for operating a combustor includes firing a bundled tube fuel nozzle assembly within a combustion liner of the combustor to generate combustion gases at a first temperature within a first combustion zone length. The method further includes firing a fuel injector downstream from the bundled tube fuel nozzle assembly within the combustion liner of the combustor to generate combustion gases at a second temperature within a second combustion zone length. The first combustion zone length is less than the second combustion zone length. The combustion gases travel through the first combustion zone length in a first time period and through the second combustion zone length in a second time period. The second time period is less than the first time period.

Abstract: A system includes a gas turbine having a turbine shaft disposed along a rotational axis, a turbine casing disposed circumferentially about the turbine shaft, a combustion gas path disposed between the turbine shaft and the turbine casing, a turbine stage disposed in the combustion gas path, wherein the turbine stage includes a plurality of turbine vanes disposed upstream from a plurality of turbine blades. The system includes an isothermal expansion system coupled to the turbine stage. The isothermal expansion system includes a plurality of fluid injectors configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion, wherein at least one fluid injector of the plurality of fluid injectors is coupled to each of the plurality of turbine vanes.

Abstract: A mobile fracking system and methods may include a gas turbine housed at least partially inside a trailer and an exhaust attenuation system configured to receive exhaust gas from the gas turbine. The exhaust attenuation system may include a lower elongated plenum configured to receive exhaust gas from the gas turbine and an upper noise attenuation system that is movably connected relative to a distal end of the lower elongated plenum.

Abstract: A combustor for a gas turbine engine has an annular inner liner and an annular outer liner forming a combustion chamber therebetween. A dilution horn pair includes dilution horns that provide a flow of an oxidizer gas into the combustion chamber in a dilution zone. At least one of the dilution horns forming the dilution horn pair is arranged so as to provide a lateral flow component of the flow of oxidizer gas therethrough into the combustion chamber, the lateral flow component having a flow direction extending laterally across and non-orthogonal to an axial flow direction of combustion gases within the combustion chamber.

Abstract: An aircraft dual-flow turbine engine assembly includes: an internal shroud for externally delimiting a primary flow path of the turbine engine gases; an external shroud for internally delimiting a secondary flow path of the turbine engine gas; and at least one air discharge duct extending between the internal shroud and the external shroud, the air discharge duct opening into the secondary flow path through an outlet orifice equipped with discharge fins. At least some of the discharge fins are movably mounted so as to be able to be incidence-control between a propulsion position, and a reverse thrust position.

Abstract: Methods and apparatus to operate a gas turbine engine with hydrogen gas are disclosed. An example combustor nozzle apparatus of a gas turbine engine includes injecting an other combustible gas into a combustor, comparing a power output of the gas turbine to a rated power threshold, and in response to the power output of the gas turbine satisfying the rated power threshold: injecting water into the combustor, injecting hydrogen into the combustor, and terminating injections of the other combustible gas.

Abstract: A sealing system for a fuel injector is provided. The fuel injector includes a body defining a plurality of fuel passages. Each of the plurality of fuel passages is defined by a first portion of the body and a second portion of the body. The sealing system includes a sealing structure adapted to be disposed between the first portion and the second portion. The sealing structure extends circumferentially around a central axis of the fuel injector. The sealing structure includes a ring and a plurality of lobes. Each lobe from the plurality of lobes is adapted to accommodate and seal a corresponding fuel passage from the plurality of fuel passages. The sealing system also includes a plurality of mechanical fasteners adapted to couple the first portion with the second portion, and to apply a pressure on the sealing structure to seal the plurality of fuel passages.

Abstract: A propulsion system comprises a propulsive hydrogen-burning gas turbine engine, a first tank storing liquid hydrogen with an ullage and a first fuel line including a fuel pump and a vaporiser, the first fuel line providing gaseous hydrogen to the engine during operation of the system. A second tank storing gaseous hydrogen is coupled by a second fuel line to the first fuel line at a position thereon between the vaporiser and the engine, providing for engine start-up (when the vaporiser is inoperative) and power-boosting during operation of the system. A duct connects gaseous hydrogen within the second tank to the ullage in the first tank in order maintain pressure in the first tank as liquid hydrogen within it is depleted, preventing cavitation of liquid hydrogen within the fuel pump.

Abstract: An assembly is provided for a gas turbine engine. This engine assembly includes a vane array and a heat exchanger integrated with the vane array. The vane array includes an inner platform, an outer platform and a plurality of vanes. The inner platform extends circumferentially about a centerline and forms an inner peripheral boundary of a flowpath through the vane array. The outer platform extends circumferentially about the centerline and forms an outer peripheral boundary of the flowpath through the vane array. The vanes extend across the flowpath between the inner platform and the outer platform. The vanes include a first vane and a second vane. The heat exchanger includes a first vane passage and a second vane passage fluidly coupled with and downstream of the first vane passage. The first vane passage extends through the first vane. The second vane passage extends through the second vane.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a flowpath duct, a fuel injector assembly and a secondary duct. The flowpath duct includes a flowpath. The fuel injector assembly includes an inner flow passage. The fuel injector assembly is configured to mix fuel with first gas within the inner flow passage to provide a fuel-gas mixture. The fuel injector assembly is configured to direct a jet of the fuel-gas mixture into the flowpath. The secondary duct is configured to direct second gas into the flowpath about the jet of the fuel-gas mixture.

Abstract: A diffuser for a combustor of a turbomachine engine. The diffuser includes a body having a first annular surface and a second annular surface that define a passage, the passage being configured to cause air to flow along the passage to a first component of the combustor. The first annular surface includes an opening that is configured to cause air to flow in a direction away from the passage to a second component of the combustor.