Abstract: A cryogenic fuel supply system includes a storage tank, a mixing chamber, an auxiliary heating device, a heat exchanger, and flow distribution devices, and a controller. The storage tank stores cryogenic fuel in a liquid state. The mixing chamber receives various flows of cryogenic fuel in a supercritical or gaseous state, the mixing chamber being connected to a combustion chamber to supply the combustion chamber with cryogenic fuel in the supercritical or gaseous state. The auxiliary heating device heats the cryogenic fuel. The heat exchanger assembly includes a cryogenic fuel/oil heat exchanger and a heat exchanger between the cryogenic fuel and the air circulating in a primary duct of the turbine engine. A flow distribution device is upstream of the auxiliary heating device, and one or more flow distribution devices are disposed upstream of the heat exchanger assembly. The controller controls opening and closing of the flow distribution devices.

Abstract: An apparatus is provided for a turbine engine. This apparatus includes a fuel-air mixer, and the fuel-air mixer includes an inner passage, a sidewall, a steam passage, a fuel nozzle and an air swirler. The inner passage extends axially along an axis within the fuel-air mixer. The sidewall extends circumferentially around and axially along the inner passage. The steam passage is embedded within the sidewall and extends along the inner passage. The fuel nozzle is configured to direct fuel into the inner passage. The air swirler is configured to direct swirled air into the inner passage for mixing with the fuel.

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.

Abstract: A method for controlling a gas turbine, having a measurement step, a prediction step which is carried out after the measurement step, and a control step which is carried out after the prediction step. In the measurement step, a state variable of a combustion within a gas turbine is measured. In the prediction step, a future combustion dynamic is predicted using the measured state variable. In the control step, a control signal is output using the prediction of the future combustion dynamic.

Abstract: A fuel system of an aircraft engine, having: a boost pump having an input and an output; one or more selector valves; a first component pump having an input fluidly coupled to the output of the boost pump and an output of the first component pump is configured to direct fuel to a first component via the one or more selector valves; and a second component pump having an input that is selectively coupled to either the input or the output of the boost pump by the one or more selector valves, and an output of the second component pump is fluidly coupled to a second component and selectively coupled to the first component by the one or more selector valves.

Abstract: The present disclosure is directed to a system comprising a recuperated gas turbine engine with a catalytic combustor, and methods of operating the same, the catalytic combustor comprising: (a) an upstream section comprising an electrical heater and (b) a downstream catalyst section, wherein the upstream section and the downstream catalyst section are disposed adjacent to and in fluid communication with one another.

Abstract: A system and method of igniting a coal air-fuel mixture, including a burner having a burner tube operable to carry a flowing mixture of fuel and air to a furnace for combustion therein and a first flow directing device disposed within the tube, operable to direct a first portion of the flowing fuel and air mixture to a location in the burner tube. The system also includes a laser igniter within the burner tube, the laser igniter including a laser tube having a first end with a laser light input and a second end with a light output, and a laser light source operably coupled to the laser light input. The laser light source, including a laser. The laser ignitor directing photons from the light output at the location in the burner tube to ignite at least a part of the first portion of the fuel.

Abstract: A power generation installation in which exhaust gas from a gas turbine is fed to a thermal energy accumulator, wherein energy in the thermal energy accumulator can be employed for various purposes, a method for operating such an installation, and a method for the modification of existing installations. The thermal energy accumulator has sufficient capacity to permit the operation of a steam turbine in isolation for the storage of thermal energy from exhaust gas in the thermal energy accumulator.

Abstract: Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to the combustor by supplying fuel to the first fuel manifold via a first flow divider valve and supplying fuel to the second fuel manifold via a second flow divider valve. While supplying fuel to the combustor by supplying fuel to the first fuel manifold, the method includes stopping supplying fuel to the second fuel manifold and supplying pressurized gas to the second fuel manifold via the second flow divider valve to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated nozzles.

Abstract: A combustor includes a combustor liner defining a combustion chamber and a fuel and air mixing body connected to the combustor liner to deliver mixed fuel and air into the combustion chamber. The mixing body includes an inner housing member centered on a center axis and an intermediate housing member. A mixing passage is defined between the inner and intermediate housing members. The mixing passage extends along a direction from an upstream end to a downstream end with a circumferential component, a component in an axially downstream direction, and a radially inward component with at least one air inlet into the mixing passage. A fuel supply extends into the mixing passage at a location downstream of the air inlet. The mixing passage extends downstream to supply fuel and air into the combustion chamber. A gas turbine engine is also disclosed.

Abstract: An aircraft including a combustion engine with a combustion chamber an injection device for injecting fuel in the combustion chamber and an air separation device adapted to separate air into an oxygen-enriched gas mix and an nitrogen-enriched gas mix, wherein the oxygen-enriched gas mix is injected into the combustion chamber with the fuel while the nitrogen-enriched gas mix is used to inert at least some portions of the aircraft in the environment of said combustion engine.

Abstract: An embodiment of a torch ignitor system for combustor of a gas turbine engine includes a torch ignitor, the torch ignitor having a combustion chamber oriented about an axis, the combustion chamber having axially upstream and downstream ends defining a flow direction through the combustion chamber, along the axis. The torch ignitor system also includes a cap defining the axially upstream end of the combustion chamber and oriented about the axis, wherein the cap is configured to receive a fuel injector and at least one glow plug, a tip at a downstream end of the combustion chamber, and a passage for pressurized oxygen containing gas passing through the cap from an exterior of the combustion chamber and in fluid communication with the combustion chamber. An embodiment of a method for starting a gas turbine engine is also disclosed.

Abstract: The present application discloses a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method comprises: determining, during use of the gas turbine engine, one or more exhaust content parameters by performing a sensor measurement on an exhaust of the gas turbine engine; and determining one or more fuel characteristics of the fuel based on the one or more exhaust parameters. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system adjusts parameters of the fuel based on data received from the sensors.

Abstract: A turbine engine having a combustion section. The combustion section can have a combustor, at least one fuel cup, and at least one ignition tube. The combustor can include a combustor liner and a dome wall, which together at least partially define a combustion chamber. The at least one fuel cup can include a swirler and a fuel injector. The at least one ignition tube can include an outlet directly fluidly coupled to the combustion chamber.

Abstract: Provided is a combustion nozzle including a center cylinder having a cylinder shape, a middle cylinder coaxially disposed with the center cylinder and surrounding the center cylinder, an outer cylinder coaxially disposed with the middle cylinder and surrounding the middle cylinder, and a perforated plate disposed inside the center cylinder to create turbulent flow therethrough. In addition, the combustion nozzle further includes inner vanes disposed between the center cylinder and the middle cylinder and outer vanes disposed between the middle cylinder and the outer cylinder.

Abstract: A mixing assembly for a combustor includes: a pilot mixer including a pilot housing extending along a mixer centerline and a pilot fuel nozzle; a main mixer surrounding the pilot mixer; a fuel manifold between the pilot and main mixers; a mixer foot extending from a main housing of the main mixer; a main swirler body surrounding the main housing defining a mixing channel between the main housing and the main swirler body; and a main fuel ring in the mixing channel connected to the main housing by main fuel vanes, at least one of the main fuel ring and main fuel vanes including fuel injection ports for discharging fuel into the mixing channel, wherein the fuel injection ports are disposed non-uniformly relative to the mixer centerline, so as to produce a static pressure difference therebetween in response to mixer air flow passing around the main fuel ring.

Abstract: A combustion staging system for fuel injectors of a multi-stage combustor of a gas turbine engine. The system includes plural fuel injectors, each having respective pilot and mains injection stages. It further includes a splitting unit which, to perform staging control of the combustor, receives a metered fuel flow and, for pilot and mains operation, controllably splits the received fuel flow into a pilot flow for injecting at the pilot stages of the injectors and a mains flow for injecting at the mains stages of the injectors, and for pilot-only operation, controllably splits the received fuel flow into a first part of the pilot flow for injecting at the pilot stages of a first portion of the injectors and a second part of the pilot flow for injecting at the pilot stages of a second portion of the injectors.

Abstract: A combined combustor and recuperator is formed with the recuperator surrounding the combustor. Cold gas conduits (14, 16, 20) through the recuperator follow along involute paths toward the combustor. Hot has conduits (26) through the recuperator follow counterflow paths along corresponding involute curves outward from the combustor. The openings (18) in the combustion chamber wall through which cold gas enters the combustion chamber may be directed to impart flow direct to the cold gas to support particular desired behaviour of the cold gas in the portions of the combustion chamber concerned, e.g. supporting a stable vortex flame, enhancing mixing, providing a protective barrier layer.

Abstract: A fuel injector includes a forward end wall and an aft end wall. The fuel injector further includes side walls that extend between the forward end wall and the aft end wall. The forward end wall, the aft end wall, and the side walls collectively define an opening for passage of air. At least one fuel injection member is disposed within the opening and extends between the end walls. A fuel circuit is defined within the fuel injector. The fuel circuit includes an inlet plenum defined within the forward end wall of the fuel injector. The fuel circuit further includes a fuel passage that extends from, and is in fluid communication with, the inlet plenum. The fuel passage is defined within the at least one fuel injection member. The fuel passage has a cross-sectional area that varies along a length of the fuel injection member.

Abstract: A method of starting a gas turbine and related apparatus, the method including keeping a main fuel line in a sealed state while allowing fuel to flow from a fuel source to at least one nozzle of a nozzle array via an auxiliary fuel line; firing a gas turbine while keeping the main fuel line in the sealed state; and after combustion has started: opening the main fuel line to commence a flow of fuel from the fuel source to the at least one nozzle of the nozzle array via the main fuel line, sealing the auxiliary fuel line, and after the auxiliary fuel line is sealed, increasing the flow of fuel from the fuel source to the at least one nozzle of the nozzle array via the main fuel line.

Abstract: An aircraft gas turbine engine system comprises first and second gas turbine engines. The first gas turbine engine has first and second spools. A first power linkage connects the second gas turbine engine to the first spool of the gas turbine engine, and a second power linkage connects the second gas turbine engine to the second spool of the first gas turbine engine.

Abstract: Various embodiments include an apparatus for ascertaining a heating temperature of a heating element for an electrically heatable catalytic converter comprising: a catalytic converter housing surrounding the heating element and the catalytic converter; a first temperature sensor arranged in the housing; and a second temperature sensor arranged in the housing downstream of the first temperature sensor with regard to an exhaust gas flow direction within the housing. The first temperature sensor is exposed to radiation from the heating element and the second temperature sensor is shielded from radiation from the heating element.

Abstract: The gas turbine engine can have a structure for holding bearings within a casing, with a shaft being rotatably mounted to the casing via the bearings and via the structure, the structure having a first wall segment and a base structure receiving the bearings, the first wall segment having a proximal end structurally joined to the base structure, the first wall segment extending away from the base structure, and having a portion extending at least partially axially and thereby being radially flexible relative to the second wall segment, the structure providing both structural resistance and radial stretchability.

Abstract: A through-flow gas turbine engine includes a core comprising multiple spools rotatable about a center axis. An accessory gearbox (AGB) is drivingly engaged to the core and disposed aft of the outlet. A reduction gearbox (RGB) is drivingly engaged to the core and disposed forward of the inlet. The RGB has an RGB output to provide rotational output to a rotatable load. An electric motor is drivingly engaged to the rotatable load. An electric generator is configured to provide electrical power to the electric motor. One of the electric motor and the electric generator is disposed axially between the outlet and the AGB and the other of the electric motor and the electric generator is disposed axially between the inlet and the RGB.

Abstract: A control device is a control device for a gas turbine including a plurality of combustors and is configured to select combustors to ignite in accordance with a target load on the basis of a predictor which defines a relationship between a load and the number and arrangement of combustors to ignite and a combustion temperature.

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 a first circuit to transport a blend of hydrogen gas, inert gas, and/or other combustible gas from a supply to a gas turbine combustor, the blend of hydrogen gas, inert gas, and/or other combustible gas including between 100% hydrogen gas, 100% inert gas, or 100% other combustible gas, a second circuit to transport water from the supply to the gas turbine combustor, and a nozzle tip. The nozzle tip includes a first outlet in connection with the second circuit to provide the water to the gas turbine combustor, and a second outlet in connection with the first circuit, the second outlet concentrically positioned within the first outlet to provide the blend of hydrogen gas, inert gas, and/or other combustible gas to the gas turbine combustor.

Abstract: The present invention relates to an arrangement for cooling of an electrical machine. The electrical machine comprises a rotor rotatably arranged around a rotation axis, a stator including a stator winding arranged radially outside of the rotor, a housing enclosing the rotor and the stator, at least one drain hole configured to drain a cooling fluid from the housing, at least one spraying device configured to spray cooling fluid on the stator winding and a pump configured to pump cooling fluid to the spraying device. The arrangement comprises a control unit, which is configured, when the electrical machine is in operation, to receive information of the temperature in at least one position of the stator winding and to control the pump such that it pumps a flow rate of the cooling fluid to the spraying device, which is related to the estimated temperature of the stator winding.

Abstract: The gas turbine comprises a compressor, a combustor, and a turbine. The method comprises: compressing air with the compressor and feeding compressed air continuously to the combustor, feeding fuel to the combustor, continuously firing the mixture of fuel and gas in the combustor, feeding combustion gases from the combustor to the turbine, and supplying at least a portion of the total amount of fuel that is supplied to the combustor intermittently.

Abstract: Pre-ground plastics of small particle size not more than 2 mm are co-fed into a solid fossil fuel fed entrained flow partial oxidation gasifier. A syngas composition can be made by charging an oxidant and a feedstock composition comprising recycle plastics and a solid fossil fuel to a gasification zone within a gasifier; gasifying the feedstock composition together with the oxidant in said gasification zone to produce said syngas composition; and discharging at least a portion of said syngas composition from said gasifier; wherein the recycled plastics are added to a feed point comprising a solid fossil fuel belt feeding a grinder after the solid fossil fuel is loaded on the belt, a solid fossil fuel belt feeding a grinder before the solid fossil fuel is loaded onto the belt, or a solid fossil fuel slurry storage tank containing a slurry of said solid fossil fuel ground to a size as the size fed to the gasification zone.

Abstract: A system and method for providing bleed air compensation for a continuous power assurance analysis of a gas turbine engine includes estimating bleed air flow rate from the gas turbine engine, estimating a shift in power turbine inlet temperature based on the estimated bleed air flow rate, and applying the estimated shift in power turbine inlet temperature to the continuous power assurance analysis of the gas turbine engine.

Abstract: A fuel nozzle assembly and a gas turbine combustor including the same are provided. The fuel nozzle assembly may include an end plate coupled to one end of an annular casing, and a fuel nozzle configured such that one end thereof is supported by the end plate and the other end thereof extends outward. The fuel nozzle may include a center fuel nozzle and a plurality of side fuel nozzles arranged annularly to surround the center fuel nozzle. The side fuel nozzle may include a nozzle body located at a center thereof, a shroud spaced outward from the nozzle body, and a plurality of swirlers located between the nozzle body and the shroud. Each of the swirlers may include a leading edge directed toward the end plate and a trailing edge located opposite the leading edge. In each of the side fuel nozzles, distances between the leading edges are different from each other.

Abstract: A gas turbine engine has a circumferential staging configuration of fuel injectors in a combustor. As a turbine blade revolves within the gas turbine engine it is subjected to lift and drag forces based on the configuration of lit injectors. A configuration of lit injectors that results in the minimum unsteady forces the turbine blade experiences is determined in order to increase the life span of the turbine blade and limit any structural failures.

Abstract: A portable fuel preservation system is disclosed. The portable fuel preservation system may comprise a switch box configured to be coupled to an integrated fuel pump and control of a gas turbine engine. The switch box may comprise a circuit configured to cause a metering valve and a solenoid valve of the integrated fuel pump and control to open. A driver may be configured to inject preservation fluid into the integrated fuel pump and control.

Abstract: Provided is a combustion nozzle including a center cylinder having a cylinder shape, a middle cylinder coaxially disposed with the center cylinder and surrounding the center cylinder, an outer cylinder coaxially disposed with the middle cylinder and surrounding the middle cylinder, and a perforated plate disposed inside the center cylinder to create turbulent flow therethrough. In addition, the combustion nozzle further includes inner vanes disposed between the center cylinder and the middle cylinder and outer vanes disposed between the middle cylinder and the outer cylinder.

Abstract: In examples, a propulsion and electrical generation system including a gas turbine engine including a compressor and a bleed air outlet from the compressor, wherein the compressor is configured to compress a fluid, wherein a portion of the compressed fluid is directed out of the bleed air outlet to define bleed air from the compressor; and a turbo-generator including a combustor, wherein the combustor includes a fuel inlet and a bleed air inlet, wherein the bleed air inlet is in fluid communication with the bleed air outlet from the compressor, wherein the combustor is configured to receive the bleed air via the bleed air inlet from the bleed air outlet of the gas turbine engine and receive fuel via the fuel inlet, wherein the combustor is configured to combust the received fuel with the received bleed air, wherein the turbo-generator is configured to generate electrical energy via the combustion of the fuel by the combustor.

Abstract: A cab is used in a work machine. The cab includes a first pillar, a second pillar disposed to a rear of the first pillar, a beam disposed between the first pillar and the second pillar, and a reinforcing member provided to the beam.

Abstract: The present invention relates to an arrangement for cooling of an electrical machine. The electrical machine comprises a rotor rotatably arranged around a rotation axis, a stator including a stator winding arranged radially outside of the rotor, a housing enclosing the rotor and the stator, at least one drain hole configured to drain a cooling fluid from the housing, at least one spraying device configured to spray cooling fluid on the stator winding and a pump configured to pump cooling fluid to the spraying device. The arrangement comprises a control unit, which is configured, when the electrical machine is in operation, to receive information of the temperature in at least one position of the stator winding and to control the pump such that it pumps a flow rate of the cooling fluid to the spraying device, which is related to the estimated temperature of the stator winding.

Abstract: A gas turbine engine includes devices, systems, and methods for design and operation of stabilizing heat exchangers for gas turbine engines in consideration of near supercritical fuel flow conditions.

Abstract: The present invention generally relates to a sequential combustor for a gas turbine having second and/or subsequent stages of a re-heat, sequential or axially-staged combustion system. A variation in Mach number along the flow path can be used to control static temperature variation, which in turn influences the progress of auto-ignition reactions that eventually lead to the onset of combustion.

Abstract: A method and system for flue gas reclamation is described. In one embodiment, a flue gas reclamation system is provided. The system includes a combustion engine including an intake member, an output shaft, and an exhaust outlet. The intake member receives flue gas from a gas source. A generator is connected to the output shaft and a compressor is connected to the exhaust outlet of the combustion engine. At least one holding tank is connected to the compressor and the compressor stores enriched flue gas from the exhaust outlet of the combustion engine in the at least one holding tank. A battery is connected to the generator and is configured to provide electric power to the flue gas reclamation system. An algae farm in fluid communication with the at least one holding tank is configured to receive the stored enriched flue gas from the at least one holding tank.

Abstract: A fuel system for a gas turbine engine includes a plurality of duplex nozzles arranged on each side of top dead center and a plurality of simplex nozzles. A primary manifold is operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles and a secondary manifold is operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. An equalizer valve that is in communication with both the primary manifold and the secondary manifold distributes fuel at various pressures to both the primary and secondary manifolds.

Abstract: An apparatus and method for a reverse flow combustor, the reverse flow combustor including a straight portion, a dilution portion and a curved portion. The reverse flow combustor receives a flow of fuel that is ignited and mixed with cooling air to form a flow of combustion gases. The flow of combustion gases travels through the reverse flow combustor to a turbine section of an engine.

Abstract: A fuel-based thermal management system may comprise a heat exchanger configured to thermally couple a fluid and a fuel. A controller may be configured to modulate a flow of the fluid to the heat exchanger. A tangible, non-transitory memory may be configured to communicate with the controller. The controller may determine a temperature of the fluid, estimate a dissolved oxygen concentration in the fuel using a first fuel temperature, a flight cycle time, and at least one of an altitude measurement or a ambient pressure measurement, and modulate the flow of the fluid to the heat exchanger based on the dissolved oxygen concentration.

Abstract: A gas turbine system has a source of ammonia and a source of an oxygen-containing gas, a first combustion chamber connected to receive ammonia, a hydrogen-rich gas stream and oxygen-containing gas, a turbine connected to receive an exhaust gas stream from the first combustion chamber; and a second combustion chamber connected to receive an exhaust gas from the turbine, ammonia and a hydrogen-rich gas stream.

Abstract: A fuel nozzle apparatus for a gas turbine engine includes: an annular outer body extending parallel to a centerline axis, with an exterior surface having a plurality of openings; a main injection ring disposed inside the outer body and including: a circumferential main fuel gallery; and a plurality of main fuel orifices, each main fuel orifice communicating with the main fuel gallery and aligned with one of the openings; a venturi disposed inside the main injection ring; an annular splitter disposed inside the venturi, and including inner and outer walls spaced-apart from each other defining a splitter cavity, and a plurality of discharge holes communicating with the splitter cavity; an array of outer swirl vanes extending between the venturi and the splitter; a pilot fuel injector within the splitter; and an array of inner swirl vanes extending between the splitter and the pilot fuel injector.

Abstract: A burner assembly with a combustion chamber, a plurality of mixing ducts leading into the combustion chamber, where combustion air and introduced fuel are mixed. The mixing ducts are formed by mixing tubes extending axially through an annular space between a tubular outer wall, a tubular inner wall arranged radially at a distance from the outer wall, a ring-shaped end plate arranged upstream and a ring-shaped end plate arranged downstream. The end plates have through-openings, which accommodate and/or extend the mixing tubes, and the end plates have, both radially inside and outside, a circumferential edge extending in the direction of the annular space, with axial bores in the edge of the ring-shaped end plate arranged downstream. The axial bores extend from the annual space into the end plate, and at least one opening is for removing cooling air, the opening branching from the axial bore.

Abstract: A fuel injector device includes a body having a leading edge and a trailing edge and defining a streamwise direction from the leading edge to the trailing edge. The fuel injector device body includes a first surface and a second surface opposite the first surface, each surface extending between and including the leading edge and the trailing edge, and the surfaces conjoining each other at the leading edge and the trailing edge. The trailing edge, when seen in the streamwise direction, undulates along a trailing edge mean line and, along its extent, deviates in opposite directions from the mean line and includes at least one inflection point along its extent.

Abstract: A fuel injector is provided for the radial introduction of a liquid fuel/air mixture to a combustor. The fuel injector includes a body having a frame that defines an inlet portion and an outlet member that defines an outlet portion. A fuel plenum is defined within the outlet member, and a fuel injection port, which communicates with the fuel plenum, is defined through the outlet member. A fuel supply conduit, fixed to the body, communicates between a source of liquid fuel and the fuel injection port, via the fuel plenum. Alternately, the fuel injector may include a swirl-inducing device mounted to the outlet member in communication with the fuel injection port, and a fuel supply conduit fixed to the swirl-inducing device. In this embodiment, the fuel supply conduit communicates between the fuel injection port and a source of a liquid fuel and water mixture, via the swirl-inducing device.

Abstract: A fuel nozzle for gas turbine combustor of the present invention includes a diffusion combustion burner including a diffusion fuel nozzle, a premixing combustion burner including a premixing fuel nozzle, an end flange to which the diffusion combustion burner and the premixing combustion burner are mechanically joined, and a diffusion combustion burner support which fixes the diffusion combustion burner to the end flange and is mounted around the premixing fuel nozzle in a manner to cover a root portion of the premixing fuel nozzle, the root portion being a junction between the premixing fuel nozzle and the end flange. An outside diameter of the premixing fuel nozzle longitudinally varies in a manner to progressively increase toward the junction. A maximum diameter of the root portion of the premixing fuel nozzle is greater than an inside diameter of a hole formed in the diffusion combustion burner support.