Abstract: A combustor for a gas turbine engine is disclosed which is able to operate with high combustion efficiency, and low nitrous oxide emissions during gas turbine operations. The combustor consists of a can-type configuration which combusts fuel premixed with air and delivers the hot gases to a turbine. Fuel is premixed with air through a swirler and is delivered to the combustor with a high degree of swirl motion about a central axis. This swirling mixture of reactants is conveyed downstream through a flow path that expands; the mixture reacts, and establishes an upstream central recirculation flow along the central axis. A cooling assembly is located on the swirler co-linear with the central axis in which cooler air is conveyed into the prechamber between the recirculation flow and the swirler surface.

Abstract: A fuel distribution apparatus has a manifold head, a waisted portion and a nozzle head. An inlet channel is formed in a mounting surface of a casing of a jet engine. An annular inlet manifold is formed in the manifold head to receive fluid from the fuel inlet channel. Six outlet fluid passageways extend substantially tangentially away from the annular inlet manifold and open out into an annular outlet manifold. Three ports are formed in the inner wall of the annular outlet manifold and the ports are connected to fuel delivery tubes which extend away from the ports in a substantially helical fashion around the waisted portion into a nozzle head. The outlet fluid passageways are arranged to direct the fluid around the outlet manifold in a first direction and the helical fuel delivery tubes extend in the opposite direction.

Abstract: A reductant delivery unit (16) is provided unit for selective catalytic reduction (SCR) after-treatment for vehicles. The unit includes a solenoid fluid injector (18) associated with an exhaust gas flow path (13). The injector has a fluid inlet (28) and a fluid outlet (30) with the fluid inlet receiving a source of urea solution and the fluid outlet communicating directly with the exhaust flow path so as to control injection of urea solution into the exhaust gas flow path. Supply structure (32) defines the fluid inlet and includes a cup (37) coupled to a body (18) of the injector and a supply tube (29) integral with the cup to define a single member. The supply tube is coupled with the source (27) of urea solution to deliver urea solution to the fluid inlet. The supply tube is heated by a heat source (33) so that an entire volume of the urea solution delivered to the fluid inlet is heated.

Abstract: A burner assembly for a firing system for firing fluidic fuels is provided. The burner assembly has a burner hub, an air inlet channel and a fuel inlet channel, wherein the fuel inlet channel is at least partially designed in the burner hub. A screening wall is arranged in the fuel inlet channel. The screening wall is spaced apart from a wall of the fuel inlet channel so that an intermediate space is formed between the wall of the fuel inlet channel and the screening wall. A sleeve is equipped with at least one radial positioning means that ensures a clearance of the sleeve from the wall of the fuel inlet channel. The at least one radial positioning means of the sleeve is designed as a positioning projection arranged in a circular manner, protruding out radially.

Abstract: A gas turbine engine is piloted with a pilot flow of fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A second additional flow of the fuel is vaporized and delivered to the combustor as a vapor. A fuel injector may have passageways associated with each of the three flows.

Abstract: The structure and operation of a fuel nozzle for a gas turbine combustor is disclosed where the fuel nozzle provides for flameholding protection and, more specifically, to such a nozzle that provides for nondestructive protection from flameholding. The nozzle provides for differential thermal expansion between tubes forming fuel passages to allowing for the nondestructive venting of fuel during a flameholding condition. Upon extinguishing the flameholding condition, the nozzle returns to normal operating condition.

Abstract: A multiple conduit system for a gas turbine engine, the multiple conduit system extending between a plurality of conduit inlet and outlets. A channel, adapted for conveying fuel flow, is formed in a surface of a gas turbine engine component. The channel includes at least a first discrete conduit and a second discrete conduit. The first and second discrete conduits are each adapted to direct an independent fluid flow from respective inlets to respective outlets.

Abstract: A burner (1) for a combustion chamber of a turbogroup includes a swirl generator (2), a mixer (3), and a lance (4) for introducing pilot fuel into a combustion space (10). In order to stabilize combustion, the lance (4) is designed and/or arranged so that, at least in the pilot mode of the burner (1), it extends far enough into the burner interior (5) for a flame front (16) of a combustion reaction, which takes place in the combustion space (10), to extend at least partially into the burner interior (5).

Abstract: A fuel injector for use in a gas turbine engine combustor assembly. The fuel injector includes a main body and a fuel supply structure. The main body has an inlet end and an outlet end and defines a longitudinal axis extending between the outlet and inlet ends. The main body comprises a plurality of air/fuel passages extending therethrough, each air/fuel passage including an inlet that receives air from a source of air and an outlet. The fuel supply structure communicates with and supplies fuel to the air/fuel passages for providing an air/fuel mixture within each air/fuel passage. The air/fuel mixtures exit the main body through respective air/fuel passage outlets.

Abstract: A turbomachine includes a compressor, a turbine operatively coupled to the compressor, and a combustor fluidly linking the compressor and the turbine. The combustor includes at least one fuel nozzle. The at least one fuel nozzle includes a flow passage including a body having first end that extends to a second end through at least one flow channel having a flow area. A fuel inlet is provided at the first end of the body. The fuel inlet is configured to receive at least one fuel. A fuel outlet is provided at the second end of the body. A control flow passage is fluidly connected to the body between the first and second ends. The control flow passage is configured and disposed to deliver a control flow into the fuel nozzle. The control flow establishes a selectively variable effective flow area of the flow passage.

Abstract: A system for mounting an internal fuel manifold within a gas generator case of a gas turbine engine includes at least one fastener support assembly having a fastener pin and at least one receiving element, respectively engaged to either the internal fuel manifold or the gas generator case. A biasing member is located between the fastener pin and the receiving element to axially preload the joint between the fastener pin and the receiving element, such as to constrain relative movement between the two parts in an axial direction while allowing relative movement in a radial direction.

Abstract: An improved mixing tube design and fuel nozzle that allows for a more uniform and thorough mixing of fuel and air being fed to the combustor of a gas turbine engine, wherein each of a plurality of mixing tubes comprises a pair of concentric hollow cylinders that define a ring-like, annular path for the flow of fuel between the two hollow cylinders in each mixing tube, a plurality of air injection slots formed in the concentric hollow cylinders defining corresponding air flow paths from the outside into the interior of each mixing tube, and one or more fuel injection ports formed in selected ones of the plurality of air injection slots that allow for the flow of fuel from the annular path formed by the hollow cylinders into the air flow path, resulting in significantly better mixing and improved thermodynamic behavior of the fuel and air mixture downstream of the nozzle and upstream of the combustor.

Abstract: A turbomachine includes a fluid cooled injection nozzle assembly. The fluid cooled injection nozzle assembly includes an inner conduit portion that includes a body portion having first end portion to a tip end portion. The body portion includes an outer surface and an inner surface. A cooling element extends through the inner conduit portion. The cooling element includes a body element having a first end section that extends to a second end section. The body element includes an outer surface and an inner surface that defines a cooling passage. The outer surface of the body element is spaced from the inner surface of the inner conduit portion to define a return channel. Fluid passing through the cooling passage impinges upon and convectively cools the tip end portion, enters the return channel and is directed out from the nozzle member.

Abstract: A combustor comprises a liquid fuel nozzle for injecting liquid fuel to a combustion chamber, and an air supply nozzle disposed around the liquid fuel nozzle and injecting air. The air supply nozzle is disposed such that air is injected from the air supply nozzle in a direction toward an axis of the liquid fuel nozzle. A space is formed around an outlet of the liquid fuel nozzle, through which the liquid fuel is injected from the liquid fuel nozzle to the combustion chamber, upstream of a distal end of the outlet in a direction in which the liquid fuel is injected. Carbonaceous deposits on surrounding surfaces of the outlet of the liquid fuel nozzle can be suppressed regardless of the operating conditions of a combustor.

Abstract: The present invention is directed to a lean direct injection (LDI) combustion system for a gas turbine using a shell and tube heat exchanger concept to construct a shell and tube lean direct injector (“LDI”) for the combustion system. One side of the LDI injector, either the shell side or the tube side, carries an oxidizer, such as air, to the combustor, while the other side of the LDI injector carries fuel to the combustor. Straight or angled holes drilled in an end plate of the combustor allow the fuel to enter the combustor and mix with air being injected into the combustor.

Abstract: A deflector takes the form of a plate provided with a hole. The plate is a portion of a conical surface of revolution about a cone axis with a concave face and a convex face. The plate includes a contour with four sides. Two of these the sides are concentric circular arcs centered on the cone axis. The other two sides are segments of generatrices of the cone which connect the first sides. The invention is applicable to a deflector of a chamber endwall of a combustion chamber of a turbine engine.

Abstract: A nozzle includes a center body and a shroud circumferentially surrounding at least a portion of the center body to define an annular passage between the center body and the shroud. The nozzle further includes a bimetallic guide between the center body and the shroud. A method for supplying fuel to a combustor includes flowing a working fluid through a nozzle, injecting the fuel into the nozzle, and mixing the fuel with the working fluid to create a fuel and working fluid mixture. The method further includes swirling the fuel and working fluid mixture, sensing flame holding in the nozzle, and reducing the swirl in the fuel and working fluid mixture.

Abstract: A gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

Abstract: Disclosed is a fuel lance for a burner, in particular a gas turbine burner, including a tip that has a nozzle surface including at least two fuel nozzles. The nozzle surface is provided with slots between the fuel nozzles. As a rule, the lance is provided for operation with a liquid fuel.

Abstract: A fuel injection system comprises a fuel conveying member and a nozzle tip assembly threadably engaged thereto. A pair of sealing elements is engaged in a fuel passage between the fuel conveying member and the nozzle tip for sealing the junction therebetween. The pair of sealing elements includes a first and a second sealing element located proximate each other and having different cross-sectional shape.

Abstract: Disclosed is a fuel injector for a gas turbine engine combustor that includes a fuel nozzle for injecting fuel into the gas turbine engine combustor and a fiber optic microphone operatively associated with the fuel nozzle for measuring acoustic pressure differentials within a combustion chamber of the gas turbine engine combustor. The fiber optic microphone includes a fiber bundle having at least one light transmitting fiber and one light receiving fiber; and a dynamic pressure-sensing diaphragm operatively spaced apart from a sensing end of the fiber bundle. The diaphragm has a reflective surface and is formed from a material capable of withstanding temperatures associated with flame exposure. The diaphragm is adapted and configured for deflecting in response to acoustic pressure changes within the combustion chamber.

Abstract: A combustor section is provided and includes one or more annular quaternary fuel manifolds mounted within an annular passage defined between a casing and a cap assembly of a combustor through which air and/or a fuel/air mixture flows upstream from a fuel nozzle support, the manifold including a body to accommodate quaternary fuel therein, the body defining injection holes through which the quaternary fuel is injected into a section of the passage at a location upstream from the fuel nozzle support.

Abstract: A nozzle includes a center body and a shroud circumferentially surrounding at least a portion of the center body to define an annular passage between the center body and the shroud. The nozzle further includes a bimetallic guide between the center body and the shroud. A method for supplying fuel to a combustor includes flowing a working fluid through a nozzle, injecting the fuel into the nozzle, and mixing the fuel with the working fluid to create a fuel and working fluid mixture. The method further includes swirling the fuel and working fluid mixture, sensing flame holding in the nozzle, and reducing the swirl in the fuel and working fluid mixture.

Abstract: A fuel injector nozzle for dispensing fuel in the combustion chamber of a gas turbine engine, comprises a fluid feed conduit having at least one internal channel for the passage of fluid from an inlet end to an outlet end of the fluid feed conduit. The fluid feed conduit has a first annular segment receiving fluid from the inlet end and a second annular segment fluidly connected to receive fluid from the first annular segment at a junction having a circumferential length less than the circumferential lengths of the first and second annular segments. The second annular segment includes fluid dispensing openings to dispense fluid from the conduit, and the first and second annular segments are coaxial and axially separated for relative movement over a major portion of the second segment to accommodate differential thermal expansion.

Abstract: A combustion system for a gas turbine engine is provided. The system includes a forward liner; an aft liner; a combustion chamber formed by the forward liner and the aft liner, the combustion chamber defining a lean combustion zone and a pilot combustion zone; a premixing zone coupled to the combustion chamber; a pilot fuel injector coupled to the combustion chamber and configured to deliver a first flow of fuel to the pilot combustion zone; and a slinger unit configured to deliver a second flow of fuel to the premixing zone such that the second flow of fuel is mixed with air in the premixing zone and directed into the lean combustion zone of the combustion chamber.

Abstract: A flame stabilizer is in fluid communication with a combustor of a gas turbine engine. The flame stabilizer has a body with an aerodynamic shape that creates a flow recirculation zone by injection of fluid through a plurality of holes in the body of the flame stabilizer. The aerodynamic shape of the body reduces pressure loss in the combustor, particularly when no fuel is being provided to the combustor. In addition, the magnitude of the flow recirculation zone can be modulated by selectively adjusting the flow rate of fluid through the holes, or by selectively adjusting the size of the holes.

Abstract: A gas turbine fuel injector includes a nozzle body having a radially inner wall proximate to an internal air path and a radially outer wall. An insulative gap is defined between the radially inner and outer walls. The inner and outer walls are adapted and configured for relative axial movement at a first interface. An inhibitor ring is disposed proximate a downstream end of the inner wall for discouraging fuel from entering the insulative gap. A second interface is formed between the downstream end of the inner wall and an upstream end of the inhibitor ring to accommodate relative axial movement of the inner and outer walls.

Abstract: A fuel injector for a gas turbine engine of an aircraft, and more particularly a novel and unique heatshield structure for a fuel nozzle wherein a labyrinth seal is uniquely provided in the nozzle to isolate a portion of an insulating gap from an interface whereat fuel may enter the insulating gap, and the insulating gap is provided with a positive purge flow for forcing vapors out of the insulating gap.

Abstract: A fuel injector for injecting fuel into a turbomachine combustion chamber, the injector presenting an injector nose for connection to an injection system fastened to the end wall of said combustion chamber, the injector nose comprising: a first channel for passing a flow of an air/fuel pre-mixture towards the combustion chamber, the first channel opening out into the center of the injector nose via an outlet opening for the air/fuel pre-mixture; an electrical insulator surrounding said outlet opening for the air/fuel pre-mixture; a plasma generator system disposed inside said electrical insulation and downstream from said outlet opening for the air/fuel pre-mixture in order to control ignition and combustion of the air/fuel pre-mixture; a second channel for passing a flow of fuel towards the combustion chamber, the second channel opening out outside said electrical insulation; and a third channel for passing a flow of fuel towards the combustion chamber, this third channel opening out outside the second ch

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 and injector system to inject a selected fuel into a combustor of a gas powered turbine. Generally, the injector is able to inject a selected fuel into a stream of an oxidizer to substantially mix the fuel with the oxidizer stream before any of the fuel in the fuel fan reaches an auto ignition temperature. Therefore the fuel may be substantially combusted at once and without any substantial hot spots.

Abstract: A fuel injection system for an annular combustion chamber of a turbomachine, the system comprising support means for supporting and centering a fuel injector head, and a bowl arranged downstream from the support means and including at its downstream end an annular collar that extends radially outwards and that is cooled by air impacting against its upstream radial surface, and including on said surface means for disturbing the flow of cooling air and for increasing the heat exchange area between the air and the collar.

Abstract: A fuel injection nozzle includes at least one tube disposed in the nozzle having a venturi shaped profile defining a gas flow path including an inlet operative to receive a first gas, at least one port operative to emit a second gas into the gas flow path, and an outlet operative to emit a mixture of the first gas and the second gas into a combustor.

Abstract: An injection nozzle (10), especially for injecting liquid fuel, preferably crude oil, into the combustion chamber of a gas turbine, includes an inner chamber (16) which extends along a nozzle axis (19), conically tapers to a concentric nozzle orifice (17), and to which the medium which is to be injected is fed from outside through a plurality of inlet ports (18) which are arranged in a distributed manner around the nozzle axis (19). The inlet ports (18) are oriented perpendicularly to the nozzle axis (19) and each lead tangentially into the inner chamber (16). With such an injection nozzle, an improved spray cone is achieved by the fact that a pin (14), which extends in the axial direction, is concentrically arranged in the inner chamber (16) and passes through the region of the mouths of the inlet ports (18) and extends right into the nozzle orifice (17).

Abstract: A fuel conveying member for a gas turbine engine, the fuel conveying member including a fuel manifold having a plurality of fuel nozzles extending therefrom. The fuel manifold is mounted within the engine by a support system, including integral attachment lugs on the fuel manifold for mounting the fuel manifold within the gas turbine engine. The attachment lugs re adapted to receive pins therein and provide a mounting mechanism which allows for thermal expansion of the fuel manifold relative to a surrounding support structure to which the fuel manifold is mounted via the attachment lugs.

Abstract: Fuel injector assemblies with frictionally damped fuel supply members, including fuel feed strips. More particularly, the invention provides friction dampers and/or assemblies that frictionally damp movement of fuel supply members in at least one direction. Some of the embodiments provide a friction damper that is easily serviceable, and can be installed after final assembly of a fuel injector. Aspects of the invention are applicable to other components of fuel injectors and gas turbine engines in addition to fuel supply members.

Abstract: Fuel injection apparatus 54, for a gas turbine engine, includes pilot fuel injection circuits 42 which supply pilot injectors 44 with fuel from a manifold 60, through a pilot control valve 46. The state of the control valve 46 is set by pneumatic pressure in the manifold 56. Main injectors 50 are supplied with fuel from the manifold 60 by main injection circuits 48, through main valves 52. The state of the main valves 52 is set by hydraulic pressure in a manifold 58. Accordingly, control signals to the valves 46, 52 are conveyed independently by means of the pneumatic manifold 56 and the hydraulic manifold 58.

Abstract: A combustor assembly for a turbine engine includes a fuel nozzle located at an upstream end of the combustor assembly. The fuel nozzle includes both a fuel delivery passageway and a purge air passageway. The purge air passageway conveys a flow of purge air to cool the fuel nozzle. The combustor assembly also includes a combustion product return line that conveys a flow of combustion products from a position downstream of a combustion zone of the combustor back to the fuel nozzle. The combustion products are mixed with purge air, and the mixture is delivered into a combustion zone located just downstream of the fuel nozzle. The addition of the combustion products into the purge air reduces the amount of oxygen in mixed flow, which helps to reduce the generation of undesirable nitrogen oxide combustion byproducts. Also, the greater heat of the combustion products helps to increase the temperature of the mixture, which helps to maintain the stability of a pilot flame fed by the mixture.

Abstract: A combustor liner assembly includes a liner assembly, including a hot dome, a hot wall, and an impingement shield. The liner assembly further includes a convector assembly. The combustor liner assembly further includes an impingement dome assembly, including a ring-shaped portion defining a plurality of apertures configured to receive a plurality of fuel injectors, and a face. The ring-shaped portion and the face define an annular passage, and the face defines a plurality of apertures and a plurality of orifices. The convector assembly and the impingement dome assembly are operably coupled to one another, and the plurality of orifices of the face are configured such that the hot dome is exposed to air flowing in the annular passage of the impingement dome assembly. The combustor liner assembly is configured to provide flow communication between the annular passage of the impingement dome assembly and the plurality of fuel injectors.

Abstract: Disclosed is a method of operating a secondary fuel nozzle for a turbomachine combustor including delivering a flow of pilot fuel through a pilot fuel channel toward a combustion zone and delivering a flow of air through a plurality of transfer passages arranged around the pilot fuel channel toward the combustion zone. The flow of pilot fuel and the flow of air are combusted in the combustion zone, and a flow of transfer fuel is delivered through the plurality of transfer passages for combustion. A secondary fuel nozzle includes a pilot fuel channel extending along the fuel nozzle to deliver a flow of pilot fuel to a combustion zone. A plurality of transfer passages are arranged around the pilot fuel channel and are configured to deliver a flow of air for combustion with the flow of pilot fuel and to deliver a flow of transfer fuel to the combustion zone.

Abstract: A fuel nozzle for a gas turbine includes a fuel nozzle body or endcover. The fuel nozzle body or endcover includes at least one seal surface. The fuel nozzle also includes at least one flexible hose and at least one flange connected to the flexible hose. The at least one flange is fastened to the fuel nozzle body at the at least one seal surface.

Abstract: A secondary fuel nozzle for a turbine includes a passive purge air passageway which provides purge air to the secondary nozzle at all times that the nozzle is in operation. The passive purge air passageway draws in air from a location adjacent an upstream end of the nozzle. Because of a pressure differential between air located at the downstream end of the nozzle and air located at the upstream end of the nozzle, purge air will run through the passive purge air passageway at all times the nozzle is in operation. There is no need for a supply of compressed purge air.

Abstract: An article, such as a bucket, a blade, a nozzle, a vane, a strut, a fuel nozzle, a combustion casing, and a transition piece, the article having a damping coating comprising approximately 8 wt % to approximately 15 wt % Y2O3 and approximately 19 wt % to approximately 28 wt % Ta2O5 with the balanced weight of ZrO2 is presented.

Abstract: Fuel injector assemblies with frictionally damped fuel supply members, including fuel feed strips. More particularly, the invention provides friction dampers and/or assemblies that frictionally damp movement of fuel supply members in at least one direction. Some of the embodiments provide a friction damper that is easily serviceable, and can be installed after final assembly of a fuel injector. Aspects of the invention are applicable to other components of fuel injectors and gas turbine engines in addition to fuel supply members.

Abstract: A fuel injector for a gas turbine combustor is disclosed which includes a fuel nozzle for issuing fuel into the combustor for combustion, at least one photodetector responsive to flame radiation indicative of an OH chemiluminescence peak, and at least one photodiode responsive to flame radiation indicative of a CH chemiluminescence peak.

Abstract: A gas turbine combustion chamber includes an annular combustion chamber (3) confined by an outer (1) and an inner combustion chamber wall (2) and is provided with several fuel injection nozzles (4). Air deflectors (6) are arranged in a front-side area of the annular combustion chamber (3) over the entire combustion chamber annulus (5) concentrically around a combustion chamber axis (12) and fuel injection nozzles (4) are distributed essentially uniformly over the combustion chamber annulus (5).

Abstract: In one embodiment, a pulse detonation system includes a pulse detonation tube and an air valve disposed at an upstream end of the pulse detonation tube. The air valve is configured to provide an air flow into the pulse detonation tube. The pulse detonation system also includes a fuel injector configured to inject fuel into the air flow to establish a fuel-air mixture configured to support detonation, and to establish a region in the fuel-air mixture having a fuel to air ratio insufficient to support a detonation wave. The pulse detonation system further includes an ignition source configured to detonate the fuel-air mixture when the region is disposed adjacent to the air valve.

Abstract: The present invention discloses a combustor system and method of measuring impurities in the combustion system. The combustion system includes an up-stream fuel injection point; a down-stream turbine combustor; a flame zone in the turbine combustor comprising a plurality of axial sub-zones; an optical port assembly configured to obtain a non-axial, direct, optical view of at least one of the plurality of axial sub-zones, and an impurity detection system in optical communication with the optical port assembly.

Abstract: In one embodiment, a system includes a fuel nozzle that includes a fuel injector that includes a fuel port and a premixer tube. The premixer tube includes a wall disposed about a central passage, multiple air ports extending through the wall into the central passage, and a catalytic region. The catalytic region includes a catalyst, disposed inside the wall along the central passage, configured to increase a reaction of fuel and air.

Abstract: A secondary fuel nozzle is formed by brazing and welding a central core, a cylindrical body, and a unitary secondary fuel nozzle portion. The central core has a central bore for receiving a liquid fuel cartridge, and at least two passages concentric with the central bore. The secondary fuel nozzle portion includes a base flange and a main shaft having a central bore for aligning with the central bore of the core and for receiving the liquid fuel cartridge, and a plurality of peripheral bores disposed at spaced locations about the central bore for flow communication with the concentric passages of the core. The core is brazed to the base flange and the cylindrical body is disposed around the core and welded to the base flange to define a one-piece secondary fuel nozzle assembly.