Abstract: An injector includes a plurality of flat, circular plates which provide fuel delivery and cooling of the injector. Fuel delivery passages in the plates have swirl chambers and spray orifices which are formed by chemical etching. A pair of fuel delivery plates define a fuel cavity therebetween, and include a plurality of radially-outwardly extending spokes, with the spokes from one fuel plate in adjacent, surface-to-surface relation with opposing spokes from the adjacent fuel plate. A fuel passage is defined between each of the opposing spokes, leading from the fuel cavity to a fuel outlet opening at the distal end of each spoke. A fuel tube delivers fuel to the fuel cavity between the plates, from where the fuel is then directed through the outlet openings. Downstream plates shape the fuel into appropriate sprays for ignition. An upstream cooling plate assembly directs air against the upstream fuel plate, and radially outwardly along the spokes of the upstream plate.

Abstract: A multiplex injector system comprising an injector head, a first fuel path located in the injector head, and a first set of injector tips located in the injector head and in fluid communication with the first fuel path. The first set of injector tips includes at least one first injector tip. The multiplex injector further includes a second fuel path located in the injector head and a second set of injector tips located in the injector head and in fluid communication with the second fuel path. The second set of injector tips includes at least one second injector tip. A flow of fuel in each of the first and second fuel paths can be selectively controlled to control the flow of fuel through the first and second sets of injector tips.

Abstract: A gas turbine engine fuel injector conduit includes a single feed strip having a single bonded together pair of lengthwise extending plates. Each of the plates has a single row of widthwise spaced apart and lengthwise extending parallel grooves. Opposing grooves in each of the plates are aligned forming internal fuel flow passages through the strip from an inlet end to an outlet end. The feed strip includes a substantially straight middle portion between the inlet end and the outlet end. In one alternative, the middle portion has a radius of curvature greater than a length of the middle portion. The feed strip has at least one acute bend between the inlet end and the middle portion and a bend between the outlet end and the middle portion. The feed strip has fuel inlet holes in the inlet end connected to the internal fuel flow passages.

Abstract: A premix fuel nozzle and method of operation for use in a gas turbine combustor is disclosed. The premix fuel nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of fuel and compressed air in order to provide a more uniform injection pattern. The fuel and compressed air mixes upstream of the combustion chamber and flows into the combustion chamber as a homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide fuel and compressed air to the fin assembly, as well as compressed air to cool the nozzle cap assembly. An alternate embodiment includes an additional fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly.

Abstract: A dual fuel premix nozzle and method of operation for use in a gas turbine combustor is disclosed. The dual fuel premix nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of fuel and compressed air in order to provide a more uniform injection pattern. When in gas operation, the fuel and compressed air mixes upstream of the combustion chamber and flows into the combustion chamber as a homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide gaseous fuel and compressed air to the fin assembly. When in liquid fuel operation, the gas circuits are purged with compressed air and liquid fuel and water pass through coaxial passages to the tip of the dual fuel premix fuel nozzle, where they inject liquid fuel and water into the secondary combustion chamber.

Abstract: A dual fuel premix nozzle and method of operation for use in a gas turbine combustor is disclosed. The dual fuel premix nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of gas fuel and compressed air in order to provide a more uniform injection pattern and homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide gaseous fuel and compressed air to the fin assembly. When in liquid fuel operation, the gas circuits are purged with compressed air and liquid fuel and water pass through coaxial passages to the tip of the dual fuel premix fuel nozzle, where they inject liquid fuel and water into the secondary combustion chamber. An alternate embodiment includes an additional gas fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly.

Abstract: A premix fuel nozzle and method of operation for use in a gas turbine combustor is disclosed. The premix fuel nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of fuel and compressed air in order to provide a more uniform injection pattern. The fuel and compressed air mixes upstream of the combustion chamber and flows into the combustion chamber as a homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide fuel and compressed air to the fin assembly, as well as compressed air to cool the nozzle cap assembly. An alternate embodiment includes an additional fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly. A second alternate embodiment is disclosed which reconfigures the injector assembly and fuel injection locations to minimize flow blockage issues at the injector assembly.

Abstract: A premix fuel nozzle and method of operation for use in a gas turbine combustor is disclosed. The premix fuel nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of fuel and compressed air in order to provide a more uniform injection pattern. The fuel and compressed air mixes upstream of the combustion chamber and flows into the combustion chamber as a homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide fuel and compressed air to the fin assembly, as well as compressed air to cool the nozzle cap assembly. An alternate embodiment includes an additional fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly.

Abstract: A dual fuel premix nozzle and method of operation for use in a gas turbine combustor is disclosed. The dual fuel premix nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of gas fuel and compressed air in order to provide a more uniform injection pattern and homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide gaseous fuel and compressed air to the fin assembly. When in liquid fuel operation, the gas circuits are purged with compressed air and liquid fuel and water pass through coaxial passages to the tip of the dual fuel premix fuel nozzle, where they inject liquid fuel and water into the secondary combustion chamber. An alternate embodiment includes an additional gas fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly.

Abstract: A premix fuel nozzle and method of operation for use in a gas turbine combustor is disclosed. The premix fuel nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of fuel and compressed air in order to provide a more uniform injection pattern. The fuel and compressed air mixes upstream of the combustion chamber and flows into the combustion chamber as a homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide fuel and compressed air to the fin assembly, as well as compressed air to cool the nozzle cap assembly.

Abstract: A fuel injector nozzle for dispensing fuel in the combustion chamber of a gas turbine engine, includes an elongated, multi-layered, convoluted nozzle feed strip having an internal passage for directing fuel through the length of the strip from the inlet end to an outlet end; and a cylindrical, multi-layered fuel dispensing nozzle unitary with the feed strip and fluidly connected to the outlet end of the feed strip for dispensing the fuel. The multi-layered feed strip and nozzle allows complex porting of fuel circuits through the injector. The internal fluid passages through the feed strip and nozzle are formed by etching.

Abstract: A dual circuit pilot burner fuel delivery arrangement (60) for a gas turbine engine combustor (18). A first fuel delivery circuit (64) delivers a gaseous fuel from a fuel supply (62) to a first diffusion outlet orifice (74). A second fuel delivery circuit (78) connects a second diffusion outlet orifice (82) to the first fuel delivery circuit through a pressure-regulated valve (86). Fuel is delivered to the second diffusion outlet orifice only when the fuel pressure in the first fuel delivery circuit is above a predetermined fuel pressure. At low fuel flow rates, the valve to the second fuel delivery circuit is closed, thereby generating a relatively high nozzle pressure drop across the first diffusion outlet orifice. This relatively high pressure drop makes the system less susceptible to acoustic instabilities that may otherwise result from feedback between the combustion dynamics and the fuel system.

Abstract: A duplex fuel injector has concentric arrangement of an inner air swirler, an inner fuel injection port, an intermediate axial air filmer, a second air swirler, an outer fuel injection port and an outer air swirler. The intermediate air filmer produces a curtain of air separating an outer air and fuel spray from an inner air and fuel spray. The outer spray is used in a pilot combustion zone whereas the inner spray is used in a main combustion zone.

Abstract: A fuel nozzle for a gas turbine engine injects a liquid fuel flow from a liquid fuel passage in the swirler vane. An air flow over the swirler vane atomizes the liquid fuel flow to form a fuel air mixture. The fuel nozzle eliminates the need for a conventional air blast atomizer.

Abstract: A gas turbine (12) capable of combusting a low quality gaseous fuel having a ratio of flammability limits less than 2, or a heat value below 100 BTU/SCF. A high quality fuel is burned simultaneously with the low quality fuel to eliminate instability in the combustion flame. A sensor (46) is used to monitor at least one parameter of the flame indicative of instability. A controller (50) having the sensor signal (48) as input is programmed to control the relative flow rates of the low quality and high quality fuels. When instability is detected, the flow rate of high quality fuel is automatically increased in relation to the flow rate of low quality fuel to restore stability.

Abstract: In a gas turbine having a compressor, a combustor and a turbine, a gaseous fuel supply coupled to provide gaseous fuel to the combustor, a liquid fuel supply coupled to provide liquid fuel to the combustor via nozzle assembly. The nozzle assembly includes a plurality of passageways for flowing a fluid into the combustor, one of the passageways being a fuel passageway conduit interconnect the fuel passageway to one of the plurality of passageways to enable fluid flow therebetween. High pressure air from one of the plurality of passageways is diverted into the fuel passageway via the conduit to protect nozzle from ingestion of hot combustor gases.

Abstract: An assembly method in which a brazing metal is initially filled into radial wells pierced in an annular injection piece having first injection orifices for discharging a primary fuel, and in a cylindrical endpiece surrounding the annular injection piece and having secondary injection orifices for discharging a secondary fuel; thereafter the annular injection piece is fitted inside the cylindrical endpiece and these pieces are together fitted on a first tube for feeding primary fuel and a second tube for feeding secondary fuel and surrounding the first tube, and also to an outer wall of the injector; finally, the end portion of the injector as assembled in this way is placed in an enclosure where it is heated so as to melt the brazing metal to unite the parts.

Abstract: A single circuit fuel injector apparatus having a bifurcated recirculation zone is provided. The single circuit injector includes an injector tip having an aft facing tapered surface which is communicated with a plurality of fuel injector ports. A radially inward tapered conical air splitter directs sweep air over the tapered injector tip. An air blast atomizer filmer lip is disposed concentrically outward from the tapered tip. In a low power operating mode, fuel exiting the fuel injector ports is entrained within a centralized sweep air stream. In a high power operating mode, the majority of the fuel exiting the fuel injection ports has sufficient momentum to carry it across the sweep air stream so that it falls upon the main fuel filmer lip and is entrained in an outer main air stream.

Abstract: A fuel and gas mixer for a gas turbine combustor includes a body member. The body member has a truncated conical annulus, with its smaller end which communicates with a smaller end of a downstream truncated conical chamber. Gaseous fuel is injected into the annulus through hollow spokes arranged radially in the annulus and mixes with air flow introduced from air upstream passages. The velocity of the primary mixture of the fuel and air in the annulus increases when the flow is directed towards the downstream chamber because the truncated conical shape of the annulus reduces the cross-section of the flow passageway downstream-wise. With increased velocity, the primary mixture is diffused when entering the truncated conical chamber and is further mixed with air introduced from downstream air passages.

Abstract: A gas-turbine engine combustion system has a combustor with a burner head having both pilot gas and pilot liquid-fuel injection arrangements, the pilot gas arrangement comprising an annular gallery communicating with a downstream face of the head and a deflecting arrangement adjacent the gallery for directing the pilot gas-fuel towards a longitudinal axis of the combustor and over a central part of the downstream face. The combustion system is designed so that, during both gas- and liquid-fuel operations, the flame front face is located close to the burner head and, during liquid-fuel operation, air is forced across the downstream face to cool the head. Advantageously, the cooling air is made to replace the pilot gas-fuel in the annular gallery, so that it is deflected, like the gas-fuel, and contacts the central part of the downstream face. The burner head also features main gas and liquid-fuel injection arrangements, these communicating with one or more passageways in a radial swirler attached to the head.

Abstract: Gas turbine engine fuel nozzles are illustrated which induce swirling to fuel flowing to the engine to facilitate reducing fuel coking. Each fuel nozzle includes an inlet, an outlet and a fuel delivery system extending therebetween. The fuel delivery system includes an inner fuel supply tube and an outer fuel supply tube. The inner fuel supply tube is concentrically aligned within the outer fuel supply tube and includes contoured fuel passageways and a center axis of symmetry. As fuel enters the contoured passageways, the fuel is accelerated locally and directed angularly with respect to the axis of symmetry.

Abstract: A burner for use in a combustion system of a heavy-duty industrial gas turbine includes a fuel/air premixer having an air inlet, a fuel inlet, and an annular mixing passage. The fuel/air premixer mixes fuel and air into a uniform mixture for injection into a combustor reaction zone. The burner also includes an inlet flow conditioner disposed at the air inlet of the fuel/air premixer for controlling a radial and circumferential distribution of incoming air. The pattern of perforations in the inlet flow conditioner is designed such that a uniform air flow distribution is produced at the swirler inlet annulus in both the radial and circumference directions. The premixer includes a swozzle assembly having a series of preferably air foil shaped turning vanes that impart swirl to the airflow entering via the inlet flow conditioner.

Abstract: A dual fuel nozzle is provided with two different size injection holes. The first injection holes have larger diameters and are used only for injecting gaseous fuel into a combustion chamber. On the other hand, the second injection nozzles have smaller diameters and are used for injecting either gaseous fuel or liquid fuel as required. When gaseous fuel is used, if the fuel injection amount is large or medium, both of the first and the second injection holes or first injection holes only are used for injecting gaseous fuel depending upon the required fuel injection amount. When the fuel injection amount is low, only the second injection hole is used for injecting gaseous fuel. Therefore, the pressure drop across the fuel nozzle can be kept at sufficiently high level even when the fuel injection amount is low, and thereby combustion vibration is suppressed. Further, when liquid fuel is used, a premixed fuel and steam mixture is injected from the second injection holes.

Abstract: An improved bulkhead for dual fuel industrial and aeroengine gas turbines, that may be formed by one or more bulkhead elements. Each element has a fuel-air channel, having an inlet for introducing air and an outlet for exiting a fuel-air mixture. Each element further includes at least one manifold for introducing a liquid or a gaseous fuel into the fuel-air channel. The channel and the manifold are formed by a plurality of etched layers of macrolaminate or platelet material. Each element may further include an air cooling channel for cooling the backside of the bulkhead. The bulkhead may be used in can and in annular combustors.

Abstract: A method and apparatus for reducing emissions in combustion systems, particularly gas turbines. A mixture of diluent and fuel is created, wherein the diluent and the fuel are at a predetermined diluent-to-fuel ratio. The mixture is homogenized to create a homogenized mixture having a uniform concentration distribution of the diluent and the fuel at the predetermined diluent-fuel ratio. Thereafter, the homogenized mixture is introduced into a flame zone and the homogenized mixture is combusted.

Abstract: A gas turbine combustor burning LBTU fuel gas serves to be applied to the combustion system of a small turbogenerator. The combustion system is composed of a combustor outer case, a combustor liner, a combustor transition section, a radial swirler with adjustable blades and fuel supply passages. The small turbogenerator (10 KW) with the redesigned combustion system is integrated by a LBTU gas generator. A recirculation bubble with proper size and strength is aerodynamically formed by the interaction among the swirling air jet, inclined fuel jet and the primary jets. Since an adjustable swirler is installed, the swirl number of the swirling air jet can be modified to meet the requirements of combustion load sharing and burning different LBTU fuel gases. To eliminate the possible hot spots and thus reduce the pattern factor value, two rows of cooling jet holes are arranged in the rear section of the combustor.

Abstract: The present invention relates to a combustion chamber of a gas turbine working on liquid fuel, comprising a tubular enclosure (2) having at least one pressurized air inlet, a liquid fuel injection means (6) positioned on or in proximity to the longitudinal axis XX′ of the tubular enclosure, an outlet to the turbine.

Abstract: A fuel delivery system and method for operating a dual fuel and steam turbine engine that utilizes gas fuel at the engine start phase through a combination of gas and liquid fuel circuits. Gas fuel initially flows through the liquid fuel circuits to a dual fuel and steam nozzle. A boiler connected to the system generates steam, and the steam is supplied to the nozzle to raise the fuel pressure to an acceptable range. The gas fuel is injected into both the liquid and gas fuel circuits and the engine can then operate on either gas or liquid fuel, or any combination thereof.

Abstract: A pre-mixing chamber (10) for gas turbines, wherein the gas turbine is of the type which comprises at least one compressor connected to the turbine, and a combustion chamber, wherein the combustion chamber has a converging portion (12), positioned such as to give rise to combustion inside the combustion chamber, and wherein the pre-mixing chamber (10) has a plurality of pipes (19) provided with holes (20), which open into the combustion chamber on a front portion of the body of the pre-mixing chamber (10), in order to generate a series of pilot flames, which are appropriately regulated, in order to stabilise a main flame, which is primed inside the combustion chamber, in which a front area of the converging portion (12) of the pre-mixing chamber (10) has at least one circular groove (28) provided at the holes (20) which belong to the pies (19).

Abstract: A fuel/mixer injection system for a combustor of a gas turbine engine comprising two major assemblies including a burner-mounted swirler with two outer air passages surrounding a central passage and a piloted air fuel nozzle containing a main fuel injection orifice and a pilot fuel injection orifice the main fuel mixes with swirling air in the central passage and the pilot fuel mixes with the swirling air in the outer air passage. In another embodiment, the piloted fuel nozzle contains an airblast-atomized main fuel injector annulus and a concentric ring with a plurality of circumferentially spaced holes feeds air to the main fuel injector annulus.

Abstract: A fuel injection system for a gas turbine engine combustor, wherein the combustor includes a dome inlet module having a plurality of flow passages formed therein and at least one cavity formed in a liner downstream of said dome inlet module. The fuel injection system includes a fuel supply and a plurality of fuel injector bars positioned circumferentially around and interfacing with the inlet dome module. The fuel injector bars are in flow communication with the fuel supply, with each of the fuel injector bars further including a body portion having an upstream end a downstream end, and a pair of sides. A plurality of injectors formed in the body portion and in flow communication with the fuel supply, whereby fuel is provided to the dome inlet module flow passages and/or the cavity through the fuel injector bars.

Abstract: A gas turbine fuel injection system of the lean direct injector type designed to reduce nitrous oxide (NOx) emissions is provided. The configuration includes a pilot fuel injector for injecting a pilot fuel stream, and a pilot swirler for providing a swirling pilot air stream to atomize and entrain the pilot fuel stream. A main airblast fuel injector is located concentrically about the pilot fuel injector, for injecting a main fuel stream concentrically about the pilot fuel stream. Inner and outer main swirlers provide a swirling main air stream to atomize and entrain the main fuel stream. An air splitter is located between the pilot swirler and the main swirler. The air splitter is so arranged and constructed as to divide the pilot air stream exiting the pilot swirler and the air splitter, from the main air stream exiting the inner main swirler, whereby a bifurcated recirculation zone is created.

Abstract: A three stage lean burn combustion chamber (28) comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (76,78,80,92). The primary fuel and air mixing ducts (78,78) comprise first and second radial flow swirlers (60,62). The primary fuel injectors (64,66) inject fuel into the first and second swirlers (60,62). The primary fuel injectors (44,66) and the first and second swirlers (60,62) are arranged such the fuel to air ratio of the fuel and air flowing from the passages (61) of the first swirler (60) into the primary combustion zone (36) is different to the fuel to air ratio of the fuel and air flowing from the passages (65) of the second swirler (62) into the primary combustion zone (36).

Abstract: A fuel/mixer injection system for a combustor of a gas turbine engine comprising two major assemblies including a burner-mounted swirler with two outer air passages surrounding a central passage and a piloted air fuel nozzle containing a main fuel injection orifice and a pilot fuel injection orifice the main fuel mixes with swirling air in the central passage and the pilot fuel mixes with the swirling air in the outer air passage. In another embodiment, the piloted fuel nozzle contains an airblast-atomized main fuel injector annulus and a concentric ring with a plurality of circumferentially spaced holes feeds air to the main fuel injector annulus.

Abstract: A fuel injector (32) for a gas turbine engine combustion chamber (22) comprises a fuel injector nozzle (34) and a fuel feeding member (36). A wire mesh (76) is located in a space (74) defined between a hollow outer member (72) arranged around and spaced from a hollow inner member (70) of the fuel feeding member (36) and a fuel nozzle body (54) of the fuel injector nozzle (34). The wire mesh (76) rubs against the hollow outer member (72), the hollow inner member (70) and the fuel nozzle body (54) to frictionally damp vibrations of the fuel injector (32).

Abstract: A fuel nozzle including a nozzle stem having an annular overhang and a heat shield secured to the overhang is described. More specifically, and in one embodiment, the nozzle stem includes an upstream end and a downstream end. The annular overhang is intermediate to the upstream end and the downstream end of the stem. The heat shield includes a first end and a second end, and the heat shield is welded to the annular overhang at the heat shield first end. An annular air gap is between the nozzle stem and the heat shield.

Abstract: Compressor discharge air flows through a heat exchanger in heat exchange relation with cooling water and is supplied at reduced temperature in atomizing air and purge modes to combustors in a dual-fuel gas turbine. The heat exchanger has a pair of bypass passages in parallel with a temperature regulating valve controlling flow of water through the heat exchanger in response to the temperature of the compressor discharge air exiting the heat exchanger. Should the flow control valve close in response to low temperature of the compressor discharge air, in the air atomizing mode, the bypass passages with orifices therein provide a minimum protective flow of cooling water to the heat exchanger. In the purge mode where only half the air flow and heat rejection is required in the heat exchanger and the flow control valve closes, a bypass valve in one of the bypass passages closes to enable reduced flow of cooling water through the heat exchanger.

Abstract: A stem member for a gas turbine fuel nozzle includes inlet and outlet ends which are respectively adapted to be connected to a fuel adapter which is coupled to a fuel injector and a tip assembly having at least one spray orifice for atomizing fuel into a combustion chamber. The stem member further includes at least one slot which is sealed throughout the length thereof by a slot cover so as to define at least one fuel conduit for directing fuel flow from the inlet end to the outlet end of the stem member. An outer shield can be disposed outwardly of the stem member to protect and limit the transfer of heat from the surroundings to the stem member.

Abstract: A combustor for a gas turbine having first and second passages for pre-mixing primary fuel and air supplied to a primary combustion zone. The flow of fuel to the first and second pre-mixing passages is separately regulated using a single annular fuel distribution ring having first and second row of fuel discharge ports. The interior portion of the fuel distribution ring is divided by a baffle into first and second fuel distribution manifolds and is located upstream of the inlets to the two pre-mixing passages. The annular fuel distribution ring is supplied with fuel by an annular fuel supply manifold, the interior portion of which is divided by a baffle into first and second fuel supply manifolds. A first flow of fuel is regulated by a first control valve and directed to the first fuel supply manifold, from which the fuel is distributed to first fuel supply tubes that direct it to the first fuel distribution manifold.

Abstract: A low emissions combustor can 18 for an aircraft gas turbine engine includes a louvered combustor liner 24 with three arrays of dilution holes 52, 54, 56. The first hole array comprises twelve equally sized, equiangularly distributed holes that penetrates the liner about midway along its axial length. The second hole array comprises twelve equally sized holes, smaller than the holes of the first array, that penetrate the liner a predetermined distance aft of the first hole array. The holes of the second array are equiangularly distributed and each second hole is circumferentially aligned with a first hole. A third hole array penetrates the liner a predefined distance aft of the first array. The third holes are nonuniformly sized and nonequiangularly distributed to regulate the spatial temperature profile of combustion gases exiting the combustor can. The quantity, size, distribution and location of the holes mitigates undesirable exhaust emissions without affecting the performance or durability of the engine.

Abstract: A gas turbine fuel injector has a fuel mixing barrel, including an air swirler incorporating a fuel flow outlet, suspended in the combustor air stream by a cantilevered injector body. The cantilevered body is mounted in the combustor wall and includes a fuel flow control valve in a valve chamber located in a part of the body in a relatively cool air surrounding the combustor. The fuel outlet is located towards the downstream end of the swirler barrel. A purge passage is provided from a location towards the upstream end of the barrel extending into the outlet side of the valve chamber. When fuel flow is shut-off a pressure differential between the upstream and downstream ends of the swirler barrel causes an air flow up the purge passage into the valve chamber and down the main fuel passage to purge residual fuel.

Abstract: A vibration damper 38 for a fuel nozzle 24 having a central axis about which are disposed an inner 34 and outer 36 concentric fuel tubes. The damper includes a sleeve 40 and at least two legs 42, each leg having a radial portion extending from the sleeve and a resilient, longitudinally extending portion 46. The sleeve engages the inner tube while the longitudinally extending portions of the legs bear against the inner surface 37 of the outer tube to dampen vibrations between the concentric tubes. Besides damping objectionable vibratory forces experienced by the fuel tubes in the fuel nozzle, the present invention offers minimal fuel flow blockage in the concentric fuel tubes.

Abstract: The burner arrangement for a gas turbine comprises at least one burner (2), which is arranged in the plenum of the gas turbine and leads with an inner injection space into a combustion chamber and to which compressed air is admitted on the outside from a compressor stage of the gas turbine. Furthermore, a fuel lance (9) for the alternative feeding of liquid and/or gaseous fuels is allocated to the burner (2), which fuel lance (9) has a central liquid-fuel tube (19) and a pilot gas tube (11) concentrically surrounding the liquid-fuel tube (19), the tubes (11, 19) ending in associated outlet openings (22, 23, 23a-c) in a lance head (12) at the tip of the fuel lance (9).

Abstract: A fuel injection system is disclosed for a gas turbine engine having a fuel injector issuing a primary fuel flow during all operating modes of the gas turbine engine and a secondary fuel flow when operating parameters of the gas turbine engine exceed predetermined values. The fuel injection system includes a first air swirler, a second air swirler located rearwardly of the first air swirler and a venturi having a wall with a converging-diverging inner wall surface forming a venturi throat, the venturi being located such that the wall separates first and second air flows issuing from the first and second air swirlers and such that the primary fuel flow emanating from the fuel injector in a conical configuration does not impact against the inner wall surface of the venturi.

Abstract: A fuel nozzle of a variety for dispensing an atomized fluid spray into the combustion chamber of a gas turbine engine. The nozzle includes a first fluid conduit extending along a longitudinal central axis from an upstream end to a downstream end portion terminating to define a primary discharge orifice, a second fluid conduit which is received coaxially over the first fluid conduit as extending along the central axis from an upstream end to a downstream end portion terminating to define a secondary discharge orifice disposed generally concentrically with said primary discharge orifice, and a third fluid conduit which is received coaxially over the second fluid conduit as extending along the central axis from an upstream first end to a downstream second end. The outer surface of the downstream end portion of the first fluid conduit and the inner surface of the downstream end portion of the third fluid conduit define therebetween a generally annular, internal mixing chamber.

Abstract: A low emissions combustion system for a gas turbine engine, including a generally annular combustor with a plurality of tangential fuel injectors to introduce a fuel/air mixture. The annular combustor includes a skirt shaped flow control baffle from the inner liner and air dilution holes in the inner liner underneath the flow control baffle and also in the cylindrical outer liner. The fuel injectors extend through the recuperator housing and into the combustor through an angled tube and then through a tangential guide in the cylindrical outer liner of the combustor housing. The fuel injectors generally comprise an elongated injector tube with the outer end including a coupler having at least one fuel inlet tube. Compressed combustion air is provided to the interior of the elongated injector tube from either holes or slits therein which receive compressed air from the angled tube around the fuel injector which is open to the space between the recuperator housing and the combustor.

Abstract: A combustor for a gas turbine having a centrally located fuel nozzle and inner, middle and outer concentric cylindrical liners, the inner liner enclosing a primary combustion zone. The combustor has an air inlet that forms two passages for pre-mixing primary fuel and air to be supplied to the primary combustion zone. Each of the pre-mixing passages has a circumferential array of swirl vanes. A plurality of primary fuel tube assemblies extend through both pre-mixing passages, with each primary fuel tube assembly located between a pair of swirl vanes. Each primary fuel tube assembly is comprised of two tubular members. The first member supplies fuel to the first pre-mixing passage, while the second member, which extends through the first member, supplies fuel to the second pre-mixing passage. An annular fuel manifold is divided into first and second chambers by a circumferentially extending baffle.

Abstract: In a pilot burner fuel nozzle for a premixed type combustor, diffusion pilot fuel injection ports (fuel holes 101) are disposed unevenly in the circumferential direction, whereby a nonuniform pilot flame is formed in the circumferential direction in the premixed type combustor. Thereupon, the heat is distributed in the axial direction, so that oscillated combustion does not occur, by which stable combustion can be maintained.

Abstract: A low emissions combustor includes a premixer for premixing liquid fuel and compressed air for achieving low NOx emissions without water or steam injection. The premixer includes a centerbody disposed in a shroud defining an annular flow channel extending between an inlet and outlet of the shroud. A plurality of fuel injection orifices are spaced circumferentially around the centerbody with each having an outlet being substantially flush with an outer surface of the centerbody. The fuel injection orifices inject liquid fuel into the flow channel wherein it is atomized by compressed air channeled through the shroud inlet. In a preferred embodiment, the fuel injection orifices are inclined at an acute angle for injection the fuel toward the shroud inlet to increase differential mixing velocity with the compressed air.

Abstract: An air fuel mixer is disclosed having a mixing duct, a shroud surrounding the upstream end of the mixing duct in which a fuel manifold is provided in flow communication with a fuel supply and control means, a set of inner and outer counter-rotating swirlers adjacent the upstream end of the mixing duct for imparting swirl to an air stream, a hub separating the inner and outer annular swirlers to allow independent rotation of the air stream, and a centerbody located axially along and substantially the full length of the mixing duct. In order to inject one type of fuel into the mixing duct, fuel is supplied to the outer annular swirlers which include hollow vanes with internal cavities, wherein the internal cavities of the outer swirler vanes are in fluid communication with the fuel manifold in the shroud. The outer swirler vanes further include a plurality of fuel passages therethrough in flow communication with the internal cavities.