Abstract: In a fuel injector assembly, for an internal combustion engine such as a gas turbine engine, a generally cylindrical air transfer sleeve is, in one embodiment, juxtaposed around a shroud, the sleeve having a peripheral inner recessed cavity, the outer end thereof encompassing a plurality of air inlets and the inner end thereof having a plurality of air openings connecting the cavity with tertiary air for a positive air wash, or in another embodiment, is located intermediate a fuel body and the shroud while having such inner radial dimensions so as to maintain first and second fuel/air venting gaps, with the transfer sleeve also having a peripheral outer recessed cavity wherein the inner end thereof encompasses the air inlets, with an outer end thereof being provided with a plurality of air openings, these openings connecting the recessed cavity with the first fuel/air gap downstream of a restrictor.

Abstract: An annular fairing for covering the annular chamber end wall of a turbomachine combustion chamber, and in particular of an airplane turbojet. The fairing presents openings for passing fuel injectors that are supported by the chamber end wall. The fairing is subdivided into a plurality of adjacent sectors, each sector presenting inner and outer fastener edges capable of being fastened on either side of said chamber end wall. Each sector includes a lip on one of its side edges, which lip is connected to the remainder of the sector by a step, said lip being designed to over the side edge of the adjacent sector.

Abstract: Location rings 3, 23 are generally known to present a sealing ring 2, 12 in combustor parts of a gas turbine engine. The location rings 3, 23 act to present the sealing rings 2, 12 such that fuel spray nozzles are appropriately presented towards the combustor whilst these location rings 3, 23 are more susceptible to wear than a panel 4, 24 within which the arrangement 1, 10 is secured. Previously location rings 3 have been located by dedicated bolts 6 provided in drilled and tapped apertures in the panel 4. Each bolt 6 is typically spot welded to provide a lock feature such that upon repair and overhaul it is necessary to remove a heat shield 5 and then remove the spot welds before disassembly. By integrating the heat shield 25 with location ring 23 portions it is possible to utilize the existing heat shield retainer positions as fixing points 19 with associated rail or strap elements 16, 17 such that through overlap along retainer edges 26, 27 a desired retained configuration is achieved.

Abstract: A gas turbine fuel nozzle comprises a body and a sheath adapted to surround the body. A snap-on device is provided for releasably retaining the sheath on the body. The snap-on device is displaceable between a first position for allowing the sheath to be fitted over the body and a second position for retaining the sheath in place about the body.

Abstract: In a fuel lance by means of which fuels can be injected, via at least two separate passages, into a combustion chamber alternately or simultaneously at an injection location arranged substantially at the lance tip, reliable operation is achieved, without the risk of flashbacks and also without coking, by virtue of the fact that the fuel lance, in addition to fuel, also passes purge air to the injection location, and that the purge air, at the injection location, is routed between the two fuel systems, in such a manner that these systems are shielded from one another by the purge air.

Abstract: A fuel-injection device in a turbomachine, comprising a high-pressure pump (10) supplying a flow control valve (16), whose outlet is connected via a pressurization and cut-off valve (20) to a fuel-injector (18) feed pipe (22), the valve (20) being connected to the inlet and to the outlet of the pump (10) in order to define two fuel pressurization thresholds, one being used to start-up and restart the turbomachine, and the other being used for operating the turbomachine as from an idle speed and for commanding an equipment with variable geometry (24).

Abstract: A gas turbine combustor burns a fuel mixture at a high combustion efficiency and a low NOx emission, is simple in construction and exercises improved ignition performance, and an ignition method can efficiently igniting a fuel mixture in the gas turbine combustor. A gas turbine combustor provided with fuel nozzles each having a pilot fuel injection nozzle and a main fuel injection nozzle, and fuel nozzles each having a pilot fuel injection nozzle and a main fuel injection nozzle. The fuel nozzle disposed close to an igniter is provided with a local fuel injection port through which fuel is jetted out from a predetermined position in an air passage in the main fuel injection nozzle to create a combustible fuel mixture zone in the vicinity of the igniter at least while the igniter is in an ignition operation.

Abstract: In a fuel lance by means of which fuels can be injected, via at least two separate passages, into a combustion chamber alternately or simultaneously at an injection location arranged substantially at the lance tip, reliable operation is achieved, without the risk of flashbacks and also without coking, by virtue of the fact that the fuel lance, in addition to fuel, also passes purge air to the injection location, and that the purge air, at the injection location, is routed between the two fuel systems, in such a manner that these systems are shielded from one another by the purge air.

Abstract: A device for fastening a fluid duct such as a fuel duct in an orifice of a turbojet casing, the device comprising a ring screwed into the orifice of the casing and having an annular rim coming to bear against a washer at the end of the duct, and a nut screwed onto the end of the duct so as to clamp the annular rim of the ring against the washer of the duct in order to fasten the duct to the casing.

Abstract: A fuel injector for a gas turbine engine is disclosed which includes an inlet fitting for receiving fuel, a nozzle for issuing fuel into the gas turbine engine, an injector body extending between the inlet fitting and the nozzle and having and interior bore extending therethrough, and an elongated fuel tube disposed within the interior bore of the injector body for delivering fuel from the inlet fitting to the nozzle. The fuel tube has an inlet end portion dynamically associated with the inlet fitting and an outlet end portion rigidly connected to the nozzle. An adapter is operatively associated with either the inlet end section of the fuel tube or a reception bore formed with the inlet fitting of the injector for accommodating a seal member in a manner that permits thermal growth of the injector body relative to the fuel tube.

Abstract: A fuel injector for a gas turbine engine is disclosed wherein the engine has an engine case that includes a reception bore for accommodating the fuel injector, and wherein the fuel injector includes a fuel inlet fitting having an annular mounting flange defining opposed upper and lower end surfaces and a cylindrical body portion which depends axially from the lower end surface of the mounting flange, the inlet fitting having integrally formed alignment structure located beneath the lower end surface of the mounting flange for guiding the fuel injector into an installed position within the reception bore of the engine case.

Abstract: A combustor nozzle is provided. The combustor nozzle includes a first fuel system configured to introduce a syngas fuel into a combustion chamber to enable lean premixed combustion within the combustion chamber and a second fuel system configured to introduce the syngas fuel, or a hydrocarbon fuel, or diluents, or combinations thereof into the combustion chamber to enable diffusion combustion within the combustion chamber.

Abstract: A gas sleeve (120) for a combustor (408) of gas turbine engine (400) attaches to a support housing (100) of the combustor (408) to convey a fuel gas and to fit within a fuel rocket (110). The gas sleeve (120) comprises a plurality of apertures (128) formed to provide impingement cooling. The apertures (128) comprise a tilt angle directed toward a structure in need of impingement cooling, for instance a weld joint (114) that attaches the fuel rocket (110) to the support housing (100). The apertures (128) additionally may comprise a rotational angle effective to create a rotationally swirling flow of the portion of fuel gas that passes through the apertures (128). A method of operation using this structure also is provided.

Abstract: A heat shield for a fuel member of a gas turbine engine having at least one heat shield segment having a connection edge with at least one joint receiving portion defined therealong for receiving a welded joint attaching the heat shield segment to a fuel member, each joint receiving portion including a stress relieving feature defined along at least one end thereof.

Abstract: An aerofoil assembly for the supply of fluid to a gas turbine engine, the aerofoil having a first part with a profile intended to provide a leading edge when located in a gas turbine engine, a second part with a profile intended to provide a trailing edge when located in a gas turbine engine, wherein the first part has features that engage with complementary features on the second part to form a releasable arrangement.

Abstract: An injection assembly for use with a combustor is provided. The injection assembly includes an effusion plate that has a plurality of plate openings and a plate sleeve having a sidewall portion that includes a forward edge. The forward edge is coupled to the effusion plate such that the effusion plate is oriented obliquely with respect to a centerline extending through the combustor. The injection assembly also includes a plurality of ring extensions where each of the ring extensions is coupled to one of the plurality of plate openings. Each ring extension extends rearwardly into the plate sleeve.

Abstract: A pilot nozzle heat shield includes a generally cylindrical body having a first end for receiving a pilot nozzle. The body includes a plurality of radial retention pin cavities for receiving retention pins. A frustoconical flow tip is located at a second end of the body that includes a proximal periphery and a distal periphery. A plurality of flow jets are circumferentially spaced about the proximal periphery of the frustoconical flow tip which further includes a plurality of slots extending distally from the plurality of flow jets. The plurality of slots define a plurality of tangs and the plurality of tangs define an aperture at the distal periphery of the flow tip. At least two of the tangs are connected about the distal periphery of the flow tip and the pilot nozzle heat shield generally includes at least two sets of connected tangs.

Abstract: A fuel injection device for fuel injection systems, e.g. for aircraft gas turbines, includes at least one fuel injection opening, through which a continuous flow of fuel 1 is issued, with the fuel injection opening having a non-circular cross-section.

Abstract: In a fuel injector assembly, for an internal combustion engine, a curved outer housing, fixed at one end, fully encloses a curved flexible fuel feed member, affixed to the housing inlet end and has a nozzle assembly operatively connected to an inner end, wherein the improvement comprises that the housing inlet includes at least one first shaped surface portion, and the nozzle assembly includes a movable nozzle spray-tip having another shaped surface portion that mates conformingly with and is in contact with the at least one shaped surface portion, resulting in relative motion therebetween upon operation of this engine, as a result of the thermal differential arising due to the differing temperatures of the housing and feed member.

Abstract: In a multimode fuel injector for combustion chambers, in particular a jet engine, a secondary circuit is connected to a distribution chamber perforated by a plurality of fuel ejection orifices and the primary circuit comprises at least one passage section adjacent the distribution chamber, for its cooling.

Abstract: An automatic recovery pilot nozzle support system for supporting a pilot nozzle in a fuel injection system of a turbine engine. The support system includes a sleeve extending from an orifice in a support housing, wherein the sleeve has a hollow opening for containing a pilot nozzle. The sleeve may increase the natural frequency of the pilot nozzle above an excitation zone, thereby limiting destructive vibrations. The sleeve may be attached to the support housing with an interference fit. The automatic recovery aspect of the support system enables the sleeve to maintain an interference fit with the support housing such that anytime the sleeve looses contact with the support housing, an insulative film of air forms between the sleeve and the support housing. The insulative film of air causes the temperature of the sleeve to increase, and the sleeve to expand and reform the interference fit with the support housing.

Abstract: The present invention includes a fuel distributor for a fuel nozzle in a gas turbine engine comprising an inner tubular body and an outer tubular body respectively having an outer body inner surface and an inner body outer surface adapted to be in sealing contact one with the other, at least two helical fuel channels defined in at least one of the inner and outer surfaces and being in fluid communication with a fuel inlet, and a channel exit port for each helical fuel channel. The present invention also includes a method of distributing fuel in a fuel nozzle comprising the steps of providing at least two helical channels in the fuel nozzle, each having a channel exit port, providing a fuel inlet cavity in fluid communication with the helical channels, and flowing fuel in the fuel inlet cavity, the helical channels and the channel exit ports.

Abstract: A fuel injection nozzle has a cylindrical nozzle body having a cavity where fuel passes through. A plurality of hollow spokes, each having an aerofoil cross section, are provided around the nozzle body. Each hollow spoke has four fuel injection holes in total on both side surfaces, i.e., two fuel injection holes on each side surface, to inject the fuel, with a distance from the surface of the nozzle body. The inside of each hollow spoke is hollow. The hollow spoke injects the fuel sent to the hollow nozzle body, from the fuel injection holes through the inside of the hollow spoke.

Abstract: The invention relates to power plants, in particular to devices for mixing and atomizing propellant components, and may be used for development injection elements and injector heads of liquid-fuel rocket engines (LRE). The task of the invention is to increase the propellant component combustion efficiency via the increase of propellant component contact area. This task is solved by provision of radial grooves 3 in the outlet portion of a hollow tip 1 of the coaxial-jet injection element, the said hollow tip connecting the cavity of one propellant with the combustion zone, a sleeve 5 surrounding the said tip providing an annular gap 4 connecting the cavity of another propellant with the combustion zone. The sleeve inner surface outlet portion is equidistant relatively to the tip grooves outer surface, and the outlet cross-section area between the sleeve equidistant surface and the tip grooves is Fr=(0,6-2,2)Fo, where Fo is the groove cross-section area at the tip outlet, 2 dependent claims, 3 Figures.

Abstract: A gas turbine fuel nozzle head includes an outlet passage having a curved annular transitional surface extending smoothly to a substantially radial and annular surface relative to a central axis of the fuel nozzle head, for providing a low cost and compact structure with improved cold start and altitude re-light capabilities.

Abstract: In order to allow dual fuel operation, that is to say gas and liquid fuel, a fuel injection arrangement and an assembly are provided in which an injection conduit provides constriction to an airflow towards an outlet. In an assembly a plurality of arrangements including liquid fuel jets are arranged to such that through high angle incidents by the airflow appropriate vaporization and droplet formation is achieved for a fuel/air mixture combustible in a combustor of a turbine engine. At low liquid fuel flow rates below a pre-determined level the liquid fuel is atomized and entrained by the airflow. At high liquid fuel flow rates the fuel splash impinges upon a wash surface to develop a wash film which is again vaporized and atomized about an edge for a fuel/air mixture capable of combustion.

Abstract: A method and a device for the introduction of fuel into a premixing burner, with a pilot gas feed and a premix gas feed for the operation of a gas turbine in the entire load range, in which the pilot gas feed is carried out via a burner lance provided in the premixing burner and the premix gas feed is carried out via side wall shells of the premixing burner. The premix gas feed and the pilot gas feed are carried out in combination, in such a way that a continuous mixture ratio between premix gas and pilot gas can be set within the premixing burner.

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 an atomizing air passageway conduit interconnecting the atomizing air passageway to one of the plurality of passageways to enable fluid flow therebetween, while not allowing the flow of fluid back into the atomizing air passageway. High pressure air from the atomizing air passageway is diverted into one of the plurality of passageways via the conduit to protect the nozzle from ingestion of hot combustor gases, thus eliminating a need for a dedicated air purge system for that one of the plurality of passageways.

Abstract: A unitary swirler assembly that provides improved service life, having a ferrule body and a unitary swirler body. The unitary swirler body has oppositely-directed primary swirler passages and secondary swirler passages. A pair of rails formed on the unitary swirler body engage the ferrule body to limit relative movement, and prevent relative rotation, of the ferrule body and the unitary swirler body.

Abstract: The present invention relates to a premixed fuel injector for low NOx emission and high heating load combustion. The premixed fuel injector of the present invention comprises a center nozzle and a plurality of outer nozzles arranged around the center nozzle, wherein a distance between the adjacent outer nozzles is 15 to 25 times as large as the diameter of the outer nozzle. Further, an air/fuel mixture is supplied to the outer nozzles and fuel is supplied to the center nozzle at a level of 0.1 to 10% of the total amount of fuel supplied to the outer nozzles. According to the premixed fuel injector for high heating load combustion in which nozzles are arranged as described above, there are various advantages in that it's simple, without additional devices, stable flames can be obtained, and NOx emissions are greatly reduced during combustion.

Abstract: A gas turbine combustion chamber with a burner 7, includes means for the supply of fuel and an atomizer 6, wherein the means for the supply of fuel are provided such that the fuel is injected in areas with maximum airflow velocity.

Abstract: In a fuel lance by means of which fuels can be injected, via at least two separate passages, into a combustion chamber alternately or simultaneously at an injection location arranged substantially at the lance tip, reliable operation is achieved, without the risk of flashbacks and also without coking, by virtue of the fact that the fuel lance, in addition to fuel, also passes purge air to the injection location, and that the purge air, at the injection location, is routed between the two fuel systems, in such a manner that these systems are shielded from one another by the purge air.

Abstract: A method enables a gas turbine engine to be assembled. The method comprises coupling a fuel nozzle within the engine to inject fuel into the engine, wherein the fuel nozzle includes three independent injection circuits arranged such that the second injection circuit is between the first and third injection circuits, coupling a liquid fuel source to a first injection circuit defined within the nozzle and including an annular discharge opening, and coupling a water source to one of the second injection circuit and the third injection circuits such that the water source is coupled in flow communication to an annular discharge opening.

Abstract: A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.

Abstract: Disclosed is a device for increasing entrainment and mixing in an air/fuel zone of a direct fuel injection system. The device comprises an ejector nozzle in the form of an inverted funnel whose central axis is aligned along the central axis of a fuel injector jet and whose narrow end is placed just above the jet outlet. It is found that effective ejector performance is achieved when the ejector geometry is adjusted such that it comprises a funnel whose interior surface diverges about 7° to about 9° away from the funnel central axis, wherein the funnel inlet diameter is about 2 to about 3 times the diameter of the injected fuel plume as the fuel plume reaches the ejector inlet, and wherein the funnel length equal to about 1 to about 4 times the ejector inlet diameter. Moreover, the ejector is most effectively disposed at a separation distance away from the fuel jet equal to about 1 to about 2 time the ejector inlet diameter.

Abstract: An injector assembly includes a barrel-shaped housing and an injector, the injector including a feed ring formed of multiple, etched, T-shaped plates. A plurality of nozzles are arranged in an evenly-spaced array around the injector and direct fluid radially inward into the central annulus of the injector assembly. The injector includes an air inlet port with an internal feed passage fluidly connected to swirl chambers and exit orifices to provide individual sprays of fuel from the nozzles. Integral air swirlers impart a swirling component of motion to the fuel sprays. A venturi configuration is provided by the housing. The nozzles are provided along the venturi configuration, which maintains fluid separation from the walls of the housing downstream from the venturi.

Abstract: A premix secondary fuel nozzle for use in transferring a flame from a first combustion chamber to a second combustion chamber is disclosed. The secondary fuel nozzle includes multiple fuel circuits, each of which are fully premixed, and neither of which are injected in a manner to directly initiate or support a pilot flame, thereby lowering emissions. Multiple embodiments are disclosed for alternate configurations of a first fuel injector, including an annular manifold and a plurality of radially extending tubes. Alternate premix secondary fuel nozzles are also disclosed incorporating improved tip cooling schemes that reduce the amount of cooling flow and increase the local heat transfer effectiveness. Reduced cooling flow to the tip region helps to improve flame stability and lower combustion dynamics by eliminating unnecessary cooling air from the fuel nozzle recirculation zone.

Abstract: Systems, methods and apparatus' of converting an engine into a multi-fuel engine are provided. One embodiment reduces particulate emissions and reduces the amount of combusted gasoline or diesel fuel by replacing some of the fuel with a second fuel, such as natural gas, propane, or hydrogen. One feature of the present invention includes a control unit for metering the second fuel. Another feature of the present invention includes an indicator that indicates how much second fuel is being combusted relative to the diesel or gasoline. This Abstract is provided for the sole purpose of complying with the Abstract requirement rules that allow a reader to quickly ascertain the subject matter of the disclosure contained herein. This Abstract is submitted with the explicit understanding that it will not be used to interpret or to limit the scope or the meaning of the claims.

Abstract: The present invention relates to a fuel nozzle for use in an engine such as a gas turbine engine. The fuel nozzle includes a fuel injector for injecting fuel into a combustion chamber of the engine and a plurality of rows of holes surrounding the fuel injector for eliminating recirculation of hot gas products onto a face of the fuel nozzle. The holes eject air primarily in an axial direction.

Abstract: The present invention relates to a burner, substantially comprising a swirl generator for a combustion air stream and means for introducing fuel into the combustion air stream, the swirl generator having combustion-air inlet openings for the combustion air stream that enters the burner, and the means for introducing fuel into the combustion air stream comprising one or more first fuel feeds having a group of first fuel outlet openings, arranged distributed around the burner axis at a combustion chamber-side end of the burner. The burner is distinguished by the fact that the one or more first fuel feeds having the group of first fuel outlet openings are mechanically decoupled from the swirl generator. The present burner allows reliable and safe use of synthesis gas in both dilute and undiluted form as fuel.

Abstract: The present invention provides a widely variable and yet precisely controlled system for metering the flow of fuel to an engine component. The present invention utilizes a plurality of solenoid activated injector valves to meter fuel into an output plenum where the fuel is accumulated to avoid pressure pulses and pressurized if necessary for introduction to the combustion chamber of a gas turbine engine. This system is particularly well suited to high pressure engine systems such as helicopter jet or turbo shaft engines.

Abstract: A fuel injector includes an annular main fuel nozzle received within an annular nozzle housing, a main nozzle fuel circuit having at least one annular leg, and a pilot nozzle fuel circuit. Spray orifices of the leg extend through the fuel nozzle and spray wells through the housing are aligned with the orifices. The nozzle is designed to generate sufficient static pressure differentials between at least two different ones of the spray wells to purge the main nozzle fuel circuit. Spray well portions may be asymmetrically flared out with respect to a spray well centerline in different local streamwise directions. Some of the spray well portions may be asymmetrically flared out in a local upstream direction and others in a local downstream direction. The local streamwise direction may have an axial component parallel to a nozzle axis about which the annular nozzle housing is circumscribed and a circumferential component around the nozzle housing.

Abstract: A method facilitates assembling a gas turbine engine. The method comprises coupling a combustor including a dome assembly and a combustor liner that extends downstream from the dome assembly to a combustor casing that is positioned radially outwardly from the combustor, coupling a fuel injector including a fuel inlet and an air inlet to the combustor casing such that the fuel injector extends axially through the dome assembly such that fuel may be discharged from the primer nozzle into the combustor, and coupling the air inlet to an air source such that cooling air received therethrough is circulated through the fuel injector to facilitate cooling the fuel injector.

Abstract: There is provided a combustor comprising a fuel nozzle which is comprised of a rodlike body which has a fuel passage and which is located in an air passage; a plurality of hollow members which are connected to the fuel passage and which extend in radial directions from the rodlike body into the air passage; at least one injection port formed in each hollow member to inject a fuel from the fuel passage into the air passage; and a projection which extends from a farmost inner wall of each hollow member that is most distant from an axis of the rodlike body to the injection port that is most distant from the axis. A hole for leaking fuel that is connected to an air passage may be formed in a farmost inner wall of the hollow member that is more distant from an axis of a rodlike body than the injection port that is most distant from the axis, or may be formed to be adjacent to all the injection ports on an upstream or downstream side in the direction of the airflow.

Abstract: A fuel injector for gas turbine engines includes an inlet fitting, a housing stem, and a spray nozzle. A pair of fuel passages in the stem deliver fuel in a primary (continuous) flow and a secondary (non-continuous or reduced) flow to primary and secondary discharge orifices in the nozzle. The fuel passages are defined by inner and outer concentric conduits, with the primary flow directed through the inner conduit, and the secondary flow directed through the outer conduit. Fins are provided unitary with the inner conduit along the length of the housing stem, and are fixed at their radial outer ends to the outer conduit. The fins provide structural support and thermally connect the inner and outer conduits. During low power operation, the fins transfer heat energy between the outer conduit to the inner conduit to cool the outer conduit and thermally protect the fuel in the secondary fuel passage.

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 pilot air system for a combustor of a gas turbine including: a pilot air compressor having inlet connectable to in to a main passageway, wherein the main passageway receives compressed air from a compressor for the gas turbine; a pilot air platform positioned adjacent to the combustor of the gas turbine, wherein the pilot air compressor is positioned on the platform; an inline throttling valve coupled to the first main passageway; a by-pass passageway for the pilot air, proximate to the platform, and arranged in parallel to the main passageway and compressor, wherein the by-pass passageway receives pilot air from the main passageway downstream of the pilot air compressor and passes a portion of the compressed pilot air to the main passageway upstream of the pilot air compressor; a by-pass throttling valve inline with the by-pass passageway to meter pilot air flowing through the by-pass passageway, and the main passageway having an outlet connectable to the combustor.

Abstract: A device for the injection of fuel into the combustion chamber of a gas turbine includes at least one swirler (1) arranged in an air path with at least one swirler vane (2) and with at least one fuel injection nozzle (3), wherein the fuel injection nozzle (3) is arranged in a wake area of the swirler vane (2) and is separate from the swirler vane (2).

Abstract: A pilot air system for providing pilot air to a combustor of a gas turbine wherein the system includes: an inlet to receive a portion of compressed air discharged by a compressor of the gas turbine, wherein the portion of the compressed air is pilot air; a main passageway coupled to the inlet and providing a passage for the pilot air; an inline throttling valve coupled to the main passageway and metering a pressure of the compressed air in the main passageway; a pilot air compressor in series with the main passageway; a by-pass passageway for the pilot air and arranged in parallel to the main passageway and compressor, wherein the by-pass passageway receives pilot air from the main passageway downstream of the compressor and passes a portion of the compressed pilot air to the main passageway upstream of the compressor; a by-pass throttling valve inline with the by-pass passageway to meter pilot air flowing through said by-pass passageway, and the main passageway is connectable to the combustor.

Abstract: A fuel injector includes an annular nozzle housing circumscribed about a nozzle axis and an annular fuel nozzle circumscribed about the nozzle axis within the housing. The annular fuel nozzle has at least one main nozzle fuel circuit, a pilot nozzle fuel circuit, and spray orifices extending radially away from the main nozzle fuel circuit through the annular fuel nozzle. Spray wells extend radially through the nozzle housing and are aligned with the spray orifices. An asymmetric cyclone means is located radially outwardly of the housing for generating a swirling flow around the nozzle and an asymmetric static pressure differential around the housing. The asymmetric cyclone means may be a symmetrical radial inflow swirler circumscribed about a cyclone axis that is not collinear with the nozzle axis or may have a plurality of angled flow swirling elements angled with respect to and asymmetrically circumscribed about the cyclone axis.