Class / Patent application number | Description | Number of patent applications / Date published |
060746000 | Plural distinct injectors | 89 |
20080196414 | Strut cavity pilot and fuel injector assembly - A ramjet engine has an inlet, an outlet, and a gas flowpath from the inlet to the outlet. A pilot recess is along the flowpath. A plurality of struts extend within the flowpath. Each strut has first and second side surfaces extending between leading and trailing ends the trailing end sweeping away from the recess in a downstream direction across from the pilot recess. Each strut has first and second distributed fuel outlets positioned to discharge fuel from the first and second side surfaces respectively. | 08-21-2008 |
20080236165 | DUAL-INJECTOR FUEL INJECTOR SYSTEM - A fuel injector system for injecting fuel into a turbomachine combustion chamber, the system comprising first and second fuel injectors wherein the first injector ( | 10-02-2008 |
20080307792 | TURBOMACHINE COMBUSTION CHAMBER WITH HELICAL AIR FLOW - The invention relates to a turbomachine combustion chamber comprising an inner annular wall, an outer annular wall surrounding the inner wall so as to co-operate therewith to define an annular space forming a combustion area, a plurality of fuel injector systems comprising pilot injectors alternating circumferentially with full-throttle injectors, and at least one air admission opening out into the upstream end of the combustion area in a substantially longitudinal direction. The outer wall has a plurality of pilot cavities extending between the two longitudinal ends of the outer wall and extending radially towards thereof, the pilot cavities being fed with air from outside the combustion chamber in a common substantially circumferential direction. Each pilot injector opens out radially into a pilot cavity, and each full-throttle injector opens out radially between two adjacent pilot cavities. | 12-18-2008 |
20090056338 | TURBOMACHINE COMBUSTION CHAMBER WITH HELICAL AIR FLOW - The invention relates to a turbomachine combustion chamber having an inner wall, an outer wall surrounding the inner wall so as to co-operate therewith to define a space forming a combustion area, a transverse wall interconnecting the inner and outer walls, and fuel injection systems. The inner wall has a plurality of inner steps each extending radially towards the outside of the inner wall, the circumferential spacing between two adjacent inner steps defining an inner cavity. The outer wall includes a plurality of outer steps each extending radially towards the inside of the outer wall, the circumferential spacing between two adjacent inner steps defining an outer cavity. At least some of the inner and outer cavities are fed with air from outside the combustion chamber in a common direction that is circumferential, and with fuel in a direction that is radial. | 03-05-2009 |
20090139241 | COMBUSTING SYSTEM, REMODELING METHOD FOR COMBUSTING SYSTEM, AND FUEL INJECTION METHOD FOR COMBUSTING SYSTEM - The present invention improves the reliability of a combusting system including multiple combustors and improves an environmental performance thereof. The invention includes a device that delays the start time of fuel injection from a fuel nozzle of a combustor including a ignitor from the start time of fuel injection from a fuel nozzle of the combustor not including the ignitor. | 06-04-2009 |
20090173076 | Fuel injector - A fuel injector head for a gas turbine engine the head comprising a pilot injector and a main injector located radially outwardly of the pilot injector. A concentric splitter separates the pilot injector from the main injector and has a toroid chamber which is supplied with air in use to generate a toroidal flow which delays mixing of the pilot and main air flows. | 07-09-2009 |
20090249789 | BURNER TUBE PREMIXER AND METHOD FOR MIXING AIR AND GAS IN A GAS TURBINE ENGINE - A burner for use in a gas turbine engine includes a burner tube having an inlet end and an outlet end; a plurality of slots formed in the burner tube and configured to introduce air flows tangentially into the burner tube and impart swirl to the air flows; a plurality of fuel passages extending axially along the burner tube; and a plurality of fuel injection holes provided to each fuel passage. At least one of the fuel injection holes of each fuel passage is configured to inject a fuel flow tangentially into the burner tube between air flows of adjacent slots to form a fuel and air co-flow. A method of mixing air and fuel in a burner of a gas turbine is provided. The burner includes a burner tube including a plurality of slots formed in the burner tube. The method includes introducing air flows tangentially into the burner tube through the slots and imparting swirl to the air flows; and injecting fuel between air flows of adjacent slots to form fuel and air co-flows. This eliminates jet cross flow, which tends to cause flame holding. The tangentially entered fuel and air co-flow layers then flow down axially with quick mixing and dump to a combustor for a stable premixed combustion. | 10-08-2009 |
20090255263 | Gas-turbine burner for a gas turbine with purging mechanism for a fuel nozzle - A gas-turbine burner for a gas turbine includes a fuel nozzle | 10-15-2009 |
20090266080 | OPTIMIZING THE ANGULAR POSITIONING OF A TURBINE NOZZLE AT THE OUTLET FROM A TURBOMACHINE COMBUSTION CHAMBER - A turbomachine including an annular combustion chamber fitted with fuel injectors and nozzle vanes arranged at the outlet from the chamber, the number of nozzle vanes being an integer multiple of the number of fuel injectors, and the head of each injector being situated angularly half-way between the leading edges of two consecutive nozzle vanes, these leading edges being in alignment with primary air holes and/or with dilution air holes. | 10-29-2009 |
20090320482 | Fuel control arrangement - The present invention addresses fuel supply in gas turbine engines. A splitter valve | 12-31-2009 |
20100011771 | COANDA INJECTION SYSTEM FOR AXIALLY STAGED LOW EMISSION COMBUSTORS - The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole. The tertiary coanda nozzles are configured to inject a third fluid comprising air, at least one other diluent, fuel, or combinations thereof to one or more remaining combustion zones among the plurality of combustion zones. | 01-21-2010 |
20100018210 | COMBUSTOR APPARATUS IN A GAS TURBINE ENGINE - A combustion apparatus in a gas turbine engine comprises a combustor shell for receiving air, a fuel injection system associated with the combustor shell, a first fuel supply structure, and a shield structure. The fuel supply structure is in fluid communication with a source of fuel for delivering fuel from the source of fuel to the fuel injection system and comprises a first fuel supply elements including a first section extending along a first path having a component in an axial direction and a second section extending from the first section along a second path having a component in a circumferential direction. The shield structure is associated with at least a portion of the second section of the first fuel supply element. | 01-28-2010 |
20100170254 | LATE LEAN INJECTION FUEL STAGING CONFIGURATIONS - A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel supplied thereto is combustible, a turbine, including rotating turbine blades, into which products of at least the combustion of the first fuel are receivable to power a rotation of the turbine blades, a transition zone, including a second interior in which a second fuel supplied thereto and the products of the combustion of the first fuel are combustible, the transition zone being disposed to fluidly couple the combustor and the turbine to one another, and a plurality of fuel injectors, which are supported by the transition zone and coupled to a fuel circuit, and which are configured to supply the second fuel to the second interior in any one or more of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages. | 07-08-2010 |
20100199676 | FUEL DELIVERY SYSTEM WITH REDUCED HEAT TRANSFER TO FUEL MANIFOLD SEAL - A gas turbine engine fuel supply assembly which includes a fuel manifold mounted to a casing surrounding a combustor of the engine, and a plurality of fuel nozzles mounted to the fuel manifold. A mating interface between the fuel nozzles and the fuel manifold has at least one sealing element, and at least one aperture defined in the fuel nozzle proximate to the mating interface. The aperture is a non fuel conveying passage, and reduces conductive heat transfer to the sealing element by preventing a direct conductive heat transfer path between a hot region of the fuel nozzle and the sealing element. | 08-12-2010 |
20100229560 | Small gas turbine engine with multiple burn zones - A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor. | 09-16-2010 |
20100242482 | METHOD AND SYSTEM FOR REDUCING THE LEVEL OF EMISSIONS GENERATED BY A SYSTEM - An embodiment of the present invention provides a method and system of operating a combustion system that has Lean direct injection (LDI) functionality. The method and system provides a passive cooling system for an injector of the LDI system. The method and system may also provide a means to direct the flow of the fluid exiting the injector of the LDI system. | 09-30-2010 |
20100307159 | FUEL INJECTOR FOR A GAS TURBINE ENGINE - An improved fuel spray nozzle for a gas turbine engine is proposed, in order to address problems associated with the nozzles being wetted with fuel purged from fuel lines upon engine shutdown. The nozzle has a heat shield provided around a fuel discharge orifice, the heat shield incorporating a sliding expansion joint and having a drip collar arranged to cover the expansion joint so as to protect it from being wetted by fuel ejected through the fuel discharge orifice and falling on the heat shield. The fuel spray nozzle is particularly suited to marine or industrial gas turbine engines having a plurality of radially oriented combustion chambers with respective fuel spray nozzles. | 12-09-2010 |
20110000215 | Combustor Can Flow Conditioner - The present application provides a combustor for a gas turbine engine. The combustor may include a combustor can with a number of nozzles therein and a flow conditioner positioned around the combustor can. The flow conditioner may include a number of apertures therein. | 01-06-2011 |
20110072824 | APPARTUS AND METHOD FOR A GAS TURBINE NOZZLE - A nozzle includes an inlet, an outlet, and an axial centerline. A shroud surrounding the axial centerline extends from the inlet to the outlet and defines a circumference. The circumference proximate the inlet is greater than the circumference at a first point downstream of the inlet, and the circumference at the first point downstream of the inlet is less than the circumference at a second point downstream of the first point. A method for supplying a fuel through a nozzle directs a first airflow along a first path and a second airflow along a second path separate from the first path. The method further includes injecting the fuel into at least one of the first path or the second path and accelerating at least one of the first airflow or the second airflow. | 03-31-2011 |
20110083441 | SYSTEM AND METHOD FOR DISTRIBUTING FUEL IN A TURBOMACHINE - A turbomachine includes a compressor, a turbine operatively connected to the compressor, and a combustion assembly fluidly connected between the compressor and the turbine. The combustion assembly includes an end cover including a plurality of fuel circuits, a fuel distributing flange fluidly connected to at least one of the plurality of fuel circuits, a nozzle assembly fluidly linked to the at least one of the plurality of fuel circuits, and a fuel distribution system configured and disposed to deliver fuel to the fuel distributing flange. The fuel distribution system is selectively activated to stage fuel delivery to the at least one of the plurality of fuel circuits at a pressure sufficient to achieve atomization at the nozzle assembly without requiring a supplemental atomization air flow prior to delivering fuel to others of the plurality of fuel circuits at a pressure to achieve atomization without requiring a supplemental atomization air flow. | 04-14-2011 |
20110107765 | COUNTER ROTATED GAS TURBINE FUEL NOZZLES - In certain embodiments, a system includes a gas turbine controller. The gas turbine controller includes a first operational mode enabling fuel flow only through a first plurality of fuel nozzles having a first swirl direction. The gas turbine controller also includes a second operational mode enabling fuel flow only through a second plurality of fuel nozzles having a second swirl direction opposite from the first swirl direction. | 05-12-2011 |
20110162375 | Secondary Combustion Fuel Supply Systems - Systems are provided for supplying fuel to secondary combustion zones within gas turbine engines. In one embodiment, a system includes a transition piece support structure extending from a compressor casing and configured to support a combustor transition piece of a gas turbine engine. A fuel passageway is integrated with the support structure. | 07-07-2011 |
20110162376 | GASIFICATION SYSTEM AND METHOD USING FUEL INJECTORS - A system is provided that comprises a gasifier with an enclosure disposed about a chamber, wherein the enclosure comprises a top wall, a bottom wall, and a side wall between the top and bottom walls. The gasifier also comprises an outlet disposed in the bottom wall, a first injector disposed in the top wall, and a second injector disposed in the side wall, wherein the first and second injectors are configured to inject fuel, oxygen, or a combination thereof, into the chamber. | 07-07-2011 |
20110185735 | GAS TURBINE COMBUSTOR WITH STAGED COMBUSTION - An annular combustor for a gas turbine engine is provided that facilitates staged combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. A method is also provided for operating a gas turbine engine over a power demand range that facilitates staged combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. | 08-04-2011 |
20110197588 | Fuel Injector Nozzle - A fuel injector nozzle is disclosed. The nozzle includes a nozzle body for fluid communication of a liquid fuel to produce a liquid fuel jet and a fluid to produce a fluid jet. The nozzle body includes an adapter comprising a fuel conduit and a fluid conduit. The nozzle body also includes a nozzle tip disposed on the adapter comprising a plurality of fuel outlet conduits that are in fluid communication with the fuel conduits and a plurality of fluid outlet conduits that are in fluid communication with the fluid conduits. | 08-18-2011 |
20110197589 | Fuel Injector Nozzle - A fuel injector nozzle is disclosed. The nozzle includes a nozzle body having a fuel conduit that extends from a fuel inlet through a fuel outlet conduit to a fuel outlet and a fluid conduit that extends from a fluid inlet through a fluid outlet conduit to a fluid outlet. The fuel outlet conduit and fuel outlet configured to produce a liquid fuel jet from the fuel outlet upon introduction of a pressurized liquid fuel into the fuel conduit. The fluid outlet conduit and fluid outlet are configured to produce a liquid fluid jet from the fluid outlet upon introduction of a pressurized liquid fluid into the fluid conduit, wherein the liquid fuel jet and the liquid fluid jet are configured to impact one another and produce a flow stream of atomized fuel. | 08-18-2011 |
20110203285 | BURNER ARRANGEMENT - A burner arrangement is provided. The burner arrangement includes a support and at least two fuel nozzles attached to the support in the direction of flow, with each fuel nozzle including a support-side section which includes a contact surface on the support side with which it rests on a supporting surface of the support, with at least two fuel nozzle tips embodied in one piece extending out from the support-side section in the direction of flow and the support-side contact surface including at least two extension parts projecting in the direction of the support, with the extension parts each embodying a channel through which fuel is fed in each case to the fuel nozzle tips through passages which are arranged in the support-side contact surface of the support-side section. | 08-25-2011 |
20110209481 | Turbine Combustor End Cover - An end cover assembly for a combustor includes a first plate receptive of one of more combustor fuel nozzles and a second plate. One or more intermediate plates are located between the first plate and the second plate and define one or more cavities to distribute fuel to one or more combustor fuel nozzles. The first plate, the second plate, and the one or more intermediate plates are secured in a stack. | 09-01-2011 |
20110225974 | Multiple Zone Pilot For Low Emission Combustion System - A method of operating a combustor includes delivering a primary fuel flow through a plurality of primary fuel nozzles toward a primary combustion zone and combusting the primary fuel flow in the primary combustion zone. A secondary fuel flow is delivered through a secondary fuel nozzle toward a secondary combustion zone and combusted therein. The secondary fuel is located such that the plurality of primary fuel nozzles are arrayed around the secondary fuel nozzle. An outer swirler is located between the plurality of primary fuel nozzles and the secondary fuel nozzle and includes a plurality of outer swirler channels extending therethrough. A flow of swirler fuel is delivered through the plurality of outer swirler channels into the combustor substantially between the primary combustion zone and the secondary combustion zone to stabilize combustion in the primary combustion zone and/or the secondary combustion zone. | 09-22-2011 |
20120031099 | COMBUSTOR ASSEMBLY FOR USE IN A TURBINE ENGINE AND METHODS OF ASSEMBLING SAME - A combustor assembly that includes a combustor liner having a centerline axis and defining a combustion chamber there within. A plurality of fuel nozzles extends through the combustion liner. An annular flowsleeve is coupled radially outward from the combustor liner such that an annular flow path is defined between the flowsleeve and the combustor liner. The flowsleeve includes a forward surface that extends between an upper endwall and a lower endwall. The upper endwall is positioned a first distance from the plurality of fuel nozzles. The lower endwall is positioned a second distance from the plurality of fuel nozzles that is different than the first distance. | 02-09-2012 |
20120055164 | TURBOMACHINE COMBUSTION CHAMBER COMPRISING IMPROVED MEANS OF AIR SUPPLY - Annular combustion chamber ( | 03-08-2012 |
20120073300 | APPARATUS AND METHOD FOR A COMBUSTOR - A combustor includes an end cover and a combustion chamber downstream of the end cover. The combustor further includes nozzles disposed radially in the end cover and a shroud surrounding at least one of the nozzles and extending downstream into the combustion chamber. The shroud includes an inner wall surface and an outer wall surface. A method for operating a combustor includes flowing compressed working fluid through nozzles into a combustion chamber, flowing fuel through each nozzle in a first subset of the nozzles into the combustion chamber, and igniting the fuel from each nozzle in the first subset of nozzles in the combustion chamber. In addition, the method includes extending into the combustion chamber a separate shroud around each nozzle in a second subset of the nozzles and isolating fuel to each nozzle in the second subset of nozzles. | 03-29-2012 |
20120073301 | CRITICAL FLOW NOZZLE FOR CONTROLLING FUEL DISTRIBUTION AND BURNER STABILITY - A fuel delivery system for delivering fuel to a gas turbine engine includes a fixed critical flow nozzle, a first fuel nozzle and a second fuel nozzle. The fixed critical flow nozzle and the variable critical flow nozzle are connected to a fuel source in parallel. The fixed critical flow nozzle has an orifice throat with a fixed effective cross-sectional area. The variable critical flow nozzle has an orifice throat with a variable effective cross-sectional area. The first fuel nozzle is connected to the fixed critical flow nozzle for delivering fuel to the gas turbine engine, and the second fuel nozzle is connected to the variable critical flow nozzle for delivering fuel to the gas turbine engine. | 03-29-2012 |
20120151930 | FUEL ATOMIZATION DUAL ORIFICE FUEL NOZZLE - A pilot fuel injector tip includes concentric primary and secondary pilot fuel nozzles having a circular primary exit axially aft and downstream of an annular secondary exit respectively. A fuel nozzle assembly includes a pilot swirler flowpath section having an annular inwardly tapering conical flowpath section surrounding primary and secondary exits. An inwardly tapering conical wall section radially inwardly bounding flowpath section defines a conical surface. Exits are located at or axially forward or upstream of the conical surface. An annular secondary fuel supply passage in secondary pilot fuel nozzle includes a secondary fuel swirler with an array of helical spin slots that may have rectangular cross sections. A chamfered leading edge of an annular wall section disposed between an outer pilot swirler and an inlet to an injector cooling flowpath surrounding the second pilot swirler includes a radially inwardly facing conical chamfered surface for deflecting dirt from cooling flowpath. | 06-21-2012 |
20120151931 | Distributed Ignition Of Fuels Using Nanoparticles - An engine includes: (1) a housing defining a combustion chamber; (2) a set of injection mechanisms connected to the housing and configured to introduce a fuel and nanoparticles into the combustion chamber; and (3) an optical ignition system connected to the housing and configured to irradiate the nanoparticles to ignite the fuel. | 06-21-2012 |
20120167571 | COMBUSTOR ASSEMBLIES FOR USE IN TURBINE ENGINES AND METHODS OF ASSEMBLING SAME - A method of assembling a combustor assembly for use in a turbine engine is provided. A combustor liner that defines a combustion chamber therein is provided. The liner includes a forward end and an aft end, wherein the aft end includes at least one channel extending therethrough. The channel is aligned obliquely with respect to a centerline extending through the aft end. A plurality of fuel nozzles are coupled to the forward end such that the fuel nozzles extend through the forward end. An annular sleeve that includes at least one opening extending radially therethrough is coupled to the aft end, wherein the sleeve substantially circumscribes the aft end such that a cavity is defined therebetween and such that a fluid channeled through the at least one opening impinges against a surface of the aft end prior to the fluid being channeled into the combustion chamber. | 07-05-2012 |
20120174589 | Combustion Chamber End Cover Without Welding or Brazing - An end cover assembly is provided in combination with a gas turbine combustor for capping a back end of a combustion chamber. The end cover assembly includes an end cover plate securable to the back end of the combustion chamber, and a monolith block secured to the end cover plate. The end cover plate is constructed without welded gas passage plates and without brazed flow passage inserts and includes fuel channels therethrough, where the fuel channels are sized and positioned to interact with fuel nozzle inlet ports. The monolith block includes internal fuel gas passages positioned relative to the fuel channels in the end cover plate to commute fuel gas to the fuel channels of the end cover plate. | 07-12-2012 |
20120279224 | GAS TURBINE ENGINE COMBUSTOR - A gas turbine engine combustor is provided and includes an array of fuel nozzles, a combustion casing assembly disposed about the array of fuel nozzles and an end cap assembly disposed within the combustion casing assembly to define with the combustion casing assembly an axis-symmetric annulus through which fluid travels into each of the fuel nozzles, at least one of the combustion casing assembly and the end cap assembly being formed with lobed, three-dimensional contouring. | 11-08-2012 |
20120291442 | WALL FOR A TURBOMACHINE COMBUSTION CHAMBER INCLUDING AN OPTIMISED ARRANGEMENT OF AIR INLET APERTURES - A rotationally symmetrical wall for an aircraft turbomachine combustion chamber, including primary holes, dilution holes positioned downstream from the primary holes, and a plug installation aperture upstream from the primary holes, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane (D) in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane (D), wherein the other primary holes are distributed equidistantly, at a predefined interval P, and the distance between the two primary holes is greater than the interval P separating two of the other adjacent primary holes. | 11-22-2012 |
20120304654 | COMBUSTION LINER HAVING TURBULATORS - A combustor for a turbine is provided. The combustor includes a plurality of fuel nozzles and a combustion zone is aligned with a combustion process associated with each of the fuel nozzles. A combustion liner includes a plurality of turbulator groups, and each of the turbulator groups has or more individual turbulators. Each of the turbulator groups is aligned with a hot streak caused by the combustion zone associated with the fuel nozzle. Each of the turbulator groups are circumferentially spaced from a neighboring turbulator group. | 12-06-2012 |
20130036741 | MULTIPOINT INJECTORS WITH AUXILIARY STAGE - A multipoint combustion system for a gas turbine engine includes a housing defining a pressure vessel. A master injector is mounted to the housing for injecting fuel along a central axis. A plurality of slave injectors are each disposed outward of the master injector for injecting fuel and air in an ignition plume radially outward of fuel injected through the master injector. The master injector and slave injectors are configured and adapted so the injection plume of the master injector intersects with the ignition plumes of the slave injectors. A primary manifold is included within the pressure vessel for distributing fuel to the slave injectors. An auxiliary manifold is in fluid communication with the auxiliary nozzles of the slave injectors for issuing an auxiliary flow of fuel from the auxiliary nozzles that is separate from the fuel flow of the primary manifold. | 02-14-2013 |
20130042622 | FLOW BALANCING VALVE - A flow balancing valve for a multistage combustor includes a first pressure feedback line, a first burn line, and a first metering port fluidly connected to the first fuel injector. The flow balancing valve further includes a second pressure feedback line, a second burn line, and a second metering port fluidly connected to the second fuel injector. A metering land is located between and defines sizes of the first metering port and the second metering port. An increase in pressure differential between the first pressure feedback line and the second pressure feedback line causes a compensatory movement in the metering land to balance fuel flow for the first fuel injector and the second fuel injector. | 02-21-2013 |
20130055720 | INTERFACE RING FOR GAS TURBINE NOZZLE ASSEMBLIES - A gas turbine combustor assembly including a combustor liner and a plurality of fuel nozzle assemblies arranged in an annular array extending within the combustor liner. The fuel nozzle assemblies each include fuel nozzle body integral with a swirler assembly, and the swirler assemblies each a bellmouth structure to turn air radially inwardly for passage into the swirler assemblies. A radially outer removed portion of each of the bellmouth structures define a periphery diameter spaced from an inner surface of the combustor liner, and an interface ring is provided extending between the combustor liner and the removed portions of the bellmouth structures at the periphery diameter. | 03-07-2013 |
20130067921 | FUEL INJECTOR - A fuel injector is provided and includes a member defining a flowpath through which a first fluid flows, the flowpath having a cross-section with transverse elongate and short axes, a head defining a plenum storing a supply of a second fluid and a system fluidly coupled to the flowpath and the plenum to inject the second fluid from the plenum and into the flowpath at first and second locations along the elongate axis. The injected second fluid is formed into jets at the first and second locations, the first fluid entrains the jets such that the injected second fluid flows through the flowpath and mixes with the first fluid, and the short axis has a sufficient dimension such that the jets remain spaced from a sidewall of the member. | 03-21-2013 |
20130074505 | SYSTEM FOR DIRECTING AIRFLOW INTO A COMBUSTOR - A system includes a gas turbine combustor. The gas turbine combustor includes a combustion liner disposed about a combustion region and a sleeve disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. The gas turbine combustor also includes multiple axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor. The multiple axial injectors are asymmetrically configured to provide an uniform injection of the airflow circumferentially about an axis of the gas turbine combustor. | 03-28-2013 |
20130074506 | TURBINE BURNER - A turbine burner is provided. The turbine burner has a secondary feed unit and a primary feed unit. The primary feed unit has a primary mixing tube and a fuel nozzle that are arranged concentrically around the secondary feed unit. The primary mixing tube and the fuel nozzle have a fluid flow connection. The fuel nozzle has an annular wall that is radially spaced in the axial direction from the secondary feed unit such that a gap height is fainted by the annular wall and the secondary feed unit. The annular wall has an inside wall directed toward the secondary feed unit and having blades with a leading edge on the upstream side. The fuel nozzle has an inlet and the blades have an axial distance from the inlet. The ratio of the distance to the gap height is greater than 1 and less than the gap height. | 03-28-2013 |
20130081397 | FORWARD CASING WITH A CIRCUMFERENTIAL SLOPED SURFACE AND A COMBUSTOR ASSEMBLY INCLUDING SAME - A forward casing used in a combustor assembly for a turbine engine includes a circumferential sloped surface which reduces the total interior volume within the forward casing. The circumferential sloped surface can be flat or concave shaped. | 04-04-2013 |
20130086912 | SYSTEM FOR COOLING A MULTI-TUBE FUEL NOZZLE - A system includes a multi-tube fuel nozzle including a fuel nozzle head that includes an outer wall surrounding a chamber. The outer wall includes a downstream wall portion configured to face a combustion region. The multi-tube fuel nozzle also includes multiple tubes extending through the chamber to the downstream wall portion. Each tube of the multiple tubes includes an upstream portion, a downstream portion, and at least one fuel inlet disposed at the upstream portion, and is configured to receive air and mix the air with fuel from the at least one fuel inlet. The multi-tube fuel nozzle includes a fuel conduit extending through the chamber crosswise to and around the multiple tubes. The fuel conduit includes multiple impingement cooling orifices. A fuel flow path extends through the fuel conduit, through the impingement cooling orifices, through the chamber, and into the at least one fuel inlet of each tube. | 04-11-2013 |
20130086913 | TURBOMACHINE COMBUSTOR ASSEMBLY INCLUDING A COMBUSTION DYNAMICS MITIGATION SYSTEM - A turbomachine combustor includes a combustor cap having a cap surface and a wall that extends about the cap surface to define a cap volume, and a plurality of nozzle members that extend from the cap surface. The plurality of nozzle members include a center nozzle member and one ore more outer nozzle members. A combustor dynamics mitigation system is arranged in the combustor cap and includes plurality of divider members that extend from the wall toward the center nozzle member. The plurality of divider members define a plurality of parallel resonator volumes. The combustor dynamics mitigation system also includes a plurality of tubes that extend into corresponding ones of the plurality of parallel resonator volumes. | 04-11-2013 |
20130199191 | FUEL INJECTOR WITH INCREASED FEED AREA - Provided is a nozzle tip assembly having at least one tube member housed within a tubular shell, the tube member having a first passage connecting the first outlet port to the first inlet port and a second passage connecting the second outlet port to the second inlet port. The configuration of the tube member allows the tubular shell to be sized such that separation of fluid flowing around the tubular shell is reduced, thereby reducing the amount of aerodynamic wake that occurs in an injector. | 08-08-2013 |
20130263604 | SYSTEM AND METHOD FOR SUPPORTING FUEL NOZZLES INSIDE A COMBUSTOR - A system for supporting fuel nozzles inside a combustor includes a ring that circumferentially surrounds the fuel nozzles inside the combustor, a support plate that extends radially inside at least a portion of the ring, and a first connection between the support plate and at least one of the fuel nozzles inside the combustor. A second connection is between the support plate and the ring. A method for supporting fuel nozzles in a combustor includes surrounding the fuel nozzles with a ring, connecting a support plate to the ring, and connecting the support plate to at least one fuel nozzle. | 10-10-2013 |
20130305725 | FUEL NOZZLE CAP - Certain embodiments include a first individual sector having a first outer perimeter configured to fit together with outer perimeters of a plurality of individual sectors to form a combustor cap assembly of a turbine combustor, wherein the first individual sector comprises a first inner perimeter configured to fit about a nozzle outer perimeter of a first fuel nozzle, and the first individual sector is configured to fixedly attach to the first fuel nozzle. | 11-21-2013 |
20130305726 | INJECTOR FOR THE COMBUSTION CHAMBER OF A GAS TURBINE HAVING A DUAL FUEL CIRCUIT, AND COMBUSTION CHAMBER PROVIDED WITH AT LEAST ONE SUCH INJECTOR - A starting injector usable in all flight modes without additional cost or weight. The starting injector includes a dual fuel circuit and an air circuit. An injector for a combustion chamber of a gas turbine includes a dual fuel injection circuit including a starting fuel circuit for ignition and then for all the flight modes, and a main fuel circuit for all the flight modes after starting. The circuits include parallel pipes in a common tube having an axis. The pipe of the starting circuit is substantially in communication with a center of a spherical injector body. At the end, the pipe accommodates an injection manifold coupled to a central channel passing through a central wall of a swirler. The pipe of the main circuit is in communication with an annular channel opposite jet channels. An air circuit is guided between two portions shaped as concentric spheres. | 11-21-2013 |
20140026579 | BURNER FOR A GAS TURBINE - A burner for a gas turbine includes a burner housing, and a pilot combustor having a supply module providing pilot fuel and pilot air into a pilot combustion room being enclosed by a pilot combustor housing having a tapered exit with a throat of a defined length into the resulting main flow direction. The throat discharges a concentration of radicals are and heat generated in the pilot combustion room into a main combustion room enclosed by the burner housing. The interior cross section area of the throat deviates from a circle by means of flow guiding elements provided as protrusions with a defined radial height or as recesses with a defined depth extending longitudinally along the direction of the burner axis to give the discharging flow a defined velocity distribution with regard to a circumferential direction. | 01-30-2014 |
20140083105 | GAS TURBINE COMBUSTOR - An annular type gas turbine combustor having a plurality of fuel nozzle assemblies ( | 03-27-2014 |
20140157781 | FUEL INJECTOR AND A GAS TURBINE ENGINE COMBUSTION CHAMBER - A gas turbine engine lean burn fuel injector includes a fuel injector head which has a first air swirler, a second air swirler arranged around the first air swirler, a pilot fuel injector arranged radially between the first air swirler and the second air swirler. A third air swirler arranged around the second air swirler, a fourth air swirler arranged around the third air swirler and a main fuel injector arranged radially between the third air swirler and the fourth air swirler. A shroud is arranged around the fourth air swirler. A downstream end of the shroud is generally circular in cross-section in a plane perpendicular to the axis (Y) of the fuel injector head and an upstream end of the shroud is generally elliptical in cross-section in a plane perpendicular to the axis (Y) of the fuel injector head. | 06-12-2014 |
20140190170 | FUEL INJECTOR FOR SUPPLYING FUEL TO A COMBUSTOR - A fuel injector for a combustor generally includes an annular outer body having an inlet and an outlet. The outer body at least partially defines an outer flow passage. An inner flow passage extends at least partially through the outer flow passage and a radial swirler is disposed at the inlet of the outer body. The radial swirler includes a first radial passage separated from a second radial passage. The first radial passage has a first plurality of swirler vanes and the second radial passage has a second plurality of swirler vanes. The first radial passage is in fluid communication with the outer flow passage and the second radial passage is in fluid communication with the inner flow passage. | 07-10-2014 |
20140202161 | COMBUSTOR AND ROTATING MACHINE - A combustor includes: a combustor basket configured to surround a fuel nozzle from an outer circumferential side, and a plurality of connecting members formed in a circumferential direction at intervals to connect a rear end of the combustor basket and a casing and configured to define a flow path through which compressed air introduced into the combustor basket flows. A circulation direction of the compressed air flowing through the flow path is configured to be reversed at a rear end of the combustor basket and the compressed air is introduced into the fuel nozzle. The flow path is partially or entirely inclined in the circumferential direction to blow the compressed air. | 07-24-2014 |
20140202162 | FUEL EFFICIENT ULTRA-LOW EMISSION AND IMPROVED PATTERN FACTOR COLORLESS DISTRIBUTED COMBUSTION FOR STATIONARY AND PROPULSION GAS TURBINE APPLICATIONS - Colorless distributed combustion (CDC) reactors or green combustion gas turbine combustors having a combustion chamber are presented for improved performance of gas turbine combustion engines. The combustors are configured and designed for providing a superior pattern factor (uniform thermal field in the combustion zone) and a reduction or complete elimination of pollutants emission from the combustor (i.e., zero emission gas turbine combustor) and uniform thermal field in the entire combustion zone to provide significantly improved pattern factor. Colorless distributed combustion is achieved with fuel and air entering the combustion chamber via one or more injection ports as non-premixed, or premixed. Rectangular, cylindrical, stadium and elliptical shaped combustors are presented with injection ports and exit ports located in various locations of the combustors. The mixture preparation between fuel and air with the hot combustion products is carried out either with the gases present in the combustion chamber or via a communication link between the exit gases from the combustor back to the combustion chamber. | 07-24-2014 |
20140245741 | STATOR VANE ROW - In a gas turbine engine, each vane has pressure and suction surfaces extending radially from an inner to outer endwall of an annular working gas engine passage, and extending axially from a leading to a trailing edge of the vane. Each vane has transverse sections providing respective aerofoil sections. Neighbouring vanes are arranged in unequally-shaped pairs in which either: (i) the first vane of each pair exhibits compound lean, and the second vane of the pair exhibits reverse compound lean or has substantially no tangential lean, (ii) the first vane of each pair has substantially no tangential lean, and the second vane of the pair exhibits reverse compound lean, or (iii) the first vane of each pair exhibits reverse compound lean, and the second vane of the pair exhibits greater reverse compound lean. Within each unequally-shaped pair the first vane is on the pressure surface side of the second vane. | 09-04-2014 |
20140260276 | END COVER CONFIGURATION AND ASSEMBLY - A system includes an end cover for a multi-tube fuel nozzle. The end cover includes a first side, a second side disposed opposite the first side, a plurality of fuel injectors disposed on the first side, and at least one pre-orifice disposed within a passage within the end cover between the first and second sides. The pre-orifice is configured to be removed through the end cover from the second side. | 09-18-2014 |
20140260277 | FLOW SLEEVE FOR A COMBUSTION MODULE OF A GAS TURBINE - A combustion module for a combustor of a gas turbine includes an annular fuel distribution manifold disposed at an upstream end of the combustion module. The fuel distribution manifold includes an annular support sleeve having an inner surface. The combustion module further includes a fuel injection assembly having an annular combustion liner that extends downstream from the fuel distribution manifold and that terminates at an aft frame, and an annular flow sleeve that circumferentially surrounds the combustion liner. The flow sleeve extends downstream from the fuel distribution manifold and terminates at the aft frame. The flow sleeve extends continuously between the support sleeve and the aft frame. A forward portion of the flow sleeve is positioned concentrically within the support sleeve where the forward portion is slidingly engaged with the inner surface of the support sleeve. | 09-18-2014 |
20140331678 | SYSTEM FOR DISTRIBUTING COMPRESSED AIR IN A COMBUSTOR - A system for distributing compressed air in a combustor of a gas turbine engine. The system may include a flow splitter, a center duct, an outer duct, and an inner duct configured to separate and receive compressed air from a prediffuser exit, and to route separate flows of the compressed air to separate downstream locations for the combustion reaction and for cooling. The system may include an axial mixer configured to axially receive a flow of the compressed air and to direct the flow to mix with fuel provided by an injector. | 11-13-2014 |
20140366543 | COMBUSTION EQUIPMENT FOR USE IN A GAS TURBINE ENGINE - Combustion equipment for use in a gas turbine engine is provided. The combustion equipment includes an inner and outer case, wherein the outer case encloses the inner case creating a cavity located between the two cases. The combustion equipment also has a fuel delivery device for delivering fuel into a combustion region within the inner case, wherein the fuel delivery device passes through a hole in the outer case and a corresponding hole in the inner case. The combustion equipment also includes a seal assembly to prevent the leaking of pressurized gas into the cavity located between the outer and inner cases via the hole through which fuel delivery device passes. The seal assembly includes a flexible sleeve that extends around the fuel delivery device and interconnects the inner case with the outer case, permitting relative movement of the inner and outer cases when the combustion equipment is in use. | 12-18-2014 |
20150020528 | FUEL MANIFOLD AND FUEL INJECTOR ARRANGEMENT FOR A COMBUSTION CHAMBER - A fuel manifold and fuel injector arrangement for a gas turbine engine combustion chamber includes an annular combustion chamber casing arranged around the combustion chamber. Circumferentially spaced fuel injectors supply fuel into the combustion chamber and a fuel manifold supplies fuel to each of the fuel injectors. The fuel manifold includes flexible fuel pipes and a plurality of T piece connectors. The stem of each connector is mounted on an outer end of a respective one of the fuel feed arms and each flexible fuel pipe interconnects an arm of one connector with an arm of an adjacent connector. Each connector is arranged such that the arms are arranged are at angles to the planes containing the axis of the casing and perpendicular to the axis of the casing respectively. Adjacent connectors are arranged at opposite angles such that the fuel manifold extends around the casing sinusoidally. | 01-22-2015 |
20150047360 | SYSTEM FOR INJECTING A LIQUID FUEL INTO A COMBUSTION GAS FLOW FIELD - A system for injecting a liquid fuel into a combustion gas flow field includes an annular liner that defines a combustion gas flow path. The annular liner includes an inner wall, an outer wall and a fuel injector opening that extends through the inner wall and the outer wall. The system further includes a gas fuel injector that is coaxially aligned with the fuel injector opening. The gas fuel injector includes an upstream end and a downstream end. The downstream end terminates substantially adjacent to the inner wall. A dilution air passage is at least partially defined by the gas fuel injector. A liquid fuel injector extends partially through the dilution air passage. The liquid fuel injector includes an injection end that terminates upstream from the inner wall. | 02-19-2015 |
20150047361 | NOZZLE WITH MULTI-TUBE FUEL PASSAGEWAY FOR GAS TURBINE ENGINES - A pilot fuel nozzle for a combustor includes an igniter forming a central body extending along a longitudinal center of the nozzle. A nozzle tip includes a plurality of circumferentially spaced fuel passages and a plurality of circumferentially spaced air passages extending to an outer side of the nozzle tip. The central body extends through a center of the nozzle tip for producing a spark to ignite a fuel/air mixture adjacent to the nozzle tip. A plurality of fuel tubes extend along the central body, each of the fuel tubes having an outlet end engaged on the nozzle tip for delivery of fuel from the nozzle tip into a combustion chamber of the combustor An outer sleeve surrounds the fuel tubes and defines an annular space in fluid communication with the air passages of the nozzle tip between the outer sleeve and the central body | 02-19-2015 |
20150052899 | AIRBLAST FUEL INJECTOR - An airblast fuel injector for a gas turbine engine fuel spray nozzle has, in order from radially inner to outer, a coaxial arrangement of an inner air swirler passage, an annular fuel passage, an annular outer air swirler passage, and an annular shroud air swirler passage. The injector further has an annular shroud having an inner surface profile. Relative to the overall axial direction of flow through the injector the shroud inner surface profile has a convergent section followed by a divergent section, the transition of which forming a first inwardly directed annular nose. The injector further has an annular wall having an outer surface profile, and having an inner surface profile. Relative to the overall axial direction of flow through the injector the wall outer surface profile has a convergent section followed by an outwardly turning section which faces across the shroud air passage to the first nose. | 02-26-2015 |
20150082796 | BURNER - A burner of a gas turbine extending along an axis is provided having in axial order: a swirler section, mixing section, outlet section, and main combustion zone. The swirler section has swirler vanes to swirl a stream of fuel and oxygen containing gas entering therein in a circumferential direction. The mixing section conducts the premix of fuel and oxygen containing gas to the outlet section. The outlet section discharges the premix into the combustion zone expanding the flow of premix from a smaller axial cross section of the mixing section to a larger cross section of the combustion zone which streamlines the flow to diverge radially. A surface of the outlet section facing the flow of the premix is provided with first fuel nozzles injecting fuel into the premix into a radial inwardly inclined direction before the flow of the premix enters the outlet section into the combustion zone. | 03-26-2015 |
20150082797 | FUEL INJECTION DEVICE - A fuel injection device for supplying a fuel to a compressed air includes: a pilot fuel injector; a main fuel injector located at an outer periphery of the pilot fuel injector; and an air injection unit located between an outlet end portion of the pilot fuel injector and the main fuel injector. The air injection unit includes: a partition wall plate separating the air injection unit from a combustion chamber; a flame stabilization plate provided at a downstream side of the partition wall plate; a first opening at a downstream side between the flame stabilization plate and the outlet end portion of the pilot fuel injector; and a second opening provided in the partition wall plate and through which the compressed air is supplied. The flame stabilization plate includes an inclined portion inclined in a radially outward and downstream direction with respect to an axis. | 03-26-2015 |
20150323184 | ULTRA COMPACT COMBUSTOR - Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle. | 11-12-2015 |
20150369489 | TURBO MACHINE COMBUSTION ASSEMBLY COMPRISING AN IMPROVED FUEL SUPPLY CIRCUIT - Turbo machine combustion assembly ( | 12-24-2015 |
20160097539 | COMBUSTION CHAMBER AND A METHOD OF MIXING FUEL AND AIR IN A COMBUSTION CHAMBER - A combustion chamber including a first fuel injector and a second fuel injector, the first and second fuel injectors being arranged to inject fuel into a mainstream flow of air with the second fuel injector arranged downstream of the first fuel injector. A method of mixing fuel and air in a combustion chamber, including injecting fuel into a mainstream flow of air with a first fuel injector; injecting fuel into the mainstream flow of air with a second fuel injector, which is arranged downstream of the first fuel injector; injecting fuel into the mainstream flow with the first fuel injector such that the resulting mixture between the first and second fuel injectors has an equivalence ratio less than the lean flame stability limit; and injecting fuel into the mainstream flow with the second fuel injector such that a combustion zone is provided downstream of the second fuel injector. | 04-07-2016 |
20170234542 | FLEXIBLE FUEL COMBUSTION SYSTEM FOR TURBINE ENGINES | 08-17-2017 |
060747000 | Injectors in distinct radially spaced parallel flow combustion products generators arranged to combine discharges | 15 |
20090038311 | Outer Sidewall Retention Scheme For A Singlet First Stage Nozzle - An outer sidewall retention scheme for a singlet first stage nozzle of a gas turbine. The retention scheme includes a circumferential retaining ring with a main body and a pair of circumferential retaining lands projecting inward radially. A circumferential annular retaining groove is formed between the pair of circumferential retaining lands. A first lug and a second lug mounted on an outer face of the outer sidewall of each nozzle are adapted to fit within the circumferential annular retaining groove of the retaining ring and are supported radially and circumferentially by a first retaining pin and a second retaining pin, each pin passing though the circumferential retaining lands. Each nozzle further includes a chordal hinge rail and seal on the outer sidewall and a chordal hinge rail and seal on the inner sidewall providing axial support for the nozzle. | 02-12-2009 |
20090071159 | Secondary Fuel Delivery System - A secondary fuel delivery system for delivering a secondary stream of fuel and/or diluent to a secondary combustion zone located in the transition piece of a combustion engine, downstream of the engine primary combustion region is disclosed. The system includes a manifold formed integral to, and surrounding a portion of, the transition piece, a manifold inlet port, and a collection of injection nozzles. A flowsleeve augments fuel/diluent flow velocity and improves the system cooling effectiveness. Passive cooling elements, including effusion cooling holes located within the transition boundary and thermal-stress-dissipating gaps that resist thermal stress accumulation, provide supplemental heat dissipation in key areas. The system delivers a secondary fuel/diluent mixture to a secondary combustion zone located along the length of the transition piece, while reducing the impact of elevated vibration levels found within the transition piece and avoiding the heat dissipation difficulties often associated with traditional vibration reduction methods. | 03-19-2009 |
20090241548 | GAS TURBINE ENGINE COMBUSTOR CIRCUMFERENTIAL ACOUSTIC REDUCTION USING FLAME TEMPERATURE NONUNIFORMITIES - A gas turbine engine combustor includes an annulus with one or more circular rows of burners and a means for providing a number of equiangular spaced apart flame temperature nonuniformities around the annulus during engine operation. The number of the flame temperature nonuniformities being equal to a circumferential acoustic mode to be attenuated in the combustor (i.e three, five, or seven). Fuel lines and/or water lines in supply communication with the burners and metering orifices in a portion of the fuel lines and/or the water lines may be used to produce the flame temperature nonuniformities. The annulus of the burners may have an equal number of equiangular spaced apart first and second arcuate segments of the burners and a means for operating the burners in the first segments and operating the burners in the second segments at different first and second flame temperatures respectively. | 10-01-2009 |
20090255264 | FUEL NOZZLE - A fuel distributor is disclosed comprising a body having a unitary construction, a fuel conduit located within the body, a fuel flow path located within the body that is oriented in a circumferential direction around an axis and in flow communication with the fuel conduit, and at least one orifice located in the body in flow communication with the fuel flow path such that a fuel entering the fuel conduit exits through the orifice. | 10-15-2009 |
20090320483 | Variable Orifice Plug for Turbine Fuel Nozzle - A fuel nozzle for use about an end cap of a combustor. The fuel nozzle may include a flange attached to the end cap, an aperture extending through the flange, a plug positioned in the aperture, and an orifice positioned within the plug. | 12-31-2009 |
20100058766 | Segmented Combustor Cap - A combustor cap is cooperable with an impingement plate in a turbine fuel nozzle and includes a plurality of cap segments independently securable to the impingement plate. The plurality of cap segments are radially and tangentially movable relative to the impingement plate. | 03-11-2010 |
20100162712 | QUENCH JET ARRANGEMENT FOR ANNULAR RICH-QUENCH-LEAN GAS TURBINE COMBUSTORS - A combustor for a turbine engine includes an outer liner having a first group of air admission holes and defining a plurality of outer liner regions. The combustor further includes an inner liner circumscribed by the outer liner and forming a combustion chamber therebetween, the inner liner having a second group of air admission holes and defining a plurality of inner liner regions. The combustor further includes a plurality of fuel injectors extending into the combustion chamber and configured to deliver an air-fuel mixture to the combustion chamber, each of the plurality of fuel injectors being associated with one of the outer liner regions and one of the inner liner regions. The first group within a respective outer liner region includes air admission holes that circumferentially alternate between approximately a first size and approximately a second size, the first size being different than the second size. | 07-01-2010 |
20100229561 | AT LEAST ONE COMBUSTION APPARATUS AND DUCT STRUCTURE FOR A GAS TURBINE ENGINE - A combustion apparatus and duct structure for a gas turbine engine are provided. The combustion apparatus comprises a combustion system to receive fuel and air, ignite at least a portion of the fuel and air and output a stream of first combustion products and any remaining fuel and air. The combustion apparatus further comprises structure positioned adjacent to the combustion system for receiving and accelerating the first combustion products and any remaining fuel and air from the combustion system. The duct structure receives the first combustion products and any remaining fuel and air from the combustion apparatus, allows any remaining fuel and air to combust to generate second combustion products, accelerates the first and second combustion products and outputs the first and second combustion products to a first row blade assembly. | 09-16-2010 |
20110067404 | Universal Multi-Nozzle Combustion System and Method - A gas turbine combustion system includes a supporting fuel manifold including a plurality of fuel nozzle support openings and a plurality of fuel line passages, and an aft case defining at least part of a combustion zone. The supporting fuel manifold is connected to the aft case. A plurality of fuel nozzles are supported one each in the plurality of fuel nozzle support openings, and a plurality of fuel lines communicate with the plurality of fuel line passages. The plurality of fuel lines deliver fuel to the plurality of fuel nozzles via the plurality of fuel line passages. | 03-24-2011 |
20110296840 | GAS TURBINE COMBUSTOR INJECTION ASSEMBLY, AND COMBUSTOR FUEL MIXTURE FEED METHOD - A fuel mixture is fed to a gas turbine combustor by an injection assembly, which has an outer body with combustion-supporting air inlets; a conical tubular portion housed inside the outer body and partly defining an inner conduit and an outer annular conduit; and a first and second feed circuit for feeding liquid fuel to the inner conduit and outer annular conduit respectively; the first circuit having a ring of conduits with respective axes parallel to a generating line of an outer surface of the conical tubular portion. | 12-08-2011 |
20120227408 | SYSTEMS AND METHODS OF PRESSURE DROP CONTROL IN FLUID CIRCUITS THROUGH SWIRLING FLOW MITIGATION - An injector with swirling flow mitigation includes an injector body defining a longitudinal axis. A fluid circuit is defined in the injector body and includes a plurality of flow channels defined in a cylindrical region around the longitudinal axis and being in fluid communication with an outlet orifice for passage of fluids out from the flow channels into a radial direction with respect to the longitudinal axis. A flow splitter is defined in each of the flow channels proximate the outlet orifice. Each flow splitter is configured and adapted to mitigate formation of swirling flow on fluids passing through the outlet orifice from the flow channels. | 09-13-2012 |
20130180251 | BURNER ARRANGEMENT - A burner arrangement includes a carrier and at least two burners which are mounted on the carrier in a flow direction Each burner includes a cylindrical housing having a lance which is arranged centrally therein and having a fuel duct and which is supported on the housing via swirl blades. An attachment is arranged on the side leading to a combustion chamber At least one fuel nozzle is disposed in the attachment and is connected to the fuel duct. The at least two fuel nozzles of the at least two burners have a different functional characteristic and/or spray form, and the at least two fuel nozzles of the at least two burners with a different functional characteristic and/or spray form include at least one full jet nozzle and at least one pressure swirl nozzle. | 07-18-2013 |
20130239575 | SYSTEM FOR SUPPLYING A WORKING FLUID TO A COMBUSTOR - A system for supplying a working fluid to a combustor includes a combustion chamber and a flow sleeve that circumferentially surrounds at least a portion of the combustion chamber. A tube provides fluid communication for the working fluid to flow through the flow sleeve and into the combustion chamber, wherein the tube comprises an axial centerline. A first set of injectors are circumferentially arranged around the tube and angled radially with respect to the axial centerline of the tube, wherein the first set of injectors provide fluid communication for the working fluid to flow through a wall of the tube. | 09-19-2013 |
20140338344 | System Having a Multi-Tube Fuel Nozzle with a Fuel Nozzle Housing - A system including a plurality of multi-tube fuel nozzles each having a plurality of tubes extending in an axial direction, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, and a fuel nozzle housing, including an outer wall extending circumferentially about a central axis, a plurality of radial walls extending from the outer wall inwardly toward the central axis, a plurality of fuel nozzle receptacles disposed within the outer wall, wherein the plurality of radial walls separate the plurality of fuel nozzle receptacles from one another, and the plurality of multi-tube fuel nozzles are disposed in the plurality of fuel nozzle receptacles a mounting structure including a plurality of radial support arms extending outwardly from the outer wall. | 11-20-2014 |
20150128601 | GAS TURBINE COMBUSTOR - A gas turbine combustor having a burner including a plurality of fuel nozzles for injecting fuel, air hole plates positioned on a downstream side of the fuel nozzles and a plurality of air holes arranged in pairs with each of the fuel nozzles, and a combustion chamber for mixing fuel injected from the fuel nozzles and air injected from the air holes and injecting and burning the mixed fuel. Each of the fuel nozzles configuring the burners is provided with a projection in which a part of an outer edge of a section of the fuel nozzle is protruded outward; and the projection is arranged so as to be directed toward a center of the gas turbine combustor. The projection of the fuel nozzle is positioned on a downstream side of a flow of combustion air flowing around each of the fuel nozzles. | 05-14-2015 |