Abstract: A fuel system interface for a gas turbine engine prevents a rotor from over-speeding and receives either an electrically-originated or a mechanically-originated over-speed signal. In the exemplary embodiment, the fuel system interface is coupled to an gas turbine engine fuel shut off valve and to a fuel metering valve to shut off engine fuel after receiving an over-speed condition. As a result, because the fuel system interface acts on both the shut off valve and the fuel metering valve, rotor over-speed protection is facilitated with parallel redundant fuel cutoff devices.

Abstract: An improved torch igniter for use in devices such as thrust augmenters, gas turbine engines, ramjets, combined-cycle engines and industrial burners. The torch igniter includes a housing with a combustion chamber. Fuel and oxidizer are delivered into the combustion chamber and ignited by an electronic ignition source, such as a plasma jet igniter or a spark igniter, so that an upstream recirculation zone and a downstream recirculation zone are created. The upstream recirculation zone stabilizes and pilots combustion within the combustion chamber, while the downstream recirculation zone augments the combustion event. Byproducts of the combustion event within the torch igniter provide a high mass flux with high thermal energy and strong ignition source radicals that are discharged through a neck portion of the housing and are thereafter employed to initiate a primary combustion event in a primary combustor.

Abstract: A premix burner has a premix burner casing which is formed in the manner of a tube open at the upstream end, is connected in the downstream direction to a combustion chamber via a transition contour and through which air can flow, a burner lance including an inner tube that projects into the interior of the premix burner casing on the upstream side, encloses a flow passage which is annular in cross section together with the premix burner casing, and has an inner tube wall which surrounds an inner flow passage and in which there is at least one fuel-addition unit for feeding fuel into the inner flow passage and at least one further fuel-addition unit for feeding fuel into the annular flow passage is provided so that downstream of the inner tube the fuel/air mixture ignites within the flame front in the region of the combustion chamber.

Abstract: A purge air and fuel flow divider module (15) independently directs fuel flow to a plurality of manifolds (25, 27 and 29) and independently purges fuel from these manifolds with air when they are not flowing fuel. A secondary manifold (29) is filled with fuel before flowing metered burn flow to it and fuel flow is apportioned between a primary manifold (27) and the secondary manifold (29) during secondary manifold metered burn flow. Three manifolds are disclosed, a primary manifold (27), a secondary manifold (29) and a start or pilot manifold (25), each containing fuel nozzles (31, 33). A manifold main housing (61) contains a purge valve (41) and two three-way solenoids (37, 39) controlling respectively a secondary transfer valve (43) and a pilot nozzle transfer valve (45). Depending on the requirements, the solenoids position the valves to either provide manifold fill fuel, burn flow fuel or purge air to their respective manifolds.

Abstract: A pulse detonation system for a gas turbine engine having a longitudinal centerline axis extending therethrough includes a rotatable cylindrical member having a forward surface, an aft surface, and an outer circumferential surface, where at least one stage of circumferentially spaced detonation chambers is disposed therein. The pulse detonation system further includes a shaft rotatably connected to the cylindrical member and a stator configured in spaced arrangement around the forward surface, the aft surface, and the outer circumferential surface of the cylindrical member and a portion of the shaft. The stator has at least one group of ports formed therein which sequentially align with the detonation chambers as the cylindrical member rotates. In this way, detonation cycles are performed in the detonation chambers of each detonation stage so that reaction forces induced by the detonation cycles create a torque which causes the cylindrical member to rotate.

Abstract: A pilot fuel nozzle configuration for use in a combustor is disclosed having a natural frequency outside the range of the operating frequencies of a gas turbine engine. Multiple embodiments are disclosed for the improved pilot fuel nozzle including configurations for newly manufactured nozzles, repair to existing pilot nozzles, as well as multiple natural frequency levels for the improved pilot fuel nozzle. The pilot fuel nozzle comprises an elongated housing, first and second flanges, and a nozzle tip, with the first flange fixed to the elongated housing at a first end and the nozzle tip fixed to the second end, opposite of the first end. The second flange is fixed along the elongated housing and is used for attaching the pilot fuel nozzle to a combustor. The present invention incorporates an increased wall thickness along at least a mid-span portion of the pilot nozzle to increase the stiffness and change the natural frequency.

Abstract: A fuel nozzle for dispensing an atomized fluid spray into the combustion chamber of a gas turbine engine. The nozzle includes a body assembly with an inner fuel passage and an annular outer atomizing air passage. An array of turning vanes is disposed within the outer fuel passage. Each of the vanes is configured generally in the shape of an airfoil in having a pressure side and an opposing suction side, the suction side of each vane being spaced-apart from a juxtaposing pressure side of an adjacent vane to define a corresponding one of a plurality of aligned air flow channels therebetween. Fuel is directed through the fuel flow channels to be issued from the nozzle as a generally helical flow having a substantial uniform velocity profile.

Abstract: A method of operating a gas turbine combustor to achieve overall lower emissions of nitrous oxides by supplying a mixture of natural gas and hydrogen gas to the combustion chamber of the gas turbine in a manner that the localized concentration of hydrogen gas is greater than 0.1% by mass of the mass of the mixture, and less than 20.0% by mass of the mixture prior to combusting the mixture in the combustion chamber.

Abstract: The purpose is to improve the mixture ratio of a pre-mixer by a simple arrangement to form a more uniform premixed gases so as to materialize low NOx combustion. Two fuel nozzles disposed circumferentially of a pre-mixer are combined with a single air intake window to make a set, which set is used to produce swirls in a pair, thereby expediting mixing. Further, the inlet window is shaped such that its circumferential width is changed axially of the combustor, thereby changing the strength and size of the swirls to achieve the greatest effect. By reducing both the pre-mixer inlet windows and the partition walls in number, the manufacturing cost can be reduced, and by strengthening and optimizing the swirls, a combustor with superior low NOx performance can be provided, while it is possible to reduce the length of the pre-mixer necessary to obtain the same mixture ratio, leading to a cost reduction.

Abstract: A turbine containing system is disclosed. The system includes an intake section, a compressor section downstream from the intake section, a combustor section having a primary combustion system downstream from the intake section, a secondary combustion system downstream from the primary combustion system, a turbine section, an exhaust section and a load. The secondary combustion system includes an injector for transversely injecting a secondary fuel into a stream of combustion products of the primary combustion system. The injector including a coupling, a wall defining an airfoil shape circumscribing a fuel mixture passage, and at least one exit for communication between said fuel mixture passage and said stream of primary combustion products.

Abstract: There are provided an injector and an associated method for injecting and mixing gases, comprising a carbonaceous fuel and oxygen, in a combustion chamber of a combustion device. The injector has jets, which can be used to separately inject different combustion fuels. The injector is compatible with combustion devices that inject only gases, for example, a reheater that provides initial combustion in a power generation cycle or a reheater that recombusts a discharged gas from a gas generator and turbine. Further, the injector defines an annular space through which a recycle gas can be injected into the combustion chamber to lower the combustion temperature.

Abstract: A rotary ejector enhanced wave rotor pulsed detonation (WRPDE) engine and method are provided. The rotary-ejector-WRPDE includes a housing, one or more inlet ports in the housing, and a rotor mounted within the housing. The rotor includes a plurality of shrouded forward combustion passages in which detonative combustion occurs. Each forward combustion passage has an inlet end for communication with the inlet port. The rotor also includes a plurality of rear combustion passages which are in gaseous communication with the forward combustion passages. The rotor further includes a plurality of transitional combustion passages which join the forward combustion passages to the rear combustion passages and communicate with a source of bypass gas to provide a rotary ejector.

Abstract: A system and method for tuning a turbine comprises a turbine controller coupled to the turbine, a first computer system coupled to the turbine controller and located locally to the turbine, and a second computer system for exchanging data with the first computer system. The second computer system is located remotely from the turbine and exchanges data with the first computer system via a network connection such as the internet, an intranet or a virtual private network (VPN). Data relating to a characteristic such as turbine combustion dynamics and/or emissions is transmitted by the first computer system to the second computer system. The second computer system transmits control data over the network connection to the first computer to tune the turbine.

Abstract: A gas turbine comprises a compressor for compressing a gas supplied therein and discharging the compressed gas, a combustor in which the discharged gas from the compressor and a fuel are combusted, a turbine to be driven by a combustion gas from the combustor, and a water injection unit which injects water into the gas to be supplied to the compressor, thereby lowering the temperature of the gas to be introduced into the compressor to a temperature lower than the atmospheric temperature, and causing water droplets having been injected in the gas and within the compressor to be vaporized while flowing down therein, wherein the quantity of water spray injection is controlled while monitoring operational conditions of the gas turbine.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: A pulse combustion device has a number of combustors with upstream bodies and downstream nozzles. Coupling conduits provide communication between the combustors. For each given combustor this includes a first communication between a first location upstream of the nozzle thereof and a first location along the nozzle of another. There is second communication between a second location upstream of the nozzle and a second communication between a second location upstream of the nozzle of a second other combustor and a second nozzle location along the nozzle of the given combustor.

Abstract: A pulse detonation system for a gas turbine engine having a longitudinal centerline axis extending therethrough includes a rotatable cylindrical member having a forward surface, an aft surface, and an outer circumferential surface, where at least one stage of spaced detonation passages are disposed therethrough. The pulse detonation system further includes a shaft rotatably connected to the cylindrical member and a stator configured in spaced arrangement with the forward surface of the cylindrical member and a portion of the shaft. The stator has at least one group of ports formed therein alignable with the detonation passages as the cylindrical member rotates. In this way, detonation cycles are performed in the detonation passages so that combustion gases exit the cylindrical manner in a substantially tangential direction with respect to the outer circumferential surface to create a torque which causes the cylindrical member to rotate.

Abstract: A pulse detonation system for a gas turbine engine having a longitudinal centerline axis extending therethrough includes a rotatable cylindrical member having a forward surface, an aft surface, and an outer circumferential surface, where at least one stage of circumferentially spaced detonation chambers is disposed therein. The pulse detonation system further includes a shaft rotatably connected to the cylindrical member and a stator configured in spaced arrangement around the forward surface, the aft surface, and the outer circumferential surface of the cylindrical member and a portion of the shaft. The stator has at least one group of ports formed therein which sequentially align with the detonation chambers as the cylindrical member rotates. In this way, detonation cycles are performed in the detonation chambers of each detonation stage so that reaction forces induced by the detonation cycles create a torque which causes the cylindrical member to rotate.

Abstract: A counter-flow system for use in a turbine for selecting various mixtures of fluids for use with the system. The system may be used for a combustor for a gas powered turbine which employs a heat exchanger to combust a fuel without the emission of undesired chemical species. A gas powered turbine requires expanding gases to power the turbine blades. Fuel is combusted to produce the required gases. An oxidizer is introduced into the counter-flow system in a first direction before a first portion of fuel is introduced into the oxidizer. The fuel and oxidizer mixture is then flowed through a second pathway wherein the fuel and oxidizer mixture obtains a selected amount of thermal energy. Moreover, in a second pathway an equivalence ratio of the fuel and oxidizer mixture may be altered from the original fuel and oxidizer mixture.

Abstract: A premixer for a combustor to entrain a selected volume of fuel in an oxidizer to be combusted to heat the oxidizer. The combustor is for a gas powered turbine which employs the hypergolic or high energy air stream which will allow a fuel to be combusted in the absence of an additional ignition source. A heat exchanger and a catalyst combusts a first portion of fuel in air without the production of undesired chemical species. The catalyst is employed to lower the combustion temperature of the fuel to substantially eliminate selected chemical species. The fuel premixed in the air then encounters the catalyst and the fuel is combusted to increase the temperature of the air to an auto-ignition temperature so that no other ignition source is needed to combust additional fuel added later.

Abstract: There are provided an injector and an associated method for injecting and mixing gases, comprising a carbonaceous fuel and oxygen, in a combustion chamber of a combustion device. The injector has jets, which can be used to separately inject different combustion fuels. The injector is compatible with combustion devices that inject only gases, for example, a reheater that provides initial combustion in a power generation cycle or a reheater that recombusts a discharged gas from a gas generator and turbine. Further, the injector defines an annular space through which a recycle gas can be injected into the combustion chamber to lower the combustion temperature.

Abstract: The invention relates to a method for operating a burner (7) of a gas turbine (2), whereby a fossil fuel (B) is gasified and the gasified fossil fuel (B) is fed as synthesis gas (SG) into the burner (7) for combustion, said burner being connected to the gas turbine (2). The synthesis gas (SG) is divided into a first partial current (SG1) and a second partial current (SG2) and the partial currents (SG1, SG2) are each fed into the burner (7) for combustion. The invention also relates to a power plant (3), particularly a gas and steam turbine system (1), comprising a gasification device (132) for a fossil fuel (B).

Abstract: A pulse detonation system for a gas turbine engine having a longitudinal centerline axis extending therethrough, the pulse detonation system includes an air inlet duct in flow communication with a source of compressed air, the air inlet duct including at least one port formed therein for permitting compressed air to flow therethrough, a fuel injector mounted to the air inlet duct in circumferentially spaced relation to each port, and a device mounted to the air inlet duct in circumferentially spaced relation to each fuel injector for initiating a detonation wave. A rotatable ring member is also positioned in coaxial relation around a portion of the air inlet duct, with the ring member including at least one stage of detonation disposed therein. Accordingly, a detonation wave is produced in each detonation stage and combustion gases following each detonation wave create a torque which causes the ring member to rotate.

Abstract: A method of operating a gas turbine combustor to achieve overall lower emissions of nitrous oxides by supplying a mixture of natural gas and hydrogen gas to the combustion chamber of the gas turbine in a manner that the localized concentration of hydrogen gas is greater than 0.1% by mass of the mass of the mixture, and less than 20.0% by mass of the mixture prior to combusting the mixture in the combustion chamber.

Abstract: A fuel supply control system for a gas turbine includes a plurality of solenoid valves. The solenoid valves are energized in a timing sequence with a phase relationship designed to achieve a desired fuel flow. In one example, one solenoid valve is associated with a primary portion of a fuel manifold while at least two other solenoids are associated with a secondary portion of the manifold. A controller that energizes the solenoids to achieve the desired fuel flow can receive feedback information regarding turbine performance to make adjustments to the solenoid operation to bring the turbine performance closer to a desired level.

Abstract: An improved torch igniter for use in devices such as thrust augmenters, gas turbine engines, ramjets, combined-cycle engines and industrial burners. The torch igniter includes a housing with a combustion chamber. Fuel and oxidizer are delivered into the combustion chamber and ignited by an electronic ignition source, such as a plasma jet igniter or a spark igniter, so that an upstream recirculation zone and a downstream recirculation zone are created. The upstream recirculation zone stabilizes and pilots combustion within the combustion chamber, while the downstream recirculation zone augments the combustion event. Byproducts of the combustion event within the torch igniter provide a high mass flux with high thermal energy and strong ignition source radicals that are discharged through a neck portion of the housing and are thereafter employed to initiate a primary combustion event in a primary combustor.

Abstract: A gas turbine combustor has a combustion chamber into which fuel and air are supplied, wherein the fuel and the air are supplied into said combustion chamber as a plurality of coaxial jets.

Abstract: In a gas turbine having a combustion chamber (6) and a first row of guide vanes (3) arranged on the outflow side of the combustion chamber (6), a plurality of temperature sensors (8) for measuring temperature value, and a plurality of means (7;71,72,73,74,75,76) for introducing fuel into the combustion chamber (6), which means can be controlled on the basis of the temperature values, the temperature sensors (8) are spectrometers, are designed to measure a combustion-gas temperature and are arranged so as to measure a gas temperature which prevails immediately in front of the first row of guide vanes (3).

Abstract: The engine (10) has a compressor (12) for compressing air, a plurality of combustion chambers (22a-f) in which fuel mixed with the compressed air combusts, and a turbine (16) which is driven by the products of combustion. Each combustion chamber operates in a cycle comprising the phases of: A. charging the combustion chamber with a charge of compressed air from the compressor while preventing air from escaping from the combustion chamber, B. then compounding by charging with a compounding charge of gas to form a compounded charge; C. then injecting fuel into the combustion chamber so that there is spontaneous ignition and combustion of the fuel with the compounded charge; and D. then exhausting the products of combustion to the turbine.

Abstract: A combustion chamber 10 for a gas turbine operates in a highly diluted mode of combustion has a substantially toroidal plenum 30. At least one combustion fluid injector is located at the periphery of the plenum. The shape of the plenum and the disposition of the combustion fuel injectors are such that a vortex flow 2 is established in the combustion chamber in use. Gasses are entrained in the vortex which provides high enough levels of the gas re-circulation to allow for the onset of highly diluted combustion.

Abstract: A low emissions combustion method wherein, in an embodiment, a plurality of tangential fuel injectors introduce a fuel/air mixture at the combustor dome end of an annular combustion chamber in two spaced injector planes. Each of the spaced injector planes includes multiple tangential fuel injectors delivering premixed fuel and air into the annular combustor. A generally skirt-shaped flow control baffle extends from the tapered inner liner into the annular combustion chamber downstream of the fuel injector planes. A plurality of air dilution holes in the tapered inner liner underneath the flow control baffle introduce dilution air into the annular combustion chamber while another plurality of air dilution holes in the cylindrical outer liner introduces more dilution air downstream from the flow control baffle.

Abstract: A fuel injector for a gas turbine is disclosed which includes a nozzle body having a discharge portion defining a discharge orifice, the discharge portion including a fuel circuit for directing a hollow fuel film toward the discharge orifice from a fuel pump associated with the gas turbine, and an air assist circuit for directing pressurized air from a source external to the gas turbine toward the fuel film upstream from the discharge orifice to impinge on an inner surface of the fuel film issuing from the discharge orifice.

Abstract: A fuel ratio control method and a device therefor in a gas turbine combustor are provided in which supply ratio of pilot fuel to main fuel is appropriately maintained to thereby avoid unstable combustion, such as combustion vibration or misfire, and to realize a low NOx combustion. Fuel ratio control signal CSO is set corresponding to generator output LGEN put out by a generator output input unit 1. Pilot fuel ratio control is performed by correcting the fuel ratio control signal CSO corresponding to condition change in both or either one of combustion air and fuel. The condition change in the combustion air is at least one of changes in compressor on-line vane washing, ambient humidity, ambient pressure, etc. and the condition change in the fuel is a change in fuel component. The signal correction is also made by gas turbine deterioration factor or change rate in the generator output.

Abstract: The purpose is to improve the mixture ratio of a pre-mixer by a simple arrangement to form a more uniform premixed gases so as to materialize low NOx combustion. Two fuel nozzles disposed circumferentially of a pre-mixer are combined with a single air intake window to make a set, which set is used to produce swirls in a pair, thereby expediting mixing. Further, the inlet window is shaped such that its circumferential width is changed axially of the combustor, thereby changing the strength and size of the swirls to achieve the greatest effect. By reducing both the pre-mixer inlet windows and the partition walls in number, the manufacturing cost can be reduced, and by strengthening and optimizing the swirls, a combustor with superior low NOx performance can be provided, while it is possible to reduce the length of the pre-mixer necessary to obtain the same mixture ratio, leading to a cost reduction.

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

Abstract: A gas fuel nozzle for mounting in a combustor wall of a gas turbine engine, with an at least partially radially-directed array of gas fuel outlets extending beyond an air flow head having an array of compressed air jet apertures around the gas fuel outlets. The air flow head also has a deflector for creating an axial flow of air for deflecting in an axial direction the radially-injected gas fuel.

Abstract: A fuel nozzle with a ring of fuel spray orifices directing fuel jets at a fuel vortex generator having a fuel deflecting surface disposed downstream a distance from each fuel spray orifice. A mixing chamber is defined between the fuel spray orifices and the fuel deflecting surface having a surface contour oriented to deflect fuel jets into the mixing chamber in counter-rotating adjacent pairs of fuel laden vortices. An air inlet supplies air to the mixing chamber via an airflow vortex generator having an airflow deflecting surface with a surface contour oriented to deflect airflow into the mixing chamber in counter-rotating adjacent pairs of airflow vortices. A fuel-air mixture outlet downstream from the mixing chamber releases the fuel-air mixture into a combustor for ignition.

Abstract: A method for operating a pulse detonation system. The method includes providing a pulse detonation chamber including a plurality of detonation tubes extending therein, and detonating a mixture of fuel and air within each detonation tube such that at least a first tube is detonated at a different time than at least a second detonation tube.

Abstract: A method enables a gas turbine engine including a combustor to be operated. The combustor includes a mixer assembly including an air swirler, a premixer, a dome plate, and a heat shield. The method comprises discharging fluids from the air swirler into the premixer, and directing cooling fluids through a cooling opening defined in at least one of the swirler and the dome plate towards an upstream side of the heat shield for impingement cooling of the heat shield.

Abstract: This invention is a fuel injection control system for a turbine engine. The invention uses at least one fuel injector, having means for injecting fuel in pulses to the combustion chamber of a turbine engine, and an electronic control unit to receive and interpret input sensor signals from selected operating functions of the engine and to generate and direct fuel injection signals to modify the pulse duration and/or frequency of fuel injection in response to a deviation from a selected operating function, such as the desired operating speed, caused by variable operating loads encountered by the turbine engine. This configuration provides significantly greater fuel efficiency, better operational control and response time, and a lighter weight than is currently available in turbine engines. The invention may be used in many applications such as commercial, private, experimental and military aviation, power plant turbines, and other industrial, military and mining applications.

Abstract: In a method for operating a premix burner a water quantity is introduced into at least one of the burner and the reaction zone of the burner depending on at least one command variable formed from at least one measured value. The method can also be used if a burner is operated dry, i.e., without water injection for nitrogen oxide injection, and the injected water quantity is less than 20% of the fuel quantity. The method can be used advantageously especially if a characteristic value derived from combustion pulsations or material temperatures is used as a command variable. The method enables measured parameters to be kept below a permissible upper limit.

Abstract: A method and a device are described 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.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.

Abstract: In a thermal turbomachine with a housing, a combustion chamber, an ignition device, a fuel supply, and an air supply, the ignition device includes an ignition space and an igniter arranged in the ignition space. The fuel supply and the air supply open into the ignition space. The ignition space is connected to the combustion chamber, and the ignition device is arranged outside the housing of the thermal turbomachine. A process is shown for the ignition of the thermal turbomachine outside the housing.

Abstract: A method for operating a gas turbine engine facilitates reducing an amount of emissions from a combustor. The combustor includes a mixer assembly including a pilot mixer, a main mixer, and a centerbody that extends therebetween. The pilot mixer includes a pilot fuel nozzle and a plurality of axial swirlers. The main mixer includes a main swirler and a plurality of fuel injection ports. The method comprises injecting fuel into the combustor through the pilot mixer, such that the fuel is discharged downstream from the pilot mixer axial swirlers, and directing flow exiting the pilot mixer with a lip extending from the centerbody into a pilot flame zone downstream from said pilot mixer.

Abstract: A fuel injector system that reduces and/or eliminates combustion instability. The fuel injector system includes a pilot fuel injector, a pilot swirler that swirls air past the pilot fuel injector, a main airblast fuel injector having an aft end, inner and outer main swirlers that swirl air past the main airblast fuel injector, and an air splitter located between the pilot swirler and the inner main swirler. The air splitter includes at least one aft end cone angled radially outboard and axially positioned downstream of the main airblast fuel injector aft end. The air splitter divides a pilot air stream exiting the pilot swirler from an inner main air stream exiting the inner main swirler to create a bifurcated recirculation zone.

Abstract: The invention provides a method of operating a gas turbine arranged in a power generation system and comprising a source of compressed air, a combustor having a combustion chamber and multiple burners. A gas turbine controller controls the activation and deactivation of the single burners and/or burner groups according to a switching criterion that is proportional to the difference between a combustion chamber air inlet temperature and a temperature downstream of the combustion chamber. The switching criterion according to the invention fully accounts for large temperature fluctuations of the combustion chamber inlet air, which then result in only relatively small variations in burner equivalence ratio. The invention is particularly suited for application in compressed air energy power generation plants.

Abstract: Methods and apparatus, both devices and systems, for control of Zeldovich (thermal) NOx production in catalytic combustion systems during combustion of liquid or gaseous fuels in the post catalytic sections of gas turbines by reducing combustion residence time in the HC zone through control of the HC Wave, principally by adjusting the catalyst inlet temperature. As the fuel/air mixture inlet temperature (to the catalyst) is reduced, the HC Wave moves downstream (longer ignition delay time), shortens the residence time at high temperature, thereby reducing thermal NOx production. The countervailing increase in CO production by longer ignition delay times can be limited by selectively locating the HC Wave so that thermal NOx is reduced while power output and low CO production is maintained. NOx is reduced to on the order of <3 ppm, and preferably <2 ppm, while CO is maintained <100 ppm, typically <50 ppm, and preferably <5-10 ppm.

Abstract: In a method for igniting a thermal turbomachine, the turbomachine containing a combustion chamber (30), an ignition device (10), a fuel supply (60) and an air supply (70), and the ignition device (10) consisting of an ignition space (50) and of an ignitor (51) arranged in the ignition space (50), the fuel supply (60) and the air supply (70) discharging into the ignition space (50), and the ignition space (50) being connected to the combustion chamber (30) via a flame tube (40), prior to the ignition of the combustion chamber (30), an overall fuel/air mixture ratio &PHgr;overall=1/&lgr;overall greater than 1, with &lgr;overall being the overall air ratio, is set in the flame tube (40) of the ignition device (10).