| Entries |
| Document | Title | Date |
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| 20080229756 | SYSTEM AND METHOD FOR PASSIVE VALVING FOR PULSE DETONATION COMBUSTORS - A pulse detonation device contains a pulse detonation combustor which detonates a mixture of oxidizer and fuel. The fuel is supplied through fuel ducts and the fuel flow is controlled by fuel flow control devices. Oxidizer flow is provided through a main inlet portion and a flow control device directs the oxidizer flow to either the combustor or to a bypass duct, or both. The combustor further contains an ignition source. Each of the flow control devices, fuel flow control devices and ignition source are controlled by a control system to optimize performance at different thrust/power settings for the device. | 09-25-2008 |
| 20090277185 | PROPORTIONAL FUEL PRESSURE AMPLITUDE CONTROL IN GAS TURBINE ENGINES - Valve systems for controlling a flow of fuel in a gas turbine engine and related methods are provided. In one embodiment, the valve system includes a supply conduit, a proportional valve portion, and a pulsating valve portion. The supply conduit is adapted and configured for receiving and carrying a flow of fuel. The proportional valve portion is in fluid connection with the supply conduit adapted and configured to gradually adjust a pressure drop thereacross and thus a flow rate of fuel flowing therethrough. The pulsating valve portion is in fluid connection with the supply conduit, in parallel with the proportional valve portion, and is adapted and configured to rapidly adjust a pressure drop thereacross and thus a flow rate of fuel flowing therethrough. | 11-12-2009 |
| 20100058770 | Method and System for Controlling Fuel to a Dual Stage Nozzle - A method and system for controlling delivery of fuel to a dual stage nozzle in the combustor of a gas turbine. A liquid fuel is conveyed from a single stage fuel supply through a plurality of primary fuel supply lines to a first nozzle stage including a plurality of primary nozzles. A predetermined operating condition of the gas turbine is identified and a signal is produced in response to the identified operating condition. The signal effects actuation of valves located on secondary fuel supply lines extending from each of the primary fuel supply lines to supply fuel to respective secondary nozzles. | 03-11-2010 |
| 20100132365 | Start-Up Control For a Compact Lightweight Turbine - In a turbo-engine system, oil is circulated by means of a positive-displacement pump directly driven by the output shaft. The pump output pressure is monitored to trigger fuel injection when the turbine reaches sufficient speed during the start-up sequence. | 06-03-2010 |
| 20100199681 | FUEL SYSTEM OF GAS TURBINE ENGINES - A method for purging fuel from a fuel system of a gas turbine engine on shutdown of the engine comprises, in one aspect, terminating a fuel supply to the fuel system and using the residual compressed air to create a reversed pressure differential in the fuel system relative to a forward pressure differential of the fuel system used to maintain fuel supply for engine operation, and under the reversed pressure differential substantially purging the fuel remaining in the system therefrom to a fuel source. | 08-12-2010 |
| 20100287946 | Lean direct injection atomizer for gas turbine engines - A lean direct injection fuel nozzle for a gas turbine is disclosed which includes a radially outer main fuel delivery system including a main inner air swirler defined in part by a main inner air passage having a radially inner wall with a diverging downstream surface, an intermediate air swirler radially inward of the main inner air swirler for providing a cooling air flow along the downstream surface of the radially inner wall of the main inner air passage, and a radially inner pilot fuel delivery system radially inward of the intermediate air swirler. | 11-18-2010 |
| 20100287947 | Acoustically Tuned Combustion for a Gas Turbine Engine - A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves. | 11-18-2010 |
| 20100300109 | FUEL INJECTION METHOD - A method is provided for fuel injection in a sequential combustion system comprising a first combustion chamber and downstream thereof a second combustion chamber, in between which at least one vortex generator is located, as well as a premixing chamber having a longitudinal axis downstream of the vortex generator, and a fuel lance having a vertical portion and a horizontal portion, being located within said premixing chamber. The fuel injected is an MBtu-fuel. In said premixing chamber the fuel and a gas contained in an oxidizing stream coming from the first combustion chamber are premixed to a combustible mixture. The fuel is injected in such a way that the residence time of the fuel in the premixing chamber is reduced in comparison with a radial injection of the fuel from the horizontal portion of the fuel lance. | 12-02-2010 |
| 20110000219 | METHOD AND APPARATUS FOR ACTIVELY CONTROLLING FUEL FLOW TO A MIXER ASSEMBLY OF A GAS TURBINE ENGINE COMBUSTOR - An apparatus for actively controlling fuel flow from a fuel pump to a mixer assembly of a gas turbine engine combustor, where the mixer assembly includes a pilot mixer and a main mixer. The pilot mixer further includes an annular pilot housing having a hollow interior, a primary fuel injector mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing, a plurality of axial swirlers positioned upstream from the primary fuel injector. The fuel flow control apparatus further includes: at least one sensor for detecting dynamic pressure in the combustor; a fuel nozzle; and, a system for controlling fuel flow supplied by the fuel nozzle through the valves. The fuel nozzle includes: a feed strip with a plurality of circuits for providing fuel to the pilot mixer and the main mixer; and, a plurality of valves associated with the fuel nozzle and in flow communication with the feed strip thereof. The control system activates the valves in accordance with signals received from the pressure sensor. | 01-06-2011 |
| 20110041510 | FUEL CONTROL APPARATUS FOR GAS TURBINE ENGINE - In an apparatus for controlling a flow rate of fuel to be supplied to a combustion chamber of a gas turbine engine having a turbine rotated by combustion gas injected from the combustion chamber and a compressor for compressing air to be supplied to the combustion chamber, a first fuel flow rate at starting of the engine is calculated based on at least the detected output pressure of the compressor. In addition, a second fuel flow rate at starting of the engine is calculated based on at least the detected temperature of the inflowing air, the exhaust gas temperature and the rotational speed of the compressor, one of the first and second fuel flow rates is selected based on the detected exhaust gas temperature and operation of a fuel metering valve is controlled based on the selected fuel flow rate, thereby enabling to prevent white smoke at engine starting even when the engine is in the unstable range where the exhaust gas temperature is low. | 02-24-2011 |
| 20110048027 | Turbine Driven By Predetermined Deflagration Of Anaerobic Fuel And Method Thereof - The present invention discloses a turbine assembly ( | 03-03-2011 |
| 20110083444 | LOW BTU FUEL INJECTION SYSTEM - A system includes a gas turbine compressor including multiple radial protrusions disposed about a circumference of the gas turbine compressor. Each radial protrusion includes multiple gaseous fuel injection orifices configured to inject gaseous fuel into the gas turbine compressor. | 04-14-2011 |
| 20110094239 | Low NOx Combustor for Hydrogen-Containing Fuel and its Operation - An object of the present invention is to provide a gas turbine combustor that supports hydrogen-containing gas having a high burning velocity and is capable of performing low NOx combustion without reducing reliability of a burner. A first fuel nozzle is provided upstream of a combustion chamber and supplies fuel for activation and hydrogen-containing gas. The combustor has a primary combustion zone, a reduction zone and a secondary combustion zone. In the primary combustion zone, the fuel supplied from the first fuel nozzle is combusted under a fuel rich condition to form a burned gas containing a low concentration of oxygen. In the reduction zone, a hydrogen-containing gas is injected into the combustion chamber through a second fuel injection hole from a second fuel nozzle so that NOx generated in the primary combustion zone is reduced by an oxygen reaction of the hydrogen. In the secondary combustion zone, air for lean combustion is supplied into the combustion chamber so that unburned part of fuel is combusted under a fuel lean condition. | 04-28-2011 |
| 20110094240 | Swirl Generator - A swirl generator for injecting fuel into a gas turbine by injection of the fuel into a channel of the swirl generator of the gas turbine. The fuel injection is predetermined by the arrangement of a fuel injecting insert in a swirl generator segment. | 04-28-2011 |
| 20110100017 | Open Cycle Gas Turbine System Having An Enhanced Combustion Efficiency - An open cycle gas turbine system includes a storage tank, a closed helical supplier, a combustor, a gas turbine, a compressor, a propeller and a generator. After the combustion gas in the combustor enters the gas turbine, the combustion gas is expanded to apply a work and to produce a power to the gas turbine so as to operate the gas turbine so that the compressor is operated to compress the air, the generator is operated to generate an electric power, and the propeller is operated to propel vehicles. The closed helical supplier has a closely sealing effect so that the high pressure combustion gas in the combustor will not leak from the closed helical supplier and will not touch the solid fuel powder in the storage tank. | 05-05-2011 |
| 20110100018 | OPERATIONAL CONTROL SYSTEM OF GAS TURBINE - An operational control system of a gas turbine which is driven using mainly ammonia as fuel, wherein in a deteriorated combustibility operating region where the combustibility of ammonia deteriorates compared with the time of normal operation of the gas turbine, for example, at the time of cold startup of the gas turbine and right after startup or right before stopping operation, etc., a ratio of fossil fuel in the fuel which is fed to the gas turbine is increased over the time of normal operation. Due to this, even when using nonflammable ammonia as a main fuel, it is possible to stably start up, operate, and stop the gas turbine. | 05-05-2011 |
| 20110113787 | Pilot combustor in a burner - A pilot combustor particularly for use in a burner of a gas turbine engine is provided. A method for burning a fuel in a pilot combustor zone of a pilot combustor is also provided. The pilot combustor includes rotationally symmetric walls defining a combustion room with an exit having a rich concentration of fuel in air for burning the fuel for the creation of a flow of an non equilibrium unquenched concentration of radicals elevated to a temperature above 2000 K in the combustion room and directed along a centre line of the pilot combustor through a throat at the exit of the pilot combustor, wherein a quarl is located downstream of the throat of the pilot combustor. According to the method the pilot combustor is arranged upstream of a burner for providing a main lean partially premixed combustion process. | 05-19-2011 |
| 20110154829 | BURNER, COMBUSTOR AND REMODELING METHOD FOR BURNER - A combustor with a burner maintains combustion stability. The burner includes an air hole member | 06-30-2011 |
| 20110167832 | Method for operating a premix burner, and a premix burner for carving out the method - A method for operating a premix burner for gaseous fuels having a multi-stage pilot gas system whose diffusion fuel is injected into a flame chamber of the premix burner as at least two partial streams with different orientations, and a premix burner for carrying out the method. | 07-14-2011 |
| 20110185743 | Methods For Operating Gas Turbine Engines - Fuel is delivered in a gas turbine engine including a can annular combustor array that includes at least one combustor level subset within the array supplied by an independent fuel delivery system. Fuel is supplied only to one or more subsets during a first mode of operation. Fuel is supplied to the entire array of combustors during a second mode of operation. | 08-04-2011 |
| 20110192169 | GAS TURBINE ENGINE FUEL CONTROL SYSTEM - A fuel control system having a combustive energy value evaluator determining a combustive energy value of the fuel, and a controller calculating a desired flow rate based at least on the combustive energy value and controlling a fuel metering device such that the fuel flow rate corresponds to the desired fuel flow rate. | 08-11-2011 |
| 20110197594 | Method of Controlling a Combustor for a Gas Turbine - A method of controlling a combustor of a gas turbine is disclosed. The method includes operatively disposing a combustor can in a combustor of a gas turbine. The combustor can comprising a plurality of combustor fuel nozzles, each having a fuel injector and configured to selectively provide a liquid fuel, a liquid fluid or liquid fuel and liquid fluid to a fuel injector nozzle that is configured to provide, respectively, a plurality of liquid fuel jets, a plurality of liquid fluid jets or a combination thereof, that are in turn configured to provide an atomized liquid fuel stream, an atomized liquid fluid stream, or an atomized and emulsified liquid fuel-liquid fluid stream, respectively. The method also includes selectively providing an amount of fuel, fluid or a combination thereof to the fuel injector nozzle to produce an atomized fuel stream, atomized fluid stream, or an atomized and emulsified fuel-fluid stream, respectively. | 08-18-2011 |
| 20110203291 | METHODS AND SYSTEMS RELATING TO FUEL DELIVERY IN COMBUSTION TURBINE ENGINES - A fuel delivery system for a combustion turbine engine, comprising: a fuel line having a fuel compressor and parallel branches downstream of the fuel compressor: a cold branch that includes an after-cooler; and a hot branch that bypasses the after-cooler; a rapid heating value meter configured to measure the heating value of the fuel from the fuel source and transmit heating value data relating to the measurements; means for controlling the amount of fuel being directed through the cold branch and the amount of fuel being directed through the hot branch; and a fuel-mixing junction at which the cold branch and the hot branch converge; wherein the fuel-mixing junction resides in close proximity to a combustor gas control valve. | 08-25-2011 |
| 20110232296 | OPTICAL FUEL NOZZLE FLASHBACK DETECTOR - According to one aspect of the invention, a combustor is disclosed. The combustor can include a combustor housing, a plurality of nozzles disposed within the combustor housing, and a flame detector disposed on and in optical communication with each of the plurality of fuel nozzles, wherein each flame detector is configured to detect an optical property related to at least one of a flame holding condition and a flashback condition in a respective fuel nozzle. | 09-29-2011 |
| 20110252806 | METHODS FOR OPERATING A GAS TURBINE ENGINE APPARATUS AND ASSEMBLING SAME - A method for operating a gas turbine engine including a combustor assembly includes channeling a first fluid through a plurality of first nozzles into the combustor and igniting the first fluid downstream from the first nozzles. The method also includes increasing the operating speed of the engine and attaining a first predetermined percentage of a baseload by channeling the first fluid only through the first nozzles and then channeling a second fluid through a second nozzle into the combustor. The method also includes igniting the second fluid within the combustor downstream from the second nozzle. The method further includes channeling the second fluid to the first nozzles when the engine attains a second predetermined percentage of the baseload. The second predetermined percentage of the baseload is greater than the first predetermined percentage of the baseload. The method also includes terminating a flow of the first fluid through the first nozzles. | 10-20-2011 |
| 20110289932 | SYSTEM FOR FUEL AND DILUENT CONTROL - According to various embodiments, a system includes a fuel controller configured to control a fuel transition between a first flow of a first fuel and a second flow of a second fuel into a fuel nozzle of a combustion system. The fuel controller is configured to adjust a third flow of a diluent in combination with the second flow of the second fuel to maintain a pressure ratio across the fuel nozzle above a predetermined operating pressure ratio. | 12-01-2011 |
| 20110289933 | Hybrid Prefilming Airblast, Prevaporizing, Lean-Premixing Dual-Fuel Nozzle for a Gas Turbine Combustor - A dual fuel nozzle for a gas turbine combustor includes a hub defining a fuel inlet and a plurality of liquid fuel jets disposed at a downstream end of the hub. The fuel jets are oriented to eject liquid fuel radially outward from the hub. An annular air passage includes a swirler that imparts swirl to air flowing in the annular air passage, and a splitter ring is disposed in the annular air passage and surrounds the plurality of liquid fuel jets. The nozzle allows liquid fuels to be injected into a swirling annular airstream and then atomized, dispersed and vaporized inside a lean premixing dual fuel nozzle for a gas turbine combustor. | 12-01-2011 |
| 20110302927 | FUEL CONTROL METHOD FOR STARTING A GAS TURBINE ENGINE - A method for controlling a flow of fuel to a starting gas turbine engine includes controlling the flow of fuel to the engine by controlling fuel pressure to the engine, and modulating the flow of fuel to the engine if engine exhaust gas temperature approaches a given exhaust gas temperature to lower the engine exhaust gas temperature below the given exhaust gas temperature. | 12-15-2011 |
| 20120017601 | GAS TURBINE WITH IMPROVED PART LOAD EMISSIONS BEHAVIOR - In a method for the low-CO emissions part load operation of a gas turbine with sequential combustion, the air ratio (λ) of the operative burners ( | 01-26-2012 |
| 20120023964 | LIQUID-FUELED PREMIXED REVERSE-FLOW ANNULAR COMBUSTOR FOR A GAS TURBINE ENGINE - A reverse flow annular combustor for a gas turbine includes a pre-vaporizer/pre-mixing region within a dome section, a liquid fuel injection system feeding the pre-mixing region, a combustion region downstream of the pre-vaporizer/pre-mixing region, a means for guaranteeing stable and sustained combustion in the combustion region, and a dilution region downstream of the combustion region. | 02-02-2012 |
| 20120023965 | COMBUSTION CONTROLLER - The present invention relates to a combustion controller for controlling the combustion of a flow of combustible and/or combusting fluid. The combustion controller comprises an inlet and an outlet which defines a flow path between them. A magnetic-field generator is arranged to generate a magnetic field across the flow path such that in use the fluid flows in the flow path through the magnetic field. As the fluid flows through the magnetic field an electric current is induced in the fluid. This results in energy being supplied to the fluid. This energy can assist is reignition. If the magnetic field is an alternating magnetic field then this induces an alternating current in the fluid. The frequency of this current can be controlled such that certain chemical reactions are either promoted or inhibited. | 02-02-2012 |
| 20120031102 | TURBINE COMBUSTOR WITH FUEL NOZZLES HAVING INNER AND OUTER FUEL CIRCUITS - A combustor cap assembly for a turbine engine includes a combustor cap and a plurality of fuel nozzles mounted on the combustor cap. One or more of the fuel nozzles would include two separate fuel circuits which are individually controllable. The combustor cap assembly would be controlled so that individual fuel circuits of the fuel nozzles are operated or deliberately shut off to provide for physical separation between the flow of fuel delivered by adjacent fuel nozzles and/or so that adjacent fuel nozzles operate at different pressure differentials. Operating a combustor cap assembly in this fashion helps to reduce or eliminate the generation of undesirable and potentially harmful noise. | 02-09-2012 |
| 20120031103 | Combustor and the Method of Fuel Supply and Converting Fuel Nozzle for Advanced Humid Air Turbine - A fuel control device and method of a gas turbine combustor, for advanced humid air turbines, in which plural combustion units comprising plural fuel nozzles for supplying fuel and plural air nozzles for supplying air for combustion are provided. A part of the plural combustion units are more excellent in flame stabilizing performance than the other combustion units. A fuel ratio, at which fuel is fed to the part of the combustion units is set on the basis of internal temperature of the humidification tower and internal pressure of the humidification tower to control a flow ratio of the fuel fed to the plural combustion units. | 02-09-2012 |
| 20120036861 | METHOD FOR COMPENSATING FOR COMBUSTION EFFICIENCY IN FUEL CONTROL SYSTEM - Compensation is provided for a fuel demand signal of a gas turbine controller during transition between operating modes. The compensation adjusts fuel demand to account for combustion efficiency differences between the starting and ending operating mode that otherwise can lead to severe swings in combustion reference temperature and lean blowout. | 02-16-2012 |
| 20120036862 | SYSTEM AND METHOD FOR OPERATING A GAS TURBINE - A system for operating a gas turbine includes a compressor, a combustor, and a turbine. The combustor and turbine define a hot gas path. A sensor disposed outside the hot gas path measures internal thermal emissions from the combustor or turbine and generates a first signal reflective of the internal thermal emissions. The internal thermal emissions are infrared or ultraviolet emissions. A controller connected to the sensor receives the first signal and adjusts the compressor, combustor, or turbine in response to the first signal from the sensor. A method for operating a gas turbine includes measuring internal thermal emissions from inside a combustor or turbine using a sensor disposed outside the hot gas path. The method further includes generating a first signal reflective of the internal thermal emissions and adjusting the operation of the compressor, combustor, or turbine in response to the first signal from the sensor. | 02-16-2012 |
| 20120036863 | METHOD, APPARATUS AND SYSTEM FOR DELIVERY OF WIDE RANGE OF TURBINE FUELS FOR COMBUSTION - In operating a gas turbine, there can be a difference between the desired heating value of the fuel and the actual needs of the fuel for sustainable combustion during various stages of the turbine operation. In one aspect, combustible lean limit operation of the gas turbine free of lean blow out is enabled by adjusting fuel-air-ratio of the fuel and fuel-air mixture properties, based on the operation requirements of the turbine and flammability of the fuel components. | 02-16-2012 |
| 20120055167 | APPARATUS AND METHOD FOR MIXING FUEL IN A GAS TURBINE NOZZLE - A nozzle includes a fuel plenum and an air plenum downstream of the fuel plenum. A primary fuel channel includes an inlet in fluid communication with the fuel plenum and a primary air port in fluid communication with the air plenum. Secondary fuel channels radially outward of the primary fuel channel include a secondary fuel port in fluid communication with the fuel plenum. A shroud circumferentially surrounds the secondary fuel channels. A method for mixing fuel and air in a nozzle prior to combustion includes flowing fuel to a fuel plenum and flowing air to an air plenum downstream of the fuel plenum. The method further includes injecting fuel from the fuel plenum through a primary fuel passage, injecting fuel from the fuel plenum through secondary fuel passages, and injecting air from the air plenum through the primary fuel passage. | 03-08-2012 |
| 20120073305 | COMBUSTION CHAMBER AND METHOD FOR OPERATING A COMBUSTION CHAMBER - A combustion chamber of a gas turbine including first and second premixed fuel supply devices connected to a combustion device having first zones connected to the first premixed fuel supply devices and second zones connected to the second premixed fuel supply devices. The second fuel supply devices are shifted along a combustion device longitudinal axis with respect to the first fuel supply devices. The first zones are axially upstream of the second premixed fuel supply devices. | 03-29-2012 |
| 20120079831 | METHOD, APPARATUS AND SYSTEM FOR IGNITING WIDE RANGE OF TURBINE FUELS - In operating a gas turbine, there can be a difference between the desired heating value of the fuel and the actual needs of the fuel for the supplied fuel to be ignited. In one aspect, fuel parameters related to the molecular weight of the fuel such as specific gravity and pressure drop are determined. Ignitability of the fuel is calculated based on the fuel parameters and adjusted as necessary to bring the fuel's ignitability to designed values. The fuel's ignitability can be calculated without actually igniting the fuel and also without direct knowledge of the fuel's calorific value or its composition. | 04-05-2012 |
| 20120085100 | Combustor with a Lean Pre-Nozzle Fuel Injection System - The present application provides for a combustor for combusting a flow of fuel and a flow of air. The combustor may include a number of fuel nozzles, a lean pre-nozzle fuel injection system positioned upstream of the fuel nozzles, and a premixing annulus positioned between the fuel nozzles and the lean pre-nozzle fuel injection system to premix the flow of fuel and the flow of air. | 04-12-2012 |
| 20120090331 | SYSTEMS AND METHODS FOR SUPPLYING FUEL TO A GAS TURBINE - Systems and methods for supplying fuel to a gas turbine are described. A fuel may be received, and one or more parameters associated with the received fuel may be determined. Based at least in part upon the determined one or more parameters, a desired pressure for removing one or more liquids from the fuel utilizing a separator may be calculated. The operation of a pressure changing device may then be controlled in order to achieve the desired pressure. In certain embodiments, the operations of the method may be performed by a controller that includes one or more computers. | 04-19-2012 |
| 20120111019 | SYSTEM AND METHOD FOR COMBUSTION DYNAMICS CONTROL BY ACOUSTIC CONTROL/CANCELLATION OF FUEL FLOW FLUCTUATION AT FUEL INJECTION LOCATION - A combustion dynamics control system for an aviation based or land based gas turbine engine employs an acoustic driver that is configured to drive pressure perturbations across a premixed fuel injection orifice to substantially zero in response to a control signal such that fuel flow perturbations across the fuel injection orifice are substantially zero. | 05-10-2012 |
| 20120117976 | APPARATUS AND METHOD FOR IGNITING A COMBUSTOR - A nozzle includes a center body that defines an axial centerline. A shroud circumferentially surrounds at least a portion of the center body to define an annular passage between the center body and the shroud. A plenum is inside the center body and substantially parallel to the axial centerline, and an igniter is inside the center body and generally adjacent to the plenum. A method for igniting a combustor includes flowing a fuel through a center body axially aligned in a nozzle and flowing a working fluid through an annular passage, wherein the annular passage is substantially parallel to and radially outward of the center body. The method further includes projecting at least one of a beam, spark, or flame from an igniter located inside the center body. | 05-17-2012 |
| 20120125007 | METHOD AND SYSTEM FOR ENGINE IGNITION AND MONITORING - A method and system of engine ignition and monitoring is provided. A combustor includes a casing surrounding a combustion zone and an ignitor plug including a tip, a base, and a body extending therebetween. The body extends through the combustor casing such that the tip is proximate the combustion zone. The ignitor plug is configured to receive ignition energy through the base and generate an ignition source at the tip. The ignitor plug is further configured to sense combustion dynamics in the combustion zone and generate a signal relative to the sensed combustion dynamics. | 05-24-2012 |
| 20120125008 | LOW CALORIFIC VALUE FUEL COMBUSTION SYSTEMS FOR GAS TURBINE ENGINES - A 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 defined through the pressure vessel. A plurality of slave injectors is included. Each slave injector is disposed radially outward of and substantially parallel to the master injector for injecting fuel and air in an injection 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 injection plumes of the slave injectors. The slave injectors can be staged for reduced power operation. | 05-24-2012 |
| 20120131926 | ADVANCED LASER IGNITION SYSTEMS FOR GAS TURBINES INCLUDING AIRCRAFT ENGINES - A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system comprises at least one laser light source configured to generate a laser beam and an optical beam guidance component. The optical beam guidance component is configured to transmit the laser beam to irradiate on an oxygenated fuel mixture supplied into the combustion chamber at a region of highest ignitability to generate a combustor flame in a flame region. The system further includes an integrated control diagnostic component configured to detect at least a portion of a light emission and operable to control one or more combustion parameters based in part on the detected light emission. The system further includes additional enhanced ignition control configurations. A method for igniting a fuel mixture in an internal combustion engine is also presented. | 05-31-2012 |
| 20120131927 | Advanced Optics and Optical Access for Laser Ignition for Gas Turbines Including Aircraft Engines - A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system including a laser light source configured to generate a laser beam, an ignition port configured to provide optimized optical access of the laser beam to a combustion chamber and an optical beam guidance component disposed between the laser light source and the ignition port. The optical beam guidance component is configured to include optimized optic components to transmit the laser beam to irradiate on a fuel mixture supplied into the combustion chamber to generate a combustor flame in a flame region. A method for igniting a fuel mixture in an internal combustion engine is also presented. | 05-31-2012 |
| 20120167582 | SUPERCRITICAL OR MIXED PHASE FUEL INJECTOR - An apparatus is disclosed including a main fuel supply fluidly coupled to a main fuel valve and an arcuate fuel passage receiving main fuel through the main fuel valve. The arcuate fuel passage includes a passage diameter and a radius of curvature which provides sufficient rotational acceleration to the main fuel such that a liquid portion of the main fuel is deposited on an outer wall of the arcuate fuel passage. The apparatus includes a fuel injector nozzle that receives the main fuel from the arcuate fuel passage, and injects the main fuel into a combustion chamber for a turbine engine. The apparatus further includes a pilot fuel supply fluidly coupled to a pilot fuel passage and a fuel selector structured to selectively provide the main fuel or the pilot fuel to the fuel injector nozzle. | 07-05-2012 |
| 20120167583 | Method for Starting a Burner - A method for starting a burner for combusting synthesis gases is provided. The burner includes first and second fuel passages, the first fuel passage encompasses the second fuel passage in a substantially concentric manner and the gas transferred to the burner is mixed with combusting air and is combusted. In order to start the burner, the second fuel passage is first loaded with a synthesis gas to a predefined burner power at a first starting phase and the first fuel passage is loaded with the synthesis gas at a second starting phase. | 07-05-2012 |
| 20120186264 | GAS TURBINE COMBUSTOR - A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine. | 07-26-2012 |
| 20120186265 | Stabilizing the flame of a burner - A burner of a gas turbine including a reaction chamber and a plurality of jet nozzles opening into the reaction chamber is provided. Fluid is injected through an outlet into the reaction chamber by the jet nozzles using of a fluid stream wherein the fluid is burned into hot gas in the reaction chamber. An annular gap is disposed about the fluid stream for at least one jet nozzle so that a part of the hot gas is drawn out of the reaction chamber and flows opposite the fluid flow direction into the annular gap and is mixed with the fluid stream within the jet nozzle. The ring gap is formed by means of an insert tube, and wherein the insert rube includes a thickening at the upstream end. A method for stabilizing the flame of such a burner of a gas turbine is also provided. | 07-26-2012 |
| 20120192568 | Gas Turbine Combustor - A combustor of the prior art that defines the outlet position and direction of an air hole and suppresses adhesion of flame to an air hole outlet can reduce a discharge amount of NOx by increasing a distance over which fuel and air are mixed with each other. However, such a combustor is not sufficiently discussed for measures to suppress the occurrence of combustion oscillation resulting from the variation of a flame surface. | 08-02-2012 |
| 20120204571 | COMBUSTOR AND METHOD FOR INTRODUCING A SECONDARY FLUID INTO A FUEL NOZZLE - A combustor is disclosed that includes a baffle plate and a fuel nozzle extending through the baffle plate. The combustor may also include a shroud extending from the baffle plate and surrounding at least a portion of the fuel nozzle. A passage may be defined between the shroud and an outer surface of the fuel nozzle for receiving a first fluid. Additionally, the passage may be sealed from a second fluid flowing adjacent to the shroud. | 08-16-2012 |
| 20120210727 | METHOD FOR COMBUSTING HYDROGEN-RICH, GASEOUS FUELS IN A BURNER, AND BURNER FOR PERFORMING SAID METHOD - A method for the combustion of hydrogen-rich, gaseous fuels in combustion air in a burner of a gas turbine includes injecting the hydrogen-rich, gaseous fuel at least partially isokinetically with respect to the combustion air such that the partially hydrogen-rich, gaseous fuel is injected at least partially in the same direction and at least partially at the same velocity as the combustion air. | 08-23-2012 |
| 20120227413 | IGNITER - An igniter arranged to ignite combustion in a primary flow including a fuel and air mixture, the igniter including one or more geometric features arranged to: induce a shockwave flow structure at least partially disposed in the primary flow; and ignite the fuel and air mixture by virtue of the shockwave flow structure. The present disclosure also relates to a method of igniting combustion in a primary flow including a fuel and air mixture, the method including: providing one or more geometric features arranged to induce a shockwave flow structure; inducing the shockwave flow structure at least partially in the primary flow; and igniting the fuel and air mixture by virtue of the shockwave flow structure. | 09-13-2012 |
| 20120234015 | DUAL PUMP FUEL FLOW SYSTEM FOR A GAS TURBINE ENGINE AND METHOD OF CONTROLLING - A fuel flow system for a gas turbine includes a first pump, a main fuel flow path and a second pump. The first pump is connected to an actuator and a metering valve. The main fuel flow path is formed between the first pump and the metering valve. The second pump is connected to the main fuel flow path and supplements the fuel flow from the first pump under certain conditions. The second pump and first pump are in parallel. | 09-20-2012 |
| 20120234016 | 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. | 09-20-2012 |
| 20120240592 | Combustor with Fuel Nozzle Liner Having Chevron Ribs - The present application and the resultant patent provide a fuel nozzle for mixing a flow of air and a flow of fuel within a combustor. The fuel nozzle may have one or more air passages for the flow of air, one or more fuel pegs for the flow of fuel, and a liner in communication with the air passages and surrounding the fuel pegs. The liner may include a number of ribs thereon so as to promote mixing of the flow of air and the flow of fuel therein. | 09-27-2012 |
| 20120255310 | COMBUSTOR NOZZLE AND METHOD FOR SUPPLYING FUEL TO A COMBUSTOR - A combustor nozzle includes a center body and a shroud circumferentially surrounding at least a portion of the center body to define a passage between the center body and the shroud. A guide between the center body and the shroud can pivot with respect to the center body. A method for supplying fuel to a combustor includes flowing a working fluid through a nozzle at a mass flow rate and flowing a fuel through the nozzle. The method further includes sensing a flame holding event inside the nozzle and pivoting a guide inside the nozzle to increase the mass flow rate of the working fluid flowing through the nozzle. | 10-11-2012 |
| 20120260666 | MULTI-FUEL COMBUSTION SYSTEM - A multi-fuel combustion system is provided. The system includes a combustor basket adapted to combust at least two type of fuels. The combustor basket includes a circumferential wall with a plurality of openings. The combustion system further includes a first conduit adapted to provide a first type of fuel directly to the combustor basket and a second conduit adapted to provide a second type of fuel directly to the combustor basket. The combustion system also may include a third conduit adapted to inject at least one of the first type of fuel and the second type of fuel through the openings into the combustor basket. | 10-18-2012 |
| 20120291447 | COMBUSTOR NOZZLE AND METHOD FOR SUPPLYING FUEL TO A COMBUSTOR - A combustor nozzle includes a first and second liquid fuel passages that terminate at first and second fuel ports. A first diluent passage terminates at a first diluent outlet radially surrounding the second fuel ports. A second diluent passage terminates at a second diluent outlet between the first diluent outlet and the second fuel ports. A third diluent passage surrounds at least a portion of the first and second diluent passages. A method for supplying fuel to a combustor includes flowing a liquid fuel through a first fuel passage and flowing an emulsified liquid fuel through a second fuel passage. The method further includes flowing a first diluent through a shroud surrounding the second fuel passage to a first diluent passage surrounding at least a portion of the second fuel passage and flowing a second diluent through a second diluent passage radially disposed between the first diluent passage and the second fuel passage. | 11-22-2012 |
| 20120291448 | Flexible Combustor Fuel Nozzle - The present application provides a flexible combustor fuel nozzle. The flexible combustor fuel nozzle may include a main passage in communication with a source of natural gas and a source of low BTU fuel, a secondary passage surrounding the main passage and in communication with the source of low BTU fuel and a source of purge air, and a tertiary passage surrounding the secondary passage and in communication with the source of low BTU fuel, the source of purge air, and a source of diluent. | 11-22-2012 |
| 20130008172 | Systems and Methods for Modified Wobbe Index Control With Constant Fuel Temperature - The present application provides a gas turbine engine system for combusting a flow of fuel. The gas turbine engine system may include a combustor and a fuel system for providing the flow of fuel to the combustor. The fuel system may include a fuel heat exchanger so as to provide the flow of fuel to the combustor at a substantially constant temperature. | 01-10-2013 |
| 20130014514 | SYSTEMS AND METHODS FOR BULK TEMPERATURE VARIATION REDUCTION OF A GAS TURBINE THROUGH CAN-TO-CAN FUEL TEMPERATURE MODULATIONAANM Romig; Bryan WesleyAACI SimpsonvilleAAST SCAACO USAAGP Romig; Bryan Wesley Simpsonville SC USAANM Ziminsky; Willy SteveAACI GreenvilleAAST SCAACO USAAGP Ziminsky; Willy Steve Greenville SC USAANM Simons; Derrick WalterAACI GreerAAST SCAACO USAAGP Simons; Derrick Walter Greer SC US - A gas turbine includes a plurality of combustion chambers; at least one fuel nozzle for each of the combustion chambers; at least one fuel line for each fuel nozzle in each of the combustion chambers; at least one heat exchanger for each fuel line configured to adjust a temperature of a fuel flow to each fuel nozzle; and a controller configured to control each of the heat exchangers to reduce temperature variations amongst the combustion chambers. | 01-17-2013 |
| 20130061598 | SYSTEM AND METHOD FOR CONDITIONING A WORKING FLUID IN A COMBUSTOR - A system for conditioning a working fluid in a combustor includes a primary combustion chamber, a liner circumferentially surrounding the primary combustion chamber, and a primary nozzle in fluid communication with the primary combustion chamber. A secondary combustion chamber located outside of the primary combustion chamber includes a shroud that defines a fluid passage, a secondary nozzle, and means for igniting fuel in the secondary combustion chamber. A method for conditioning a working fluid in a combustor includes flowing the working fluid through a primary combustion chamber and flowing at least a portion of the working fluid through a secondary combustion chamber located outside of the primary combustion chamber. The method further includes flowing a fuel through the secondary combustion chamber, combusting the fuel, and flowing the combustion gases from the secondary combustion chamber into the primary combustion chamber. | 03-14-2013 |
| 20130074514 | SYSTEMS AND METHODS INVOLVING IMPROVED FUEL ATOMIZATION IN AIR BLAST FUEL NOZZLES OF GAS TURBINE ENGINES - Systems and methods involving improved fuel atomization in air-blast fuel nozzles of gas turbine engines are provided. In this regard, a representative method includes: providing fuel to a chamber defined by an inner surface; and continuously atomizing a portion of the fuel via interaction of the fuel with the inner surface. | 03-28-2013 |
| 20130086918 | METHOD FOR SWITCHING OVER A COMBUSTION DEVICE - An exemplary method for switching over a combustion device from operation with a first premix fuel to a second premix fuel includes reducing and stopping a first premix fuel supply and then starting a second premix fuel supply. In an intermediate phase, after the first premix fuel supply stop and before the second premix fuel supply start, the combustion device is operated with one or more pilot fuels generating diffusion flames. | 04-11-2013 |
| 20130091857 | Fuel System - A fuel system comprises a fuel actuation arrangement | 04-18-2013 |
| 20130111918 | COMBUSTOR ASSEMBLY FOR A GAS TURBOMACHINE - A turbomachine combustor assembly includes a combustor housing, a first combustion zone arranged in the combustor housing, a second combustion zone arranged downstream from the first combustion zone, and one or more injector assemblies positioned downstream from the first combustion zone and upstream from the second combustion zone. The one or more injector assemblies includes a first injector member having a first centerline axis and a second injector member having a second centerline axis. The second injector member extends though the first injector member with the second centerline axis being off-set from the first centerline axis. | 05-09-2013 |
| 20130118181 | PROCEDURE FOR IGNITING A TURBINE ENGINE COMBUSTION CHAMBER - A procedure for igniting a combustion chamber of a turbine engine, the chamber being fed with fuel by injectors and including an igniter mechanism igniting the fuel injected into the chamber. The procedure includes an initial stage during which fuel is injected into the chamber at a constant rate while simultaneously exciting the igniter mechanism, and in event of non-ignition of the chamber at an end of the initial stage, a second stage during which a rate at which fuel is injected is increased rapidly by 20% to 30%. | 05-16-2013 |
| 20130125558 | PREHEATING A SPARK PLUG - A method of igniting a turbine engine using a spark plug including a first electrode, a second electrode, and a semiconductor body between the first electrode and the second electrode, the semiconductor body having an exposed surface, the ignition method including: generating a spark adjacent to the exposed surface by applying a voltage difference greater than a first predetermined threshold between the first electrode and the second electrode; and prior to generating a spark, a preheating applying a voltage difference less than a second predetermined threshold between the first electrode and the second electrode, the second predetermined threshold being less than the first predetermined threshold. | 05-23-2013 |
| 20130174568 | COMBUSTOR AND METHOD FOR DISTRIBUTING FUEL IN THE COMBUSTOR - A combustor includes a tube bundle that extends radially across at least a portion of the combustor. The tube bundle includes an upstream surface axially separated from a downstream surface. A plurality of tubes extends from the upstream surface through the downstream surface, and each tube provides fluid communication through the tube bundle. A baffle extends axially inside the tube bundle between adjacent tubes. A method for distributing fuel in a combustor includes flowing a fuel into a fuel plenum defined at least in part by an upstream surface, a downstream surface, a shroud, and a plurality of tubes that extend from the upstream surface to the downstream surface. The method further includes impinging the fuel against a baffle that extends axially inside the fuel plenum between adjacent tubes. | 07-11-2013 |
| 20130174569 | SYSTEM AND METHOD FOR SUPPLYING A WORKING FLUID TO A COMBUSTOR - A system for supplying a working fluid to a combustor includes a fuel nozzle and a combustion chamber downstream from the fuel nozzle. A flow sleeve circumferentially surrounds the combustion chamber, and a plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the distribution manifold provides fluid communication through the distribution manifold to the plurality of fuel injectors. A method for supplying a working fluid to a combustor includes flowing a working fluid from a compressor through a combustion chamber and diverting a portion of the working fluid through a distribution manifold that circumferentially surrounds a plurality of fuel injectors circumferentially arranged around the combustion chamber. | 07-11-2013 |
| 20130174570 | ENGINE SYSTEMS WITH ENHANCED START CONTROL SCHEDULES - An engine system for starting a gas turbine engine includes a starter coupled to the gas turbine engine and configured to provide torque to the gas turbine engine; and a controller coupled to the starter and configured to evaluate an engine system parameter and to select from a plurality of start modes for starting the gas turbine engine based on the engine system parameter. | 07-11-2013 |
| 20130174571 | Fuel Flow Control Method and Fuel Flow Control System of Gas Turbine Combustor for Humid Air Gas Turbine - Provided is a fuel flow control method of a gas turbine combustor provided in a humid air gas turbine, by which method NOx generation in the gas turbine combustor is restricted before and after the starting of humidification and combustion stability is made excellent. | 07-11-2013 |
| 20130180260 | COMBUSTOR RECOVERY METHOD AND SYSTEM - A method is disclosed for controlling gas turbine operation in response to lean blowout of a combustion can. The gas turbine comprises a pair of combustion cans. The method includes sensing that a first combustion can is extinguished during a full load operation of the gas turbine, adjusting a fuel ratio between the fuel nozzles in each can, delivering a richer fuel mixture to the fuel nozzles nearest to the cross-fire tubes, generating a cross-fire from the second combustion can to the first combustion can, detecting a recovery of the turbine load, and adjusting the fuel ratio to the normal balanced fuel distribution between the fuel nozzles in each can. | 07-18-2013 |
| 20130186098 | FLUID FLOW CONTROL DEVICE AND METHOD - A fluid flow control system includes a fluid inlet, a central chamber, a first nozzle extending from a first side of the central chamber and comprising a first throat, a second nozzle extending from a second side of the central chamber opposite the first side and comprising a second throat, and a flow control shuttle. The flow control shuttle includes a first needle having a first tapered portion positioned within the first throat for controlling flow through the first nozzle and a second needle having a second tapered portion positioned within the second throat for controlling flow through the second nozzle. | 07-25-2013 |
| 20130186099 | Gas Turbine Combustor - A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine. | 07-25-2013 |
| 20130192243 | FUEL NOZZLE FOR A GAS TURBINE ENGINE AND METHOD OF OPERATING THE SAME - A fuel nozzle assembly for use with a turbine engine includes at least one fuel conduit coupled to at least one fuel source. The fuel nozzle assembly also includes at least one swirler that includes at least one wall having a porous portion. The at least one wall is coupled to the at least one fuel conduit. The porous portion is formed from a material having a porosity that facilitates fuel flow therethrough. At least one fuel flow path is thereby defined through the porous portion of the at least one wall. | 08-01-2013 |
| 20130192244 | HYBRID APU START FUEL SYSTEM - A fuel system for an auxiliary power unit includes a mechanical fuel pump, an electric fuel pump, and a controller. The mechanical fuel pump provides fuel flow to the auxiliary power unit and has an output dependent upon an operational speed of the auxiliary power unit. The electric fuel pump provides fuel flow to the auxiliary power unit, and is located in flow series with the mechanical fuel pump. The controller causes the electric fuel pump to provide fuel flow during starting of the auxiliary power unit. | 08-01-2013 |
| 20130192245 | Gas Turbine Combustor and Operating Method Thereof - A gas turbine combustor has a chamber supplied with fuel and air and a multi-burner having a plurality of burners provided with an air hole plate having a plurality of air holes, and fuel nozzles for supplying fuel to the air holes in the air hole plate; the multi-burner is made up of a center burner disposed in the center and a plurality of outer burners around the center burner, the outer burners are divided into inner fuel nozzles and outer fuel nozzles to separately supply fuel through fuel systems, and the fuel is supplied to the fuel nozzles in the center burner or to the fuel nozzles in the center burner and the inner fuel nozzles in the outer burners disposed around the center burner in a partial load condition in which the load is lower than that of when all the fuel systems are used to supply fuel. | 08-01-2013 |
| 20130192246 | DUAL FUEL AIRCRAFT ENGINE CONTROL SYSTEM AND METHOD FOR OPERATING SAME - A dual fuel engine control system comprising a first fuel control system configured to control the flow of a first fuel to an aircraft gas turbine engine, and a second fuel control system configured to control the flow of a second fuel to the aircraft gas turbine engine. | 08-01-2013 |
| 20130199199 | DRIVE FOR A TURBINE AND DRIVE METHOD - The invention relates to the drive for a turbine, in particular for an aviation turbine, as well as to a method for operating such a turbine. An aviation turbine is a gas turbine that accelerates an aircraft. The invention further relates to an aircraft having the drive for a turbine. According to the invention, a drive for a turbine is provided with a compressor for compressing air, with a nozzle for injecting a first fuel into the compressed air, and with a combustion chamber for igniting the air-fuel mixture. Furthermore, the drive comprises another nozzle for injecting a second fuel. The nozzle for injecting a first fuel serves for starting the drive or a turbine engine comprising the drive as well as a turbine, which provides mechanical energy by the igniting the air-fuel mixture. Therefore, the first fuel is a conventional fuel, in particular kerosene. It is thus ensured that the engine can be started at any time, because it is, or at least can be, of a conventional design in this regard. The second nozzle serves for injecting a new fuel, which at least at first is a liquid gas. In particular, a mixture of and Bio LNG with a high calorific value, which is drawn from a tank and fed to the combustion chamber in an insulated pressure pipe, is used as the liquid gas. | 08-08-2013 |
| 20130199200 | FUEL DISTRIBUTION WITHIN A GAS TURBINE ENGINE COMBUSTOR - A fuel system for a gas turbine engine includes a plurality of duplex nozzles arranged on each side of top dead center and a plurality of simplex nozzles. A primary manifold is operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles and a secondary manifold is operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. An equalizer valve that is in communication with both the primary manifold and the secondary manifold distributes fuel at various pressures to both the primary and secondary manifolds. | 08-08-2013 |
| 20130205797 | FUEL HEATING SYSTEM FOR POWER PLANT - A fuel heating system for a power plant includes a fuel source for a gas turbine system, wherein the fuel source contains a fuel at a first temperature. Also included is a feedwater source for distributing a feedwater. Further included is a fuel heater configured to receive the fuel at the first temperature and distribute the fuel to the gas turbine system at a second temperature and capable of receiving the feedwater from the feedwater source. Yet further included is at least one boiler drum for containing a boiler substance. Also included is at least one heat exchanger for receiving the feedwater and the boiler substance and distributing the feedwater to the fuel heater and the boiler substance to the blowdown system. | 08-15-2013 |
| 20130213052 | COMBUSTION DEVICE WITH PULSED FUEL SPLIT - It is described a combustion device control unit and a combustion device, e.g. a gas turbine, which determine on the basis of at least one operating parameter whether the combustion device is in a predefined operating stage. In response hereto, there is generated a control signal configured for setting a ratio of at least two different input fuel flows to a predetermined value (psc | 08-22-2013 |
| 20130213053 | CONTROL OF A FUEL METERING DEVICE FOR TURBOMACHINE - A control of a fuel metering device for a turbine engine as a function of a weight flow rate setpoint includes responding to at least one validity criterion to select a weight flow rate from among: a weight flow rate calculated as a function of a position signal; a weight flow rate calculated as a function of the position signal and of at least one temperature measurement signal; a weight flow rate calculated as a function of the position signal and of at least one permittivity measurement signal; a weight flow rate calculated as a function of the position signal, of at least one temperature measurement signal, and of at least one permittivity measurement signal; and a weight flow rate calculated as a function of a temperature measurement signal, of a permittivity measurement signal, and of a volume flow rate measurement signal. | 08-22-2013 |
| 20130219910 | EXHAUST TEMPERATURE BASED THRESHOLD FOR CONTROL METHOD AND TURBINE - A gas turbine, computer software and a method for controlling an operating point of the gas turbine that includes a compressor, a combustor and at least a turbine is provided. The method comprises: determining an exhaust pressure at an exhaust of the turbine; measuring a compressor pressure discharge at the compressor; determining a turbine pressure ratio based on the exhaust pressure and the compressor pressure discharge; calculating a primary to lean-lean mode transfer threshold reference curve as a function of the turbine pressure ratio, where the primary to lean-lean mode transfer threshold curve includes points at which an operation of the gas turbine is changed between a primary mode to a lean-lean mode; and controlling the gas turbine to change between the primary mode and the lean-lean mode. | 08-29-2013 |
| 20130232986 | COMBUSTOR AND METHOD FOR REDUCING THERMAL STRESSES IN A COMBUSTOR - A combustor includes a combustion chamber and a casing that circumferentially surrounds the combustion chamber to at least partially define an annular passage between the casing and the combustion chamber. A fuel plenum extends radially through the casing to provide fluid communication through the casing to the annular passage, and a liner extends inside at least a portion of the fuel plenum to prevent fuel from directly impinging upon at least a portion of the fuel plenum. A method of reducing thermal stresses in a combustor includes flowing a fuel inside a fuel plenum that extends radially through a casing, shielding at least a portion of the fuel plenum from direct impingement by the fuel radially inward of the casing, and flowing the fuel from the fuel plenum into an annular passage between the casing and a combustion chamber. | 09-12-2013 |
| 20130232987 | GAS TURBINE FUEL INJECTOR WITH INSULATING AIR SHROUD - A fuel injector for a gas turbine engine is disclosed. The fuel injector includes an injector housing extending from a first end to a second end along a longitudinal axis. The second end of the housing is fluidly coupled to a combustor of the turbine engine and the housing includes a liquid fuel gallery annularly disposed about the longitudinal axis. The fuel injector also includes a stem extending longitudinally from the first end of the housing to a third end. The stem includes a liquid tube configured to deliver liquid fuel to the fuel injector. The fuel injector also includes an annular shell extending along the longitudinal axis from the first end to the third end and circumferentially disposed about the stem. The fuel injector further includes an insulating air shroud formed inside the shell. The air shroud includes a layer of air between the shell and the stem. | 09-12-2013 |
| 20130232988 | BURNER FOR A GAS COMBUSTOR AND A METHOD OF OPERATING THE BURNER THEREOF - A burner for a gas combustor and a method of operating the burner are disclosed. The burner includes a front surface area divided into a plurality of subareas and inlets arranged on the front surface area such that each subarea is encircled by at least four inlets and such that during operation of the burner, a gas recirculation in the combustor is facilitated corresponding to each subarea. | 09-12-2013 |
| 20130247578 | METHOD FOR OPERATING A FIXED GAS TURBINE, DEVICE FOR REGULATING THE OPERATION OF A GAS TURBINE AND POWER PLANT - A method of operating a gas turbine, a device for regulating the starting and/or the operation of a gas turbine and a power plant are provided. The method includes continuously extracting fuel from a fuel network, and combusting, in at least one combustion chamber of a gas turbine, the fuel by adding combustion air. For an increase of a fuel stream supplied to the at least one combustion chamber, a fuel volume is extracted from a fuel store and supplied to the fuel still to be supplied to the at least one combustion chamber. | 09-26-2013 |
| 20130255270 | PASSIVE EQUILIZATION FLOW DIVIDER VALVE - A method and system for providing fuel to primary and secondary fuel nozzles in a gas turbine engine fuel system comprises generating a fuel flow and routing primary fuel from the fuel flow to a primary fuel nozzle. Backpressure on the fuel flow is maintained using a valve. The valve is opened at increased fuel flow to route secondary fuel from the fuel flow to a secondary fuel nozzle. The valve is progressively opened under increasing fuel flows to reduce a pressure drop across the valve produced by the secondary fuel. | 10-03-2013 |
| 20130255271 | Fuel Supply System - A fuel supply system is provided having a first fuel gas compressor configured to be driven by a motor and a second fuel gas compressor configured to be driven by a shaft of a gas turbine system. The first fuel gas compressor and the second fuel gas compressor are configured to supply a pressurized fuel flow to a combustor of the gas turbine system, and the first fuel gas compressor and the second fuel gas compressor are coupled to one another in series. | 10-03-2013 |
| 20130276452 | Gas turbine purge process - This purge process of a gas turbine supply pipe network provided with fuel (diesel or natural gas) at least partly containing synthesis gas comprises of injection of inert gas in intervalve portions or collectors of the pipe network likely to contain fuel when the fuel supply is stopped. This injection of gas is implemented in the said portions of the network according to a sequence of respective injection. | 10-24-2013 |
| 20130283810 | COMBUSTION NOZZLE AND A RELATED METHOD THEREOF - A combustion nozzle includes at least one passage having a mixing section and an exit section. The mixing section includes an air inlet, and a fuel inlet. The mixing section has a first length and a first diameter. The exit section has a second length different from the first length, and a second diameter different from the first diameter. | 10-31-2013 |
| 20130283811 | FUEL CIRCUIT FOR AN AVIATION TURBINE ENGINE, THE CIRCUIT HAVING A FUEL PRESSURE REGULATOR VALVE - A fuel circuit for an aviation turbine engine, the fuel circuit including: a main fuel line for feeding fuel to a combustion chamber of the engine and including a positive displacement pump; an auxiliary fuel line connected to the main fuel line at a junction situated downstream from the pump and serving to feed fuel to hydraulic actuators to control variable-geometry equipment of the engine, the auxiliary fuel line including electrohydraulic servo-valves upstream from each actuator; and a fuel pressure regulator valve arranged on the main fuel line downstream from the pump. | 10-31-2013 |
| 20130291550 | AERO COMPRESSION COMBUSTION DRIVE ASSEMBLY CONTROL SYSTEM - A control system for an aero compression combustion drive assembly, the aero compression combustion drive assembly having an engine member, a transmission member and a propeller member, the control system including a sensor for sensing a pressure parameter in each of a plurality of compression chambers of the engine member, the sensor for providing the sensed pressure parameter to a control system device, the control system device having a plurality of control programs for effecting selected engine control and the control system device acting on the sensed pressure parameter to effect a control strategy in the engine member. A control method is further included. | 11-07-2013 |
| 20130298569 | GAS TURBINE AND METHOD FOR OPERATING SAID GAS TURBINE - The present disclosure relates to gas turbine including a combustion system with several burners, a conduit system with a fuel manifold for providing the burners with liquid fuel and a system for aerating the liquid fuel with gas. The system for aerating the liquid fuel with gas is located upstream to the fuel manifold. | 11-14-2013 |
| 20130305734 | Fuel Plenum Premixing Tube with Surface Treatment - The present application provides a micro-mixer fuel plenum for mixing a flow of fuel and a flow of air in a combustor. The micro-mixing fuel plenum may include an outer barrel and a number of mixing tubes positioned within the outer barrel. The mixing tubes may include one or more heat transfer features thereon. | 11-21-2013 |
| 20130305735 | GAS TURBINE SYSTEM - Provided is a gas turbine including: a first compressor which compresses air; a mixer which adds the compressed air from the first compressor to fuel and generates a fuel mixture; a combustor which combusts the generated fuel mixture from the mixer; a plurality of flow meters which adjusts an amount of the air or the fuel injected into the mixer; and a control unit which maintains the Wobbe Index of the fuel mixture within a predetermined Wobbe Index rang. | 11-21-2013 |
| 20130305736 | METHOD AND APPARATUS FOR INCREASING COMBUSTION EFFICIENCY AND REDUCING PARTICULATE MATTER EMISSIONS IN JET ENGINES - A portable on-demand hydrogen supplemental system producing hydrogen gas and mixing the hydrogen gas with the air used for combustion of the jet fuel to increase the combustion efficiency of said jet fuel. Hydrogen increases the laminar flame speed of the jet fuel during combustion thus causing more fuel to be burned and lowering particulate matter emissions. Hydrogen is supplied to the jet engine at levels well below it lower flammability limit in air of 4%. Hydrogen and oxygen is produced by an electrolyzer from nonelectrolyte water in a nonelectrolyte water tank. The system utilizes an onboard diagnostic (OBD) interface in communication with the jet's control systems, to regulate power to the system so that hydrogen production for the jet engine only occurs when the jet engine is running. The hydrogen gas produced is immediately consumed by the jet engine. No hydrogen is stored on, in or around the jet. | 11-21-2013 |
| 20130318992 | Combustor with a Brief Severe Quench Zone - The present application provides a combustor for combusting a number of flows of air and a number of flows of fuel. The combustor may include a central swirler for producing a high swirl quench air flow, a number of trapped vortex cavities surrounding the central swirler for producing a flow of combustion gases, a brief severe quench zone downstream of the trapped vortex cavities to quench the flow of combustion gases between an outer quench air flow and the high swirl quench air flow, and an expansion zone downstream of the brief severe quench zone. | 12-05-2013 |
| 20130327054 | HYBRID SLINGER COMBUSTION SYSTEM - A hybrid combustor combines two distinct fuel injection sources to spray fuel in the combustor. The combustor combines a rotary fuel slinger for spraying fuel in a first combustion zone during high power level and cruise conditions and a set of fuel nozzles for spraying fuel in a second combustion zone during lower power level and starting conditions. | 12-12-2013 |
| 20140000274 | METHODS AND APPARATUS FOR CO-FIRING FUEL | 01-02-2014 |
| 20140000275 | LNG FUEL HANDLING FOR A GAS TURBINE ENGINE | 01-02-2014 |
| 20140000276 | TURBOMACHINE COMPRISING A PRIVILEGED INJECTION DEVICE AND CORRESPONDING INJECTION METHOD | 01-02-2014 |
| 20140007584 | SYSTEM AND METHOD FOR REDUCING PRESSURE OSCILLATIONS WITHIN A GAS TURBINE ENGINE - In one embodiment, a system for reducing pressure oscillations within a gas turbine engine includes at least one fuel injector configured to inject fuel into a combustor. The system also includes a valve fluidly coupled to the at least one fuel injector. The system further includes a controller communicatively coupled to the valve. The controller is configured to cycle the valve between an open position and a closed position at a first frequency and a first duty cycle while a magnitude of pressure oscillations within the combustor is less than a threshold value, to cycle the valve between the open position and the closed position at a second frequency and a second duty cycle while the magnitude of the pressure oscillations within the combustor is greater than or equal to the threshold value, and to adjust the second frequency based on a measured frequency of the pressure oscillations. | 01-09-2014 |
| 20140007585 | LIQUID FUEL ASSIST IGNITION SYSTEM OF A GAS TURBINE AND METHOD TO PROVIDE A FUEL/AIR MIXTURE TO A GAS TURBINE - A liquid fuel assist ignition system for providing a fuel/air mixture to a gas turbine in its start-up phase includes a high pressure tank, a vacuum pump connected to the high pressure tank, a liquid fuel inlet connected to the high pressure tank, an air inlet connected to the high pressure tank, and an outlet of the high pressure tank connected to a burner of the gas turbine. | 01-09-2014 |
| 20140033731 | Method for fuel temperature control of a gas turbine - The present invention relates to a method for controlling the fuel temperature of a gas turbine, where parameters are determined as input values, where the parameters are compared with emission-optimized nominal values and an optimum fuel temperature is determined, and where the fuel to be supplied to a combustion chamber is heated or cooled. | 02-06-2014 |
| 20140053569 | METHOD FOR MIXING A DILUTION AIR IN A SEQUENTIAL COMBUSTION SYSTEM OF A GAS TURBINE - The invention concerns a method for mixing a dilution air with a hot main flow in sequential combustion system of a gas turbine, wherein the gas turbine essentially comprises at least one compressor, a first combustor which is connected downstream to the compressor, and the hot gases of the first combustor are admitted to at least one intermediate turbine or directly or indirectly to at least one second combustor. The hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery, wherein at least one combustor runs under a caloric combustion path having a can-architecture. At least one dilution air injection is introduced into the first combustor, and wherein the direction of the dilution air injection is directed against or in the direction of the original swirl flow inside of the first combustor. | 02-27-2014 |
| 20140060069 | COMBUSTOR INCLUDING COMBUSTION NOZZLE AND AN ASSOCIATED METHOD THEREOF - A combustor including a combustion nozzle. The combustion nozzle includes a mixing section and an exit section. The mixing section includes an air inlet, and a fuel inlet. The exit section includes a plurality of jets on an exit surface. The combustor further includes a combustion zone, including a combustion liner, disposed downstream and in fluidic communication with the combustion nozzle. The combustor is configured wherein a, NOx emission of the combustor is related to 1/R, where R is a Reynolds number ratio of a jet of the plurality of jets to the combustion liner. A method for achieving NOx reduction in a combustion nozzle. | 03-06-2014 |
| 20140060070 | METHOD AND APPARATUS FOR INCREASING COMBUSTION EFFICIENCY AND REDUCING PARTICULATE MATTER EMISSIONS IN JET ENGINES - A portable on-demand hydrogen supplemental system producing hydrogen gas and mixing the hydrogen gas with the air used for combustion of the jet fuel to increase the combustion efficiency of said jet fuel. Hydrogen increases the laminar flame speed of the jet fuel during combustion thus causing more fuel to be burned and lowering particulate matter emissions. Hydrogen is supplied to the jet engine at levels well below it lower flammability limit in air of 4%. | 03-06-2014 |
| 20140060071 | COOLED PILOT FUEL LANCE - A device for injecting fuel into a combustion chamber of a gas turbine is provided, having a distribution section to which a first fuel channel, a second fuel channel and an injection channel are coupled. The first fuel channel and the second fuel channel are arranged such that a) fuel is transportable by one of the first fuel channel and the second fuel channel to the distribution section, and b) a first quantity of fuel is transportable by the other one of the first fuel channel and the second fuel channel out of the distribution section. The injection channel is arranged such that a second quantity of fuel is injectable from the distribution section into the combustion chamber. The device further comprises an end cap with a protrusion having the injection channel inside, and extending inside the inner tube. | 03-06-2014 |
| 20140069105 | COMPRESSOR SURGE PREVENTION DIGITAL SYSTEM - Methods and devices for anticipating a surge in a gas turbine engine. Controlled pressure signal(s) may be compared with reference pressure signal(s), each of the controlled pressure signal(s) and reference pressure signal(s) having an associated time value. If the controlled pressure signal(s) are less than the reference pressure signal(s), a controlled pressure curve may be fitted through a predetermined number of points based on the controlled pressure value(s) and associated time value(s). A reference pressure curve may be fitted through the predetermined number of points based on the reference pressure value(s) and associated time value(s). A time to compressor surge may be estimated based on an intersection of the controlled pressure curve and the reference pressure curve. | 03-13-2014 |
| 20140083110 | SEAL FOR FUEL DISTRIBUTION PLATE - A fuel flow passes through a micromixer section of a gas turbine that includes a plurality of mixing tubes for transporting a fuel/air mixture and a distribution plate including a plurality of distribution holes and a plurality of tube holes for accommodating the mixing tubes. Each of the mixing tubes includes a plurality of fuel holes through which fuel enters the mixing tubes. The tube holes and the mixing tubes form a plurality of annulus areas between the distribution plate and the mixing tubes. The distribution holes and the annulus areas are configured to pass the fuel flow through the distribution plate toward the fuel holes. A flow management device modifies an effective size of the annulus areas to control a distribution of the fuel flow through the distribution holes and the annulus areas of the distribution plate to provide a uniform fuel/air composition in each of the mixing tubes. | 03-27-2014 |
| 20140090393 | Minimum Pressure and Shutoff Valve - An apparatus and method for controlling a pressure drop of a fluid is disclosed. A sleeve includes a plurality of flow passages in a wall of the sleeve. A member slides relative to the sleeve to reveal at least a portion of the plurality of flow passages to control flow of the fluid through the flow passages. The plurality of flow passages are at azimuthally- and axially-staggered locations around the sleeve. | 04-03-2014 |
| 20140090394 | FLOW MODIFIER FOR COMBUSTOR FUEL NOZZLE TIP - A fuel injector nozzle assembly includes a body extending along an axis and a core swirl plug positioned at least partially within the body. The core swirl plug has a flow modifying structure configured to swirl fuel at a location upstream from a distal end of the nozzle assembly. | 04-03-2014 |
| 20140090395 | GAS TURBINE ENGINE - A cooling system for a gas turbine engine ( | 04-03-2014 |
| 20140090396 | COMBUSTOR WITH RADIALLY STAGED PREMIXED PILOT FOR IMPROVED - The present invention discloses a novel apparatus and method for a mixing fuel and air in a gas turbine combustion system. The mixer helps to mix fuel and air while being able to selectively increase the fuel flow to a shear to a shear layer of a pilot flame in order to reduce polluting emissions. The mixer directs a flow of air radially inward into the combustion system and includes two sets of fuel injectors within each radially-oriented vane. A first plurality of fuel injectors operate independent of a second plurality of fuel injectors and the second plurality of fuel injectors are positioned to selectively modulate the fuel flow to the shear layer of the resulting pilot flame. | 04-03-2014 |
| 20140102112 | ONE-PIECE FUEL NOZZLE FOR A THRUST ENGINE - A nozzle formed of one piece for a jet engine includes a mixing tube, a fuel conduit integrally formed with the mixing tube, and an opening through the fuel conduit and directed radially into the mixing tube. | 04-17-2014 |
| 20140109587 | SYSTEM AND METHOD FOR REDUCING MODAL COUPLING OF COMBUSTION DYNAMICS - A system and method for reducing combustion dynamics includes first and second combustors, and each combustor includes a fuel nozzle and a combustion chamber downstream from the fuel nozzle. Each fuel nozzle includes an axially extending center body, a shroud that circumferentially surrounds at least a portion of the axially extending center body, a plurality of vanes that extend radially between the center body and the shroud, a first fuel port through at least one of the plurality of vanes at a first axial distance from the combustion chamber, the plurality of vanes being located at a second axial distance from the combustion chamber. A second fuel port is provided through the center body at a third axial distance from the combustion chamber. The system further includes structure for producing a combustion instability frequency in the first combustor that is different from the combustion instability frequency in the second combustor. | 04-24-2014 |
| 20140109588 | BURNER FOR A CAN COMBUSTOR - The present invention relation to a burner for a combustion chamber of a gas turbine with a mixing and injection device. The mixing and injection device includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies, each extending in a first transverse direction into the gas-flow channel. Each streamlined body has two lateral surfaces that are arranged essentially parallel to the main-flow direction, the lateral surfaces being joined to one another at their upstream side to form a leading edge of the body and joined at their downstream side to form a trailing edge of the body. Each streamlined body has a cross-section perpendicular to the first transverse direction that is shaped as a streamlined profile. At least one of the streamlined bodies is provided with a mixing structure and with at least one fuel nozzle located at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel, wherein at least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor. The invention also relates to a method of using said burner in a gas turbine. | 04-24-2014 |
| 20140123668 | SYSTEM AND METHOD FOR DIFFUSION COMBUSTION WITH FUEL-DILUENT MIXING IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM - A system is provided with a turbine combustor having a first diffusion fuel nozzle, wherein the first diffusion fuel nozzle has first and second passages that separately inject respective first and second flows into a chamber of the turbine combustor to produce a diffusion flame. The first flow includes a first fuel and a first diluent, and the second flow includes a first oxidant. The system includes a turbine driven by combustion products from the diffusion flame in the turbine combustor. The system also includes an exhaust gas compressor, wherein the exhaust gas compressor is configured to compress and route an exhaust gas from the turbine to the turbine combustor along an exhaust recirculation path. | 05-08-2014 |
| 20140123669 | SYSTEM AND METHOD FOR DIFFUSION COMBUSTION WITH OXIDANT-DILUENT MIXING IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM - A system is provided with a turbine combustor having a first diffusion fuel nozzle, wherein the first diffusion fuel nozzle has first and second passages that separately inject respective first and second flows into a chamber of the turbine combustor to produce a diffusion flame. The first flow includes a first fuel, and the second flow includes a first oxidant and a first diluent. The system includes a turbine driven by combustion products from the diffusion flame in the turbine combustor. The system also includes an exhaust gas compressor, wherein the exhaust gas compressor is configured to compress and route an exhaust gas from the turbine to the turbine combustor along an exhaust recirculation path. | 05-08-2014 |
| 20140123670 | GAS TURBINE BURNER - The burner of a gas turbine includes a swirl generator and, downstream of it, a mixing tube. The swirl generator is defined by at least two walls facing one another to define a conical swirl chamber and is provided with nozzles arranged to inject a fuel and apertures arranged to feed an oxidiser into the swirl chamber. The burner includes a lance which extends along a longitudinal axis of the swirl generator and is provided with side nozzles for ejecting a fuel within the burner. The side nozzles have their axes inclined with respect to the axis of the lance and can be positioned along the axis of the burner. | 05-08-2014 |
| 20140123671 | COMBUSTOR AND METHOD OF SUPPLYING FUEL TO THE COMBUSTOR - A combustor ( | 05-08-2014 |
| 20140137565 | COMBINATION AIR ASSIST AND PILOT GASEOUS FUEL CIRCUIT - When starting a dual fuel turbine engine on liquid fuel, a flow of air assist into a combined pilot gaseous fuel and air assist tube is supplied. The velocity of the flow of air assist is increased as it is expelled through an outlet of the tube. The flow of air assist is directed into a first end of a pilot injector barrel. Additional air is drawn into the first end of the pilot injector barrel from an enclosure containing both the first end of the pilot injector barrel and the tube outlet. Liquid pilot fuel is supplied at a second end of the pilot injector barrel, and this fuel is atomized by the flow of the air assist and additional air. The atomized pilot liquid fuel may then be combusted, | 05-22-2014 |
| 20140137566 | COMBUSTOR AND METHOD OF SUPPLYING FUEL TO THE COMBUSTOR - A combustor ( | 05-22-2014 |
| 20140144152 | Premixer With Fuel Tubes Having Chevron Outlets - A premixer includes an air tube formed in a burner tube defining a longitudinal axis, and a coaxially disposed fuel tube with a turbulence enhancing chevron outlet. The fuel tube may include an exterior tube and an interior tube with the interior tube, the exterior tube or both having chevron outlets. The chevron outlets may be tapered and notched. | 05-29-2014 |
| 20140144153 | SYSTEM AND METHOD TO CONTROL A GAS TURBINE SUBJECT TO FUEL COMPOSITION VARIATION - A system and method control a gas turbine subject to fuel composition variation. The method includes operating a first effector to control the gas turbine based on fuel composition. The method also includes operating a second effector to maintain operation of the first effector within a first boundary limit, the second effector operation being initiated when the operating the first effector reaches a second boundary limit within the first boundary limit. | 05-29-2014 |
| 20140150444 | GENERATING POWER USING AN ION TRANSPORT MEMBRANE - In some implementations, a system may include a compressor, a heat exchanger and an ITM. The compressor is configured to receive an air stream and compress the air stream to generate a pressurized stream. The heat exchanger is configured to receive the pressured stream and indirectly heat the pressurized stream using heat from an oxygen stream from an Ion Transport Membrane (ITM). The ITM is configured to receive the heated pressurized stream and generate an oxygen stream and the non-permeate stream, wherein the non-permeate stream is passed to a gas turbine burner and the oxygen stream is passed to the heat exchanger. | 06-05-2014 |
| 20140150445 | SYSTEM AND METHOD FOR LOAD CONTROL WITH DIFFUSION COMBUSTION IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM - A system is provided with a turbine combustor having a first diffusion fuel nozzle, wherein the first diffusion fuel nozzle is configured to produce a diffusion flame. The system includes a turbine driven by combustion products from the diffusion flame in the turbine combustor. The system also includes an exhaust gas compressor, wherein the exhaust gas compressor is configured to compress and route an exhaust gas from the turbine to the turbine combustor along an exhaust recirculation path. In addition, the system includes a control system configured to control flow rates of at least one oxidant and at least one fuel to the turbine combustor in a stoichiometric control mode and a non-stoichiometric control mode, wherein the stoichiometric control mode is configured to change the flow rates and provide a substantially stoichiometric ratio of the at least one fuel with the at least one oxidant, and the non-stoichiometric control mode is configured to change the flow rates and provide a non-stoichiometric ratio of the at least one fuel with the at least one oxidant. | 06-05-2014 |
| 20140157787 | FUEL NOZZLE FOR GAS TURBINE - A gas turbine system includes a fuel nozzle. The fuel nozzle includes a first fluid conduit defining an oxidant passage, a second fluid conduit defining a first fuel passage, and a third fluid conduit surrounding the second fluid conduit and defining a second fuel passage. A first orifice is disposed on the second fluid conduit and is configured to fluidly couple the first fuel passage to the oxidant passage. A second orifice is disposed on the third fluid conduit and is configured to fluidly couple the second fuel passage to the oxidant passage. A first diameter of the first orifice is less than a second diameter of the second orifice. | 06-12-2014 |
| 20140157788 | FUEL NOZZLE FOR GAS TURBINE - A gas turbine system includes a fuel nozzle. The fuel nozzle has a first wall extending along an axis and defines a first fluid passage. A second wall surrounds the first wall and defines a second fluid passage. A third wall surrounds the second wall and defines a third fluid passage. The first and second fluid passages are configured to collectively direct a flow of air and fuel into a combustion region to produce a flame. The third fluid passage is configured to direct a diluent into the combustion region to adjust a combustion parameter of the flame. | 06-12-2014 |
| 20140165583 | BLEED VALVE OVERRIDE SCHEDULE ON OFF-LOAD TRANSIENTS - In one aspect, the present disclosure is directed to a method for controlling a position of a bleed valve of a gas turbine engine. The onset of an off-load transient may be determined. Values representative of the turbine rotor inlet temperature and the exhaust outlet temperature may be determined. Also, the amount of time elapsed since the onset of the off-load transient may be determined. Three provisional bleed valve command positions may be determined based on value representative of the turbine rotor inlet temperature, the value representative of the exhaust outlet temperature, and the amount of time elapsed, respectively. The provisional bleed valve command position associated with the lowest relative value may be selected. Then, the bleed valve position may be adjusted to match the selected bleed valve command position. | 06-19-2014 |
| 20140165584 | CRYOGENIC PUMP SYSTEM FOR CONVERTING FUEL - A system for converting liquid fuel into gaseous fuel is provided. The system may have a supply of liquid fuel. The system may also have a combustor, and one or more pumps in fluid communication with the supply. The one or more pumps may be configured to pump liquid fuel from the supply into the combustor. The system may also have a compressor in fluid communication with an inlet of the combustor, and a turbine in fluid communication with an outlet of the combustor. The turbine may be connected to drive the compressor and the one or more pumps. The system may also have a heat exchanger in fluid communication with an outlet of the turbine and an outlet of the one or more pumps. | 06-19-2014 |
| 20140165585 | Oblong Swirler Assembly for Combustors - In accordance with one aspect of the disclosure, a swirler is disclosed. The swirler may include an outer shroud and inner shroud. The inner shroud may be positioned radially inside the outer shroud. At least one of the outer shroud and inner shroud may have a major diameter which is larger than a minor diameter such that the shrouds define an oblong shape. The swirler may further include a plurality of vanes which may be positioned between the inner and outer shrouds. | 06-19-2014 |
| 20140174096 | METHOD AND ARRANGEMENT FOR INJECTING AN EMULSION INTO A FLAME - An arrangement for injection of an emulsion of a first fluid and a second fluid into a flame of a burner has a central gas duct, an outer gas channel disposed coaxially with the gas duct, and a fluid channel disposed coaxially between the gas duct and the outer gas channel The central gas duct and the fluid channel are separated by a first frustoconical wall. The fluid channel and the outer gas channel are separated by a second frustoconical wall. The arrangement is mounted concentrically surrounding a heat source which provides through the gas duct hot gases being directed into the flame of the burner. Further, the arrangement includes a mixing device for forming an emulsion of the first fluid and the second fluid, for supplying the emulsion into the fluid channel and for injecting the emulsion from the fluid channel into the flame. | 06-26-2014 |
| 20140190174 | MICROMIXER ASSEMBLY FOR A TURBINE SYSTEM AND METHOD OF DISTRIBUTING AN AIR-FUEL MIXTURE TO A COMBUSTOR CHAMBER - A micromixer assembly for a turbine system includes a plurality of pipes each having an inlet for receiving an airflow from an annulus defined by an inwardly disposed liner and an outwardly disposed sleeve, each of the plurality of pipes also including an outlet for dispersing an air-fuel mixture into a combustor chamber. Also included is a first portion of each of the plurality of pipes. Further included is a second portion of each of the plurality of pipes, the second portion comprising the inlet for receiving the airflow. Yet further included is at least one fuel receiving path in communication with at least one of the first portion and the second portion. | 07-10-2014 |
| 20140190175 | 3D NON-AXISYMMETRIC COMBUSTOR LINER - A combustor liner with an input end and an output end includes an annular inner wall and an annular outer wall. At least one of the inner wall and outer wall is three-dimensionally contoured. The inner wall and the outer wall form a combustion chamber with the contours creating alternating expanding and constricting regions inside the chamber causing combustion gases to flow in the circumferential and axial directions. | 07-10-2014 |
| 20140190176 | FUEL CONTROL METHOD AND FUEL CONTROL APPARATUS FOR GAS TURBINE AND GAS TURBINE - A fuel control method for a gas turbine with a combustor being formed of at least two groups of a pluralities of main nozzles for supplying fuel, and that supplies fuel from the main nozzles of all groups upon ignition of the combustor (S | 07-10-2014 |
| 20140190177 | METHOD FOR RUNNING UP A STATIONARY GAS TURBINE - A method for running up a stationary gas turbine is provided. The turbine has a combustion chamber, the burners of which have a pilot burner and a main burner and by which various types of fuel are introduced into the combustion chamber for burning. The method, carried out while a rotor of the gas turbine is accelerating from a standstill to a nominal speed, includes feeding fuel of a first type of fuel to the pilot burners and feeding fuel of the first type of fuel to the main burners. In order to provide a method in which a comparatively low supply pressure is required in the fuel supply system and in which combustion vibrations that put the machine at risk are avoided during the running-up process, it is proposed that fuel of a second type of fuel is fed to the burner before the nominal speed is reached. | 07-10-2014 |
| 20140190178 | COMBUSTOR - A combustion method and combustor for use in a jet engine, the jet engine having a compressor portion and a turbine portion. The combustor includes an outer tube having a central axis that extends longitudinally intermediate the compressor portion and turbine portion and is positioned to receive air discharged by the compressor portion. An inner tube is positioned within the outer tube that includes an associated outer surface spaced from the inner surface of the outer tube thereby defining a combustion chamber. The outer tube and inner tube include fluid directing structure for communicating at least some of the air discharged by the compressor portion to the combustion chamber. The fluid directing structure directs air into the combustion chamber in a direction offset from the central axis, thereby causing rotation or swirling of the air about the central axis. | 07-10-2014 |
| 20140196465 | LEAN-RICH AXIAL STAGE COMBUSTION IN A CAN-ANNULAR GAS TURBINE ENGINE - An apparatus and method for lean/rich combustion in a gas turbine engine ( | 07-17-2014 |
| 20140196466 | METHODS AND SYSTEMS FOR OPERATING GAS TURBINE ENGINES - Methods and systems for operating a gas turbine engine including a fuel delivery system and a plurality of combustor assemblies are provided. The fuel delivery system comprises a primary fuel circuit configured to continuously supply fuel to each of the plurality of combustor assemblies during a first mode of operation and a second mode of operation. At least one secondary fuel circuit of the fuel delivery system is configured to supply fuel to each of the plurality of combustor assemblies during the second mode of operation. The secondary fuel circuit includes at least one isolation valve coupled in flow communication with each of the plurality of combustor assemblies. The at least one isolation valve facilitates preventing fluid flow upstream into the secondary fuel circuit during the first mode of operation. The fuel delivery system, using the isolation valve, replaces a purging system in the gas turbine engine. | 07-17-2014 |
| 20140196467 | OXIDIZING FUEL - A mixture of air and fuel is received into a reaction chamber of a gas turbine system. The fuel is oxidized in the reaction chamber, and a maximum temperature of the mixture in the reaction chamber is controlled to be substantially at or below an inlet temperature of a turbine of the gas turbine system. The oxidation of the fuel is initiated by raising the temperature of the mixture to or above an auto-ignition temperature of the fuel. In some cases, the reaction chamber may be provided without a fuel oxidation catalyst material. | 07-17-2014 |
| 20140216048 | Variable Volume Combustor - The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner. | 08-07-2014 |
| 20140216049 | Variable Volume Combustor with Pre-Nozzle Fuel Injection System - The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system. | 08-07-2014 |
| 20140216050 | Variable Volume Combustor with Nested Fuel Manifold System - The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles, a fuel manifold system in communication with the micro-mixer fuel nozzles to deliver a flow of fuel thereto, and a linear actuator to maneuver the micro-mixer fuel nozzles and the fuel manifold system. | 08-07-2014 |
| 20140216051 | Variable Volume Combustor with an Air Bypass System - The present application provides a combustor for use with flow of fuel and a flow of air in a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles positioned within a liner and an air bypass system position about the liner. The air bypass system variably allows a bypass portion of the flow of air to bypass the micro-mixer fuel nozzles. | 08-07-2014 |
| 20140216052 | IGNITION DEVICE AND METHOD FOR A TURBOMACHINE COMBUSTION CHAMBER - A method and device for igniting a turbomachine combustion chamber including alternately: an intake phase of a fluid into a chamber through an intake port, during which a piston compresses an elastic mechanism under pressure of the fluid such that the elastic mechanism applies onto a piezoelectric element a force sufficient for the piezoelectric element to induce between the electrodes an electric voltage enabling an electric arc to be generated, until the piston reaches a predetermined position for closing a valve for sealing the intake port; and an exhaust phase of the fluid, during which the elastic mechanism pushes back the piston to induce a fluid ejection out of the chamber through an exhaust port, and the valve is open. | 08-07-2014 |
| 20140230447 | FUEL NOZZLE FOR A GAS TURBOMACHINE - A fuel nozzle for a gas turbomachine includes an outer nozzle body including an inner surface defining a mixing zone, and an inner nozzle body arranged within the outer nozzle body. The inner nozzle body includes a fluid passage. At least one flow affector extends from the inner nozzle body to the outer nozzle body. The at least one flow affector includes an inner surface that defines an interior chamber having an inlet fluidly connected to the fluid passage and at least two openings fluidically linking the interior chamber and the mixing zone. One or more flow tuning elements are arranged at the interior chamber upstream of the at least two openings. The one or more flow affectors are configured and disposed to condition a fluid passing into the interior chamber to affect a substantially iso-kinetic distribution of the fluid within the interior chamber. | 08-21-2014 |
| 20140230448 | METHOD FOR PREVENTING FLASHBACK IN A BURNER HAVING AT LEAST ONE SWIRL GENERATOR - A method for preventing flashback in a burner having swirl generator with a central fuel distributor element and an outer wall enclosing the central fuel distributor element and bounding an axial flow channel for combustion air is provided. A separating wall encloses the central fuel distributor element radially within the outer wall, wherein the axial flow channel is divided into a radially inner channel segment and a radially outer segment by the separating wall. A tangential flow component in the radially outer channel segment is provided to the combustion air flowing through the axial flow channel, wherein, during combustion, the combustion air passes the radially inner channel segment without a tangential flow component. | 08-21-2014 |
| 20140230449 | VALVE CONTROL DEVICE, GAS TURBINE, AND VALVE CONTROL METHOD - A valve control device is provided in a gas turbine having a combustor for generating combustion gas, a turbine driven by the combustion gas generated by the combustor, a flow rate regulating valve for regulating the flow rate of the fuel to be supplied to the combustor, and a pressure regulating valve disposed upstream of the flow rate regulating valve, for regulating the fuel pressure. The valve control device controls the opening degree of the valve. The valve control device includes a load decrease detection part which detects a load decrease of the gas turbine, and a pressure control part which controls the opening degree of the valve in accordance with the output of the gas turbine. The valve control device suppresses instability of the gas turbine output even when the load rapidly decreases. | 08-21-2014 |
| 20140238036 | FUEL/AIR MIXING SYSTEM FOR FUEL NOZZLE - A fuel nozzle includes an inner wall defining a central passage extending in an axial direction of the fuel nozzle, a hub wall surrounding the inner wall and defining a first annular passage, an outer wall surrounding the hub wall and defining a second annular passage, and a shroud surrounding the outer wall and defining a third annular passage. A swirler may receive air and direct the air into the first annular passage. The swirler includes at least one swirl vane extending from the shroud to the hub wall that has an air passage extending between the shroud and the hub wall. The air passage is coupled to the first annular passage and has a first width adjacent the shroud and a second width adjacent the hub wall. The second width is larger than the first width defining a diverging outlet into the first annular passage. | 08-28-2014 |
| 20140238037 | GAS TURBINE ENGINE AND METHOD FOR OPERATING A GAS TURBINE ENGINE - One embodiment of the present disclosure is a unique method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft. Another embodiment of the present disclosure is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. | 08-28-2014 |
| 20140245744 | CONTROL OF GAS TURBINE ENGINE - Systems, devices, and methods for controlling a fuel supply for a turbine or other engine using direct and/or indirect indications of power output and optionally one or more secondary control parameters. | 09-04-2014 |
| 20140245745 | FUEL SYSTEM OF GAS TURBINE ENGINES - A method for purging fuel from a fuel system of a gas turbine engine on shutdown of the engine comprises, in one aspect, terminating a fuel supply to the fuel system and using the residual compressed air to create a reversed pressure differential in the fuel system relative to a forward pressure differential of the fuel system used to maintain fuel supply for engine operation, and under the reversed pressure differential substantially purging the fuel remaining in the system therefrom to a fuel source. | 09-04-2014 |
| 20140250907 | GAS TURBINE FUEL INJECTOR WITH METERING CAVITY - A fuel injector for a gas turbine engine may include a flow path for a fuel-air mixture extending longitudinally through the fuel injector, and a fuel gallery extending circumferentially around the flow path. The fuel gallery may be adapted to inject a liquid fuel into the flow path. The fuel injector may also include an annular casing positioned circumferentially around the fuel gallery to define an insulating chamber around the gallery. The fuel injector may also include an annular cover extending around the fuel injector to define a metering chamber. The fuel injector may further include one or more purge holes fluidly coupling the metering chamber to the insulating chamber, and one or more metering holes fluidly coupling the metering chamber to a volume exterior to the fuel injector. | 09-11-2014 |
| 20140250908 | Systems and Methods for Controlling Combustion of a Fuel - Systems and methods for controlling the composition of a combustion exhaust gas are provided. | 09-11-2014 |
| 20140260296 | SLINGER COMBUSTOR - A slinger combustor has an annular combustor shell defining a combustion chamber having a radially inner fuel inlet for receiving a spray of fuel centrifuged by a fuel slinger. The combustion chamber has a fuel atomization zone extending radially outwardly from the fuel inlet and merging into a radially outwardly flaring expansion zone leading to a combustion zone. A plurality of nozzle air inlets are defined in the fuel atomization zone of the combustor shell. The nozzle air inlets have a nozzle axis intersecting the stream of fuel and a tangential component in a direction of rotation of the fuel slinger. A plurality of dilution holes are defined in the combustor shell and have a dilution axis intersecting the combustion zone. The dilution axis of at least some of the dilution holes has a tangential component opposite to the direction of rotation of the fuel slinger. | 09-18-2014 |
| 20140260297 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor comprises an annular combustor chamber formed between the inner and outer liners, the annular combustor chamber having a central axis. Fuel nozzles are in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber. The fuel nozzles are oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. Nozzle air inlets are in fluid communication with the annular combustor chamber to inject nozzle air generally radially in the annular combustor chamber. A plurality of dilution air holes are defined through the inner and outer liner downstream of the nozzle air inlets, the dilution holes configured for high pressure air to be injected from an exterior of the liners through the dilution air holes generally radially into the combustor chamber, a central axis of the dilution air holes having a tangential component relative to the central axis of the annular combustor chamber. | 09-18-2014 |
| 20140260298 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor comprises an annular combustor chamber formed between the inner and outer liners. Fuel nozzles each have an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. A plurality of nozzle air holes are defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles. The nozzle air holes are configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber. A central axis of the nozzle air holes has a tangential component relative to the central axis of the annular combustor chamber. | 09-18-2014 |
| 20140260299 | FUEL-AIR MIXING SYSTEM FOR GAS TURBINE SYSTEM - Embodiments of the present disclosure are directed to systems and methods for premixing fuel and air prior to combustion within a combustion chamber. The system includes a plurality of fuel injectors and a plurality of mixing tubes, wherein each mixing tube has a first portion for receiving one of the plurality of fuel injectors and a second portion having a mixing chamber that is configured to mix fuel and air. The length of the mixing chamber varies among the plurality of mixing tubes to allow for different mixing times. | 09-18-2014 |
| 20140260300 | AIR DIFFUSER FOR COMBUSTOR - A system includes a multi-tube fuel nozzle of a turbine combustor. The multi-tube fuel nozzle includes a support structure defining an interior volume configured to receive an air flow; a plurality of mixing tubes disposed within the interior volume, wherein each of the plurality of mixing tubes comprises a respective fuel injector; and an outer annular wall configured to direct an air flow from an annulus between a liner and a flow sleeve of the turbine combustor at least partially radially inward into the interior volume through an air inlet and toward the plurality of mixing tubes, wherein the outer annular wall at least partially defines an air flow passage extending from the annulus to the interior volume. | 09-18-2014 |
| 20140260301 | ENGINE MANIFOLD DRAIN SYSTEM - A fuel system for a gas turbine engine includes an engine fuel manifold, a hydraulic actuator, and a drain piston assembly. The hydraulic actuator actuates in response to a change in pressures within the hydraulic actuator. The drain piston assembly is fluidically connected to both the hydraulic actuator and the engine fuel manifold. The drain piston assembly receives fuel from the engine fuel manifold and sends fuel to the hydraulic actuator during engine shut down. | 09-18-2014 |
| 20140260302 | DIFFUSION COMBUSTOR FUEL NOZZLE FOR LIMITING NOx EMISSIONS - The present application and the resultant patent provide a diffusion combustor fuel nozzle for a gas turbine engine. The fuel nozzle may include one or more gas fuel passages for one or more flows of gas fuel, a swirler surrounding the one or more gas fuel passages and positioned about a downstream face of the fuel nozzle, a number of swirler gas fuel ports defined in the swirler, and a number of downstream face gas fuel ports defined in the downstream face of the fuel nozzle. The swirler may include a number of swirl vanes and a number of air chambers defined between adjacent swirl vanes. The present application and the resultant patent further provide a method of operating a diffusion combustor fuel nozzle of a gas turbine engine. | 09-18-2014 |
| 20140260303 | METHODS RELATING TO DOWNSTREAM FUEL AND AIR INJECTION IN GAS TURBINES - A method for use in a gas turbine engine. The method includes the steps of: configuring a downstream injection system within the interior flowpath that includes two injection stages, a first stage and a second stage, wherein the first stage and the second stage are each axially spaced from the other; and circumferentially positioning the injectors of the first stage and the second stage based on: a) a characteristic of an anticipated combustion flow occurring just upstream of the first stage during a mode of operation; and b) the characteristic of an anticipated combustion flow just downstream of the second stage given an anticipated effect of the air and fuel injection from the first stage and the second stage. | 09-18-2014 |
| 20140260304 | ALGAE-DERIVED FUEL/WATER EMULSION - A method including providing a wet algae material having been subjected to a refinement process without a water separation phase; supplying the wet algae material including a water fraction and an algae-grown biofuel to a turbine engine; and operating the turbine engine with the wet algae material where a retained portion of the water fraction is retained in the wet biofuel during a manufacturing process not including a dehydration step and includes an amount sufficient to reduce generation of a quantity of nitrogen oxides, and where the turbine engine can further include a combustor capable of combusting the wet biofuel. | 09-18-2014 |
| 20140260305 | LEAN AZIMUTHAL FLAME COMBUSTOR - A combustion chamber may include a first surface and a second surface interconnected by a wall forming a chamber having a central axis. The first surface may define an exhaust opening and the second surface defining a pilot opening, wherein the exhaust opening and the pilot opening align along the central axis. A plurality of inlet ports may be configured to deliver air to the chamber. A plurality of fuel ports may be arranged on an inside of the second surface to deliver fuel to the chamber. The air flow from the inlet ports and fuel from the fuel ports may oppose each other to create a vortex of product proximal to the second surface. | 09-18-2014 |
| 20140283524 | NOZZLE SYSTEM AND METHOD FOR STARTING AND OPERATING GAS TURBINES ON LOWBTU FUELS - A fuel nozzle system for enabling a gas turbine to start and operate on low-Btu fuel includes a primary tip having primary fuel orifices and a primary fuel passage in fluid communication with the primary fuel orifices, and a fuel circuit capable of controlling flow rates of a first and second low-Btu fuel gases flowing into the fuel nozzle. The system is capable of operating at an ignition status, in which at least the first low-Btu fuel gas is fed to the primary fuel orifices and ignited to start the gas turbine, and a baseload status, in which at least the second low-Btu fuel gas is fired at baseload. The low-Btu fuel gas ignited at the ignition status has a content of the first low-Btu fuel gas higher than that of the low-Btu fuel gas fired at the baseload status. Methods for using the system are also provided. | 09-25-2014 |
| 20140283525 | TWO-BRANCH MIXING PASSAGE AND METHOD TO CONTROL COMBUSTOR PULSATIONS - A gas turbine engine combustion system including a mixing duct that separates into at least two branch passages for the delivery of a fuel and working fluid to distinct locations within a combustion chamber. The residence time for the fuel and working fluid within each of the two branch passages is distinct. | 09-25-2014 |
| 20140290266 | FUEL AND ACTUATION SYSTEM FOR GAS TURBINE ENGINE - A fuel system for an aircraft comprises a boost pump, a main fuel pump and a motive pump. The boost pump receives fuel from a storage unit. The main fuel pump receives fuel from the boost pump and delivers fuel to a distribution system. The motive fuel pump receives fuel from the boost pump, routes fuel through the storage unit, and delivers fuel to an actuator. A method for delivering fuel in an aircraft comprises pumping fuel from a fuel tank to a distribution system using a main pump, pumping fuel from a fuel tank to an actuator using a motive pump, and routing fuel from the actuator to the main pump | 10-02-2014 |
| 20140298817 | ARRANGEMENT FOR PREPARATION OF LIQUID FUEL FOR COMBUSTION AND A METHOD OF PREPARING LIQUID FUEL FOR COMBUSTION - A preparation of liquid fuel for combustion, particularly for ignition, in the burners of a gas turbine is proposed. A source of heated liquid fuel, particularly heated pressurized fuel, is arranged in a reservoir. The reservoir is arranged in parallel with a main fuel feed of a combustion system. The reservoir can be filled by the main fuel feed and/or evacuated into the main fuel feed. The reservoir contains a heating means, which heat the liquid fuel filled in the reservoir, particularly while circulating in the reservoir. | 10-09-2014 |
| 20140298818 | CONTROL METHOD AND CONTROL DEVICE FOR LEAN FUEL INTAKE GAS TURBINE - A method and a device for controlling a lean fuel intake gas turbine engine, which can stably maintain operation of the gas turbine engine by preventing misfire and burnout of a catalytic combustor even when a catalyst in the combustor is deteriorated. A difference between measured temperatures of an inlet and an outlet of the combustor with respect to a methane concentration in an intake gas that is to be taken into the lean fuel intake gas turbine engine is compared with reference temperature difference data that is data indicating difference between temperatures of the inlet and the outlet of the combustor including a catalyst in its initial state to be a reference, and at least one of the inlet temperature and the outlet temperature of the combustor is controlled based on a difference between the measured temperature difference and the reference temperature difference data. | 10-09-2014 |
| 20140298819 | TURBINE ENGINE COMPRISING AN ELECTRICALLY ACTIVATED FUEL SUPPLY PUMP, AND TURBINE ENGINE FUEL SUPPLY METHOD - A turbine engine for an aircraft including a turbine engine shaft and a pumping module, including: a pump shaft, connected to the turbine engine shaft; a pump for supplying fuel to the turbine engine, mounted on the pump shaft, configured to deliver a flow of fuel as a function of a speed of rotation of the turbine engine shaft; and an electrical device mounted on the pump shaft and configured, according to a first mode of operation, to drive the pump shaft in rotation to actuate the supply pump and, according to a second mode of operation, to be driven in rotation by the pump shaft to supply electrical power to equipment of the turbine engine. | 10-09-2014 |
| 20140318145 | HYBRID SLINGER COMBUSTION SYSTEM - There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand. | 10-30-2014 |
| 20140331685 | GAS DOSAGE CONTROL FOR GAS ENGINE - A gas engine assembly includes a compressor, a combustion system, a bypass line and a control system. The control system is configured to control gas supply parameters based on a transportation delay value. The transportation delay value corresponds to a delay between a time when a gas supply control mechanism is adjusted and a time that gas having a corresponding adjustment of a gas characteristic is received at a predetermined point downstream from the gas supply control mechanism. | 11-13-2014 |
| 20140338354 | System Having a Multi-Tube Fuel Nozzle with an Inlet Flow Conditioner - A system including a multi-tube fuel nozzle, including a plurality of tubes extending in an axial direction relative to a central axis of the multi-tube fuel nozzle, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet; and an inlet flow conditioner, including a plate extending in a radial direction relative to the central axis of the multi-tube fuel nozzle; an outer wall extending circumferentially about the plate, wherein the outer wall is coupled to the plate; and a plurality of air openings in the plate, the outer wall, or a combination thereof, wherein the plurality of air openings are disposed upstream from the air inlets in the plurality of tubes. | 11-20-2014 |
| 20140338355 | System and Method for Sealing a Fuel Nozzle - A system including a multi-tube fuel nozzle, including a first plate having a first plurality of openings, a plurality of tubes extending through the first plurality of openings in the first plate, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, and a resilient metallic seal disposed along the first plate about the plurality of tubes. | 11-20-2014 |
| 20140338356 | System Having a Multi-Tube Fuel Nozzle with an Aft Plate Assembly - A system including a first multi-tube fuel nozzle including a plurality of first tubes extending in an axial direction, wherein each first tube of the plurality of first tubes includes a first air inlet, a first fuel inlet, and a first fuel-air mixture outlet, a second multi-tube fuel nozzle including a plurality of second tubes extending in an axial direction, wherein each second tube of the plurality of second tubes includes a second air inlet, a second fuel inlet, and a second fuel-air mixture outlet, and an aft plate including a plurality of first tube apertures and a plurality of second tube apertures, wherein the plurality of first tubes extend to the plurality of first tube apertures, and the plurality of second tubes extend to the plurality of second tube apertures. | 11-20-2014 |
| 20140338357 | CAVITY SWIRL FUEL INJECTOR FOR AN AUGMENTOR SECTION OF A GAS TURBINE ENGINE - A fuel injection system for a gas turbine engine includes a fuel nozzle with a fuel injection aperture to inject a fuel jet and a multiple of airflow passages in the fuel nozzle to communicate a multiple of air streams to interact with the fuel jet. | 11-20-2014 |
| 20140338358 | INTERNAL DETONATION ENGINE, HYBRID ENGINES INCLUDING THE SAME, AND METHODS OF MAKING AND USING THE SAME - Hybrid internal detonation-gas turbine engines incorporating detonation or pulse engine technology (such as an internal detonation engine), and methods of manufacturing and using the same are disclosed. The internal detonation engine includes a detonation chamber having a fuel igniter therein, a stator at one end of the detonation chamber having at least a first opening to receive fuel, a rotor adjacent to the stator, and an energy transfer mechanism configured to convert energy from igniting or detonating the fuel to mechanical energy. The detonation chamber and fuel igniter are configured to ignite or detonate a fuel in the detonation chamber. Either the stator or the detonation chamber has a second opening to exhaust detonation gas(es). The rotor has one or more third openings therein configured to overlap with at least the first opening as the rotor rotates. | 11-20-2014 |
| 20140338359 | COMBUSTOR AND METHOD FOR SUPPLYING FUEL TO A COMBUSTOR - A combustor ( | 11-20-2014 |
| 20140345291 | Toroidal Combustion Chamber - A device comprising a combustion toroid for receiving combustion-induced centrifugal forces therein to continuously combust fluids located therein and an outlet for exhaust from said combustion toroid. | 11-27-2014 |
| 20140352321 | GAS TURBINE ENGINE SYSTEM AND AN ASSOCIATED METHOD THEREOF - A gas turbine engine system includes a compressor, a combustor, and a turbine. The combustor is coupled to the compressor and disposed downstream of the compressor. The combustor includes a secondary combustor section coupled to a primary combustor section and disposed downstream of the primary combustor section. The combustor also includes a transition nozzle coupled to the secondary combustor section and disposed downstream of the secondary combustor section. The combustor further includes an injector coupled to the secondary combustor section, for injecting an air-fuel mixture to the secondary combustor section. The turbine is coupled to the combustor and disposed downstream of the transition nozzle; wherein the transition nozzle is oriented substantially tangential to the turbine. | 12-04-2014 |
| 20140352322 | ANNULAR STRIP MICRO-MIXERS FOR TURBOMACHINE COMBUSTOR - A turbomachine combustor is provided. The turbomachine includes a combustion chamber and multiple micro-mixer nozzles arranged concentrically within a radial combustion liner and configured to receive fuel from one or more fuel supply pipes affixed to each of the plurality of micro-mixer nozzles at an upstream face. The multiple micro-mixer nozzle are also configured to receive air from a flow sleeve surrounding the radial combustion liner. Each of the micro-mixer nozzles include an annular strip having a multiple tubes or passages extending axially from the upstream face to a downstream face of each of the micro-mixer nozzles. | 12-04-2014 |
| 20140360202 | FUEL INJECTOR AND A COMBUSTION CHAMBER - A fuel injector includes pilot and main fuel injectors. The pilot fuel injector includes at least one pilot air swirler and the main fuel injector includes a main air blast fuel injector located between inner main and outer main air swirlers. A first air splitter is located between the pilot and inner main air swirlers and a second air splitter is located between the pilot and inner main air swirlers. The first air splitter has a downstream portion converging to a downstream end. The second air splitter has a downstream portion diverging to a downstream end. The second air splitter downstream end is downstream of the first air splitter downstream end and the ratio of the distance from the first air splitter downstream end to the second air splitter downstream end to the diameter of the second air splitter downstream end is in the range of 0.22 to 0.30. | 12-11-2014 |
| 20140360203 | RIJKE TYPE COMBUSTION ARRANGEMENT AND METHOD - A process of producing heat energy for use in heat exchange with other fluids and substances so as to impart said heat energy to said fluid or substances which includes several steps. In a first step, a mixture of fuel and oxidant is ignited in a combustion zone or zones ( | 12-11-2014 |
| 20140366551 | CONTINUOUS IGNITION - An ignition system includes a housing defining an interior and an exhaust outlet. The housing is configured and adapted to be mounted to a combustor to issue flame from the exhaust outlet into the combustor for ignition and flame stabilization within the combustor. A fuel injector is mounted to the housing with an outlet of the fuel injector directed to issue a spray of fuel into the interior of the housing. An igniter is mounted to the housing with an ignition point of the igniter proximate the outlet of the fuel injector for ignition within the interior of the housing. | 12-18-2014 |
| 20150013342 | AIR FLOW CONDITIONER FOR FUEL INJECTOR OF GAS TURBINE ENGINE - A fuel injector for a gas turbine engine is provided. The fuel injector includes a central body, an air inlet duct, a mixing duct, a swirler, and a flow conditioner. The air inlet duct and the mixing duct are positioned around the central body to define an air flow passage. The swirler is positioned between the air inlet duct and the mixing duct. The flow conditioner is disposed in the air flow passage upstream with respect to the swirler. The flow conditioner has a perforated plate configured to uniformly distribute air circumferentially within the air inlet duct. | 01-15-2015 |
| 20150033752 | GAS TURBINE COMBUSTION SYSTEM AND METHOD OF FLAME STABILIZATION IN SUCH A SYSTEM - A gas turbine combustion system is provided. In an embodiment, the system includes a first radial inflow swirler having first radial outer intake openings, first radial inner outlet openings and first flow passages, each first flow passage including a first angle (a) with respect to the radial direction, a second radial inflow swirler having second radial outer intake openings, each second flow passage including a second angle (p) with respect to the radial direction, where the radial outer circumference of the second radial inflow swirler has a diameter that is smaller than the diameter of the radial inner circumference of the first radial inflow swirler and the second radial inflow swirler is located coaxially with and radially inside the first radial inflow swirler. The first angle (a) has a different sign than the second angle (p) with respect to the radial direction. | 02-05-2015 |
| 20150040575 | FUEL IGNITER ASSEMBLY HAVING HEAT-DISSIPATING ELEMENT AND METHODS OF USING SAME - A combustor for a gas turbine engine includes a combustion chamber and a fuel igniter assembly. The combustion chamber is defined by an annular inner combustor liner and an annular outer combustor liner. The fuel igniter assembly is coupled to the combustor and extends radially outward from the outer combustor liner. The fuel igniter assembly includes an igniter housing configured to house a fuel igniter therein, and a heat-dissipating element coupled to the igniter housing. The heat-dissipating element includes a plurality of fins configured to dissipate heat from the fuel igniter assembly. | 02-12-2015 |
| 20150040576 | COUNTER SWIRL DOUBLET COMBUSTOR - An improved combustor for a gas turbine is operable to provide high combustion efficiency in a compact combustion chamber. The combustor includes a counter swirl doublet for improved fuel/air mixing. The enhanced combustor assembly and method of operation improves operation of the turbine. | 02-12-2015 |
| 20150047365 | SEQUENTIAL COMBUSTION WITH DILUTION GAS MIXER - The invention refers to a sequential combustor arrangement including a first burner, a first combustion chamber, a dilution burner for admixing a dilution gas and a second fuel via a dilution-gas-fuel-admixer to the first combustor combustion products. The dilution-gas-fuel-admixer has at least one streamlined body which is arranged in the dilution burner for introducing the at least one second fuel into the dilution burner through at least one fuel nozzle, and which has a streamlined cross-sectional profile and which extends with a longitudinal direction perpendicularly or at an inclination to a main flow direction prevailing in the dilution burner. The streamlined body includes a dilution gas opening for admixing dilution gas into the first combustor combustion products upstream of the at least one fuel nozzle. The disclosure further refers to a method for operating a gas turbine with such a sequential combustor arrangement. | 02-19-2015 |
| 20150052905 | Pulse Width Modulation for Control of Late Lean Liquid Injection Velocity - Systems and methods for pulse-width modulation of late lean liquid injection velocity can be provided by certain embodiments of the disclosure. In one embodiment, a gas turbine combustor utilizing a late lean injection scheme can be provided, wherein the combustor can include a combustor liner and a transition piece. Methods described herein can allow for dynamic and intelligent adjustment of the late lean injection scheme based on a duty cycle and, optionally, a measured combustion gases temperature profile. The adjustments can involve a pulse-width modification of the duty cycle, which in turn can affect a fuel introduction velocity. Dynamic control of the fuel introduction velocity can provide for improved fuel droplet penetration and moving the heat release zone away from walls of the transitional piece. | 02-26-2015 |
| 20150059352 | DUAL FUEL COMBUSTOR FOR A GAS TURBINE ENGINE - The present application and the resultant patent provide a dual fuel combustor for a gas turbine engine. The combustor may include a primary premixer positioned within a head end plenum of the combustor, and a dual fuel, injection system positioned within the head end plenum and upstream of the premixer. The injection system may be configured to inject a gas fuel about an inlet end of the premixer when the combustor operates on the gas fuel. The injection system also may be configured to vaporize and inject a liquid fuel about the inlet end of the premixer when the combustor operates on the liquid fuel. The present application and the resultant patent also provide a related method of operating a dual fuel combustor. | 03-05-2015 |
| 20150075174 | Combustor and Method of Fuel Supply and Converting Fuel Nozzle for Advanced Humid Air Turbine - A fuel control device and method of a gas turbine combustor, for advanced humid air turbines, in which plural combustion units comprising plural fuel nozzles for supplying fuel and plural air nozzles for supplying air for combustion are provided. A part of the plural combustion units are more excellent in flame stabilizing performance than the other combustion units. A fuel ratio, at which fuel is fed to the part of the combustion units is set on the basis of internal temperature of the humidification tower and internal pressure of the humidification tower to control a flow ratio of the fuel fed to the plural combustion units. | 03-19-2015 |
| 20150082800 | METHOD FOR SUPPRESSING GENERATION OF YELLOW PLUM OF COMPLEX THERMAL POWER PLANT USING HIGH THERMAL CAPACITY GAS - There is provided a method for suppressing a generation of a yellow plume from a complex thermal power plant, the method being characterized in that, in a complex thermal power generating method including combusting fuel and compressed air for combustion, supplied to a combustor, to generate exhaust gas; generating power using the exhaust gas generated in the combusting; recovering heat of the exhaust gas by a heat recovery steam generator (HRSG) and generating power using the recovered heat and a steam turbine, and controlling an amount of supplied high thermal capacity gas supplying the high thermal capacity gas together with the fuel in the combusting, in such a manner that nitrogen dioxide is contained in the exhaust gas in an amount of 10 ppm or less (based on exhaust gas containing an oxygen concentration of 15%). | 03-26-2015 |
| 20150082801 | GAS TURBINE AND METHOD TO OPERATE THE GAS TURBINE - It is proposed a gas turbine and a method to operate the gas turbine. A fluid is supplied to the gas turbine by means of fluid paths. A first fluid path is divided into a second and third fluid path. A first control valve controls a fluid mass flow in the first fluid path. A second control valve controls a ratio of fluid mass flows in the second and third fluid paths. | 03-26-2015 |
| 20150089952 | INCLINED FUEL INJECTION OF FUEL INTO A SWIRLER SLOT - A combustor for a gas turbine is provided. The combustor includes a pre-combustion chamber having a centre axis and a swirler which is mounted to the pre-combustion chamber. The swirler surrounds the pre-combustion chamber in a circumferential direction with respect to the centre axis. The swirler has a bottom surface which forms a part of a slot through which oxidant/fuel mixture is injectable into the pre-combustion chamber, wherein the bottom surface is located in a bottom plane. The swirler further includes a fuel injector which is arranged to the bottom surface such that a fuel is injectable into the slot with a fuel injection direction, wherein a first component of the fuel injection direction is non-parallel to the normal (n) of the bottom plane. | 04-02-2015 |
| 20150089953 | SYSTEMS AND METHODS FOR PREVENTING FUEL LEAKAGE IN A GAS TURBINE ENGINE - Systems and methods for preventing fuel leakage in a gas turbine engine are provided. A fuel accumulation system includes a control valve section fluidly coupled to a fuel manifold passage and an accumulator valve section fluidly coupled at a first side to the control valve section. The control valve section is configured to control expansion of a fluid flowing in the fuel manifold passage. The accumulator valve section is configured to receive fluid expanded in the fuel manifold passage via the control valve section. | 04-02-2015 |
| 20150089954 | BURNERS HAVING FUEL PLENUMS - Burners having a fuel plenum in a base are disclosed. One disclosed example apparatus includes a base of a burner, the base comprising a fuel plenum and coupled to fuel nozzles, where at least one of the fuel nozzles is in fluid communication with the fuel plenum. The disclosed example apparatus also includes a burner head of the burner comprising nozzle passages in fluid communication with an airflow path, where the burner head defines a pilot combustion space that opens towards a flame tube of the burner, and is in fluid communication with the airflow path, and where each nozzle passage is to receive a fuel nozzle to provide fuel to entrain with air from the airflow path. | 04-02-2015 |
| 20150101342 | ENGINE - The present disclosure relates to an engine having two modes of operation—air breathing and rocket—that may be used in aerospace applications such as in an aircraft, flying machine, or aerospace vehicle. The engine's efficiency can be maximized by using a precooler arrangement to cool intake air in air breathing mode using cold fuel delivery systems used for the rocket mode. By introducing the precooler and certain other engine cycle components, and arranging and operating them as described, problems such as those associated with higher fuel and weight requirements and frost formation can be alleviated. | 04-16-2015 |
| 20150107259 | GAS TURBINE WITH SEQUENTIAL COMBUSTION ARRANGEMENT - The present disclosure refers to a method for operating a gas turbine with sequential combustors having a first-burner, a first combustion chamber, and a second combustor arranged sequentially in a fluid flow connection. To minimize emissions and combustion stability problems during transient changes when the fuel flow to a second combustor is initiated the method includes the steps of increasing the second fuel flow to a minimum flow, and reducing the first fuel flow to the first-burner of the same sequential combustor and/or the fuel flow to at least one other sequential combustor of the sequential combustor arrangement in order keep the total fuel mass flow to the gas turbine substantially constant. Besides the method a gas turbine with a fuel distribution system configured to carry out such a method is disclosed. | 04-23-2015 |
| 20150107260 | GAS TURBINE AND GAS TURBINE AFTERBURNER - A gas turbine afterburner includes a gutter electrode that helps to hold an afterburner flame. A charge source applies a majority charge to be carried by a turbine exhaust gas. Electrical attraction between the majority charge and the gutter electrode helps to hold the afterburner flame. | 04-23-2015 |
| 20150113998 | Gas Turbine Combustor and Gas Turbine Combustor Control Method - The burners include a central burner and a plurality of outer burners disposed around the central burner. Each of the outer burners is equipped with a fuel supply system that includes a fuel flow regulating valve. The outer circumference of the combustor liner is provided with a cylindrical flow sleeve. At least one flow velocity measurement unit is disposed in a circular flow path formed between the combustor liner and the flow sleeve to measure the flow velocity of air flowing downward. The gas turbine combustor also includes a control device that adjusts the fuel flow rate of the fuel, which is to be supplied to the outer burners, in accordance with the flow velocity of the air in the circular flow path, which is measured by the flow velocity measurement units. | 04-30-2015 |
| 20150128606 | Combustion Casing Manifold for High Pressure Air Delivery to a Fuel Nozzle Pilot System - The present application provides a pilot manifold system for a combustor of a gas turbine engine. The pilot manifold system may include a casing with a casing manifold, an end cover connected to the casing and having an end cover passage in communication with the casing manifold, and a fuel nozzle mounted about the end cover. The fuel nozzle may include a pilot system in communication with the end cover passage. | 05-14-2015 |
| 20150128607 | Multi-Swirler Fuel/Air Mixer with Centralized Fuel Injection - A gas turbine combustor assembly has a fuel/air mixer assembly with a plurality of fuel/air mixer elements. Each fuel/air mixer element defines an air flow passage therethrough. A fuel injector is coupled to the fuel/air mixer assembly. The fuel injector has a tip portion with a plurality of fuel outlets arranged to direct fuel into the air flow passages of the air/fuel mixer elements. Each of the fuel/air mixer elements has at least one outlet arranged to supply fuel to the element. | 05-14-2015 |
| 20150135723 | COMBUSTOR NOZZLE AND METHOD OF SUPPLYING FUEL TO A COMBUSTOR - A combustor nozzle includes a fuel passage that extends generally axially in the nozzle and a surface that extends radially across at least a portion of the fuel passage. A projection in the surface extends generally axially down-stream from the surface, and an indention in the surface radially surrounds the projection. An oxidant supply is in fluid communication with an oxidant passage, and the oxidant passage is radially displaced from the fuel passage and terminates at an oxidant outlet. A method of supplying a fuel to a combustor includes flowing fuel through a projection in a surface, wherein the projection extends generally axially downstream from the surface, and flowing fuel through an indention in the surface, wherein the indention radially surrounds the projection. The method further includes flowing an oxidant through an oxidant outlet that circumferentially surrounds the indention in the surface. | 05-21-2015 |
| 20150135724 | PARALLEL METERING PRESSURE REGULATION SYSTEM WITH INTEGRATED FLOW METER PLACEMENT - A fuel pressure regulation system is provided. The fuel pressure regulation system includes a supply arrangement for supplying an outlet flow. The supply arrangement includes a primary pressure regulator and a bypass pressure regulator. The primary pressure regulator is operably connected to downstream pressures of each of a plurality of parallel metering circuits. The bypass pressure regulator is upstream from an output of a pump or pumps, which in combination with the primary pressure regulator, is operable to increase the output of the pump or pumps to accommodate for downstream pressure spikes in any one of the parallel metering circuits. | 05-21-2015 |
| 20150292416 | Gas Turbine Power Generation Equipment - This invention provides a gas turbine power generation equipment adapted to prevent step out of an electric power generator due to a system trouble such as a momentary power failure in a local grid including intermittent renewable energy generation equipment and the gas turbine power generation equipment. The gas turbine power generation equipment that supplies electric power in cooperation with the intermittent renewable energy generation equipment in the local grid interconnected to a power grid includes: a fuel flow control valve that controls a flow rate of fuel supplied to a combustor; a bleed valve or inlet air flow control valve that controls a flow rate of air compressed by a compressor and supplied to the combustor; and a control unit configured so that if voltage on the power grid decreases below a threshold V(t), the control unit outputs a control signal to at least one of the fuel flow control valve, the bleed valve, and the inlet air flow control valve. The control unit thereby reduces instantaneously at least one of the fuel flow rate and the compressed air flow rate. After an elapse of a predetermined time, the control unit instantaneously returns the reduced flow rate to its original level. | 10-15-2015 |
| 20150300646 | METHOD FOR PREMIXING AIR WITH A GASEOUS FUEL AND BURNER ARRANGEMENT FOR CONDUCTING SAID METHOD - A method for premixing air with a gaseous fuel for being burned in a combustion chamber includes:
| 10-22-2015 |
| 20150308348 | CONTINUOUS DETONATION WAVE TURBINE ENGINE - A gas turbine engine includes a transient plasma igniter in communication with a continuous detonation wave combustor. A method of operating a gas turbine engine includes maintaining ignition of a continuous detonation wave combustor with a transient plasma igniter. | 10-29-2015 |
| 20150316266 | BURNER WITH ADJUSTABLE RADIAL FUEL PROFILE | 11-05-2015 |
| 20150316267 | RECESSED FUEL INJECTOR POSITIONING - A combustion chamber for a gas turbine is provided. The combustion chamber has a pilot burner device, a fuel injector and an ignitor unit. The pilot burner device has a pilot body with a pilot surface which is facing an inner volume of the combustion chamber. The fuel injector has a fuel outlet for injecting a fuel into the inner volume. The ignitor unit is adapted for igniting the fuel inside the inner volume, wherein the ignitor unit is arranged at the pilot surface such that fuel which passes the ignitor unit is ignitable. The pilot body includes a recess, wherein the fuel outlet is arranged within the recess. | 11-05-2015 |
| 20150323187 | METHOD AND APPARATUS FOR ASSISTING WITH THE COMBUSTION OF FUEL - An apparatus and method for assisting with the combustion of fuel are described. The apparatus includes a swirler assembly and a fuel nozzle. Fuel is directed into a fuel nozzle mixing chamber and combines with air therein to form a fuel-air mixture. At least one plasma generator, at least partially within the fuel nozzle, generates an at least one of an at least partially ionized air-fuel mixture and an at least partially dissociated air-fuel mixture (“at least partially I/D air-fuel mixture”) via a plasma generator discharge. A combustion chamber inlet admits the at least partially I/D air-fuel mixture from the plasma generator into a combustion chamber internal volume. Combustion air flows through the swirler body and into the combustion chamber internal volume. Combustion of the at least partially I/D air-fuel mixture with the combustion air occurs at least partially within the combustion chamber internal volume to responsively produce products. | 11-12-2015 |
| 20150330307 | LASER-CHARGED HIGH-SPEED PROPULSION SYSTEM AND METHOD FOR PRODUCTION OF HIGH-POWERED LASER - A high-speed propulsion system for a high-speed vehicle includes an air intake duct, a compression section, a combustion section, and an expansion section. A gas dynamic laser generator is disposed proximate the expansion section and generates a source of laser light. A laser transmission mechanism receives the laser light and transmits the laser light from the gas dynamic laser generator to the combustion section and to presents the laser light into the air/fuel mixture of the flowpath which promotes mixing, combustion, and flame stability. | 11-19-2015 |
| 20150330311 | FUEL METERING VALVE AND METHOD OF MANAGING FUEL IN A METERING VALVE - A fuel metering valve includes a main flow path extending axially between an inlet and an outlet. Also included is a plunger disposed around a portion of a plunger guide, the plunger and the plunger guide configured to translate between an open position and a closed position to selectively distribute a fuel flowing through the main flow path to the outlet of the fuel metering valve. Further included is a solenoid coil disposed between a solenoid outer body and a solenoid inner body, the solenoid coil configured to magnetically attract the plunger to the open position. Yet further included is a secondary flow path for routing a stagnant volume of fuel upon translation of the plunger from the closed position to the open position. | 11-19-2015 |
| 20150330636 | SYSTEM AND METHOD FOR CONTROL OF COMBUSTION DYNAMICS IN COMBUSTION SYSTEM - A system includes a gas turbine engine that includes a first combustor and a second combustor. The first combustor includes a first fuel nozzle disposed in a first head end chamber of the first combustor. The first fuel nozzle includes a first orifice configured to inject fuel into a first combustion chamber of the first combustor. The second combustor includes a second fuel nozzle disposed in a second head end chamber of the second combustor. The second fuel nozzle includes a second orifice configured to inject the fuel into a second combustion chamber of the second combustor. The second combustor also includes a second orifice plate disposed in a fuel path upstream of the second orifice. The second orifice plate is configured to help reduce modal coupling between the first combustor and the second combustor. | 11-19-2015 |
| 20150337739 | RAMP RATE CONTROL FOR A GAS TURBINE - A method for operating a gas turbine is disclosed. The method can include receiving a command for a new load set-point for a gas turbine. In response to the received new load set-point, a fuel valve of the gas turbine can be metered to a first position, the first position corresponding to a first gas turbine load less than the new load set-point. After a first predetermined time, the fuel valve can be metered to a second position, the second position corresponding to a second gas turbine load greater than the first gas turbine load and less than the new load set-point. The fuel valve is incrementally metered more open until the gas turbine is operating at the new load set-point. | 11-26-2015 |
| 20150337742 | GAS TURBINE WITH FUEL COMPOSITION CONTROL - A method for operating a gas turbine plant is provided. According to the method a first fuel gas with a first fuel reactivity and a second fuel gas with a second fuel reactivity which is higher than the first fuel reactivity are injected into a combustor of the gas turbine, and the ratio of the mass flows of the second fuel gas to the first fuel gas is controlled depending on the combustion behavior of the combustor. A gas turbine plant configured to carry out the method is further shown. | 11-26-2015 |
| 20150338102 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor comprises an annular combustor chamber formed between the inner and outer liners. Fuel nozzles each have an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. A plurality of nozzle air holes are defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles. The nozzle air holes are configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber. A central axis of the nozzle air holes has a tangential component relative to the central axis of the annular combustor chamber. | 11-26-2015 |
| 20150345402 | SYSTEMS AND METHODS FOR VARIATION OF INJECTORS FOR COHERENCE REDUCTION IN COMBUSTION SYSTEM - A system includes a gas turbine engine having a first combustor and a second combustor. The first combustor includes a first fuel conduit having a first plurality of injectors. The first plurality of injectors are disposed in a first configuration within the first combustor along a first fuel path, and the first plurality of injectors are configured to route a fuel to a first combustion chamber. The system further includes a second combustor having a second fuel conduit having a second plurality of injectors. The second plurality of injectors are disposed in a second configuration within the second combustor along a second fuel path, and the second plurality of injectors are configured to route the fuel to a second combustion chamber. The second configuration has at least one difference relative to the first configuration. | 12-03-2015 |
| 20150345791 | APPARATUS AND A METHOD OF CONTROLLING THE SUPPLY OF FUEL TO A COMBUSTION CHAMBER - Combustion chamber includes a primary stage fuel burner to supply fuel into a primary combustion zone and secondary stage fuel burner supplies fuel into a secondary combustion zone. A sensor is positioned downstream from combustion chamber measuring concentration of one or more compounds. The sensor sends measurements of the concentration of compounds to a processor. The processor compares the measured concentration of the compound with a first threshold value and a second threshold value. The processor supplies more fuel to the primary stage fuel burner if the measured concentration is higher than the first threshold value, supplies more fuel to the secondary stage fuel burner if the measured concentration is higher than the second threshold value, maintaining fuel supply to the primary and secondary stage fuel burners if the measured concentration is lower than the first threshold value and lower than the second threshold value. | 12-03-2015 |
| 20150352483 | METHOD AND DEVICE FOR GENERATING FUEL FOR A GAS TURBINE - The invention relates to a method for fractionating a feed gas ( | 12-10-2015 |
| 20150354458 | METHOD FOR STARTING A COMBUSTION SYSTEM - A method for starting a combustion system having a first ignition device and an at least second ignition device, a processing unit and a sensor system. To obtain a reliable ignition the method includes as a starting sequence at least the following: monitoring during a working condition of the combustion system, an operational state of the first and the at least second ignition device by the sensor system; identifying a predefined state of the first and/or the at least second ignition device by the processing unit; and as a further step: in case of an identification of the predefined state changing in at least one parameter of at least one of the ignition devices by the processing unit. A combustion system is equipped to be operable with the method as well as to a flow engine with such a combustion system. | 12-10-2015 |
| 20150354459 | FUEL INJECTOR FOR HIGH ALTITUDE STARTING AND OPERATION OF A GAS TURBINE ENGINE - A fuel injector for a combustor of a gas turbine engine includes an air swirler adjacent to a pressure atomizer. | 12-10-2015 |
| 20150354460 | Distributed Spark Igniter for a Combustor - An ignition system for a combustor of a gas turbine engine is disclosed. The ignition system may include an igniter operatively associated with the combustor, and an electrode operatively associated with the combustor and spaced from the igniter, wherein an electrical potential is created between the igniter and the electrode to produce an electric arc therebetween. | 12-10-2015 |
| 20150354466 | GAS TURBINE SYSTEM, CONTROLLER, AND GAS TURBINE OPERATION METHOD - The gas turbine system has: a gas turbine having a compressor, a combustor, and a turbine; a fuel supply mechanism for supplying fuel to the combustor; a composition detection unit for detecting the composition of the fuel; and a controller for controlling the flow rate of the fuel supplied from the fuel supply mechanism to the combustor, on the basis of a function of the exhaust temperature of exhaust gas passing through the turbine and either air pressure of air expelled from the compressor to the combustor or an expansion ratio of the turbine. The controller calculates the specific heat ratio of the combustion gas from the composition of the fuel detected by the composition detection unit, corrects the function on the basis of the calculated specific heat ratio, and controls the flow rate of the fuel on the basis of the corrected function. | 12-10-2015 |
| 20150377059 | METHOD FOR MONITORING A DEGREE OF CLOGGING OF THE STARTING INJECTORS OF A TURBINE ENGINE - A method for monitoring a degree of clogging of the starting injectors of a turbine engine, which includes: a combustion chamber into which at least one starting injector supplied with fuel leads, the starting injectors being suitable for initiating the combustion in the chamber by igniting the fuel; and a turbine rotated by the gases resulting from the combustion of the fuel in the chamber, the method including the steps involving: the measurement, during a phase of starting the turbine engine, of the temperature of the exhaust gases at the outlet of the turbine; and the determination, from the changes over time in the temperature thus measured, of a degree of clogging of the starting injectors. A system for monitoring a degree of clogging capable of implementing the method, and a turbine engine including such a system. | 12-31-2015 |
| 20150377138 | SYSTEMS AND METHODS FOR A FUEL PRESSURE OSCILLATION DEVICE FOR REDUCTION OF COHERENCE - A system with a gas turbine engine is provided. The gas turbine engine includes a first combustor comprising a first fuel nozzle, a second combustor comprising a second fuel nozzle, and a first fuel pressure oscillation system. The first fuel pressure oscillation system includes a first rotary device coupled to a first fuel circuit. The first fuel circuit is disposed along a first fuel passage leading to the first fuel nozzle. The first rotary device is configured to generate a first fuel pressure oscillation through the first fuel nozzle. The gas turbine engine also includes a second fuel pressure oscillation system having a second rotary device coupled to a second fuel circuit. The second fuel circuit is disposed along a second fuel passage leading to the second fuel nozzle, and the second rotary device is configured to generate a second fuel pressure oscillation through the second fuel nozzle. | 12-31-2015 |
| 20150377490 | SUPPLEMENTARY LASER FIRING FOR COMBUSTION STABILITY - A combustion system for a gas turbine includes a combustion chamber having an end section and a pre-combustion section extending from the end section, a swirler device, an optional pilot burner device and a light emitting arrangement. Main fuel is injectable by the swirler device into an inner volume of the pre-combustion section. The main flame using main fuel is producible inside the inner volume. The pilot burner device is mounted to the end section of the combustion chamber such that a pilot fuel is injectable by the pilot burner device into the inner volume of the pre-combustion section, wherein a pilot flame using the pilot fuel is producible inside the inner volume for stabilizing the main flame. The light emitting arrangement emits an electromagnetic radiation into the inner volume, such that an energy input is generatable by the electromagnetic radiation for stabilizing the pilot flame and/or the main flame. | 12-31-2015 |
| 20160010493 | SYSTEM AND METHOD OF CONTROL FOR A GAS TURBINE ENGINE | 01-14-2016 |
| 20160010548 | SYSTEM AND METHOD FOR A TURBINE COMBUSTOR | 01-14-2016 |
| 20160010867 | SEQUENTIAL COMBUSTOR ARRANGEMENT WITH A MIXER | 01-14-2016 |
| 20160018110 | AXIALLY STAGED GAS TURBINE COMBUSTOR WITH INTERSTAGE PREMIXER - The present invention discloses a novel and improved apparatus and method for reducing the emissions of a gas turbine combustion system. More specifically, a combustion system is provided having a first combustion chamber and a premixer positioned proximate an outlet end of a combustion liner for mixing a second fuel/air mixture with hot combustion gases and burning the subsequent mixture to achieve reduced emissions levels. The premixer is positioned generally about the combustion liner and includes a plurality of channels and fuel injectors for introducing a fuel/air mixture, induced with a swirl, into a second, axially staged combustor. | 01-21-2016 |
| 20160033135 | Fuel Injector For High Flame Speed Fuel Combustion - A multiple fuel capable, pre-mixed, low emission injector is provided which is particularly suited for burning high flame speed fuels. The fuel injector includes an injector body having a preliminary pre-mixing chamber, an intermediate pre-mixing chamber and a final pre-mixing chamber. A fuel distributor separates the preliminary pre-mixing chamber from the intermediate pre-mixing chamber. A guide vane separates the intermediate pre-mixing chamber from the final pre-mixing chamber. | 02-04-2016 |
| 20160040885 | SEQUENTIAL COMBUSTION WITH DILUTION GAS - An exemplary sequential combustor arrangement includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The mixer includes at least three groups of injection tubes pointing inwards from the side walls of the mixer for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber. The first injection tubes of the first group have a first protrusion depth, the second injection tubes of the second group have a protrusion depth, and the third injection tubes of the third group have a third protrusion depth. | 02-11-2016 |
| 20160047315 | ATOMIZING FUEL NOZZLE - A fuel nozzle for a gas turbine engine. The nozzle has a body and a center axis. The body has an inner circumferential surface circumscribing a central passageway which is coaxial with the center axis. The nozzle also has air passages which extend predominantly radially inward through the body. The air passage outlets of each air passage are circumferentially spaced apart from one another along the inner circumferential surface. Each air passage conveys air through the body toward the nozzle center axis and into the central passageway. The nozzle also has fuel passages which extend through the body. Each fuel passage is disposed within the body between adjacent circumferentially spaced apart air passages and is transverse to the direction of extension of its neighboring air passages. | 02-18-2016 |
| 20160047318 | Torch Igniter - A gas turbine combustor assembly includes a primary combustion chamber in fluid communication with a primary fuel injector and a primary air inlet. A torch igniter is carried by the primary combustion chamber, and includes an auxiliary combustion chamber housing comprising a mixing chamber and a throat region converging downstream of the mixing chamber. An air swirler including a plurality of swirl openings surrounding an outlet of an auxiliary fuel injector is coupled to the auxiliary combustion chamber proximate the mixing chamber. An ignition source projects into the mixing chamber of the auxiliary combustion chamber. | 02-18-2016 |
| 20160047550 | DISTRIBUTED FUEL CONTROL SYSTEM - Distributed fuel control system and methods are disclosed. A distributed fuel control system may comprise a controller, a fuel delivery system, and fuel delivery system sensors and combustion sensors. The controller may output a control signal in response to at least one of the fuel delivery system sensor or the combustion sensors. In response, the fuel flow to individual multiplex fuel delivery unit may be controlled according to various methods. One such method includes determining a desired fuel pressure differential, directing a torque motor to set a pressure regulator to a position corresponding to the desired fuel pressure differential, determining a sensed fuel pressure differential, and adjusting the torque motor in response to a difference between the sensed fuel pressure differential and the desired fuel pressure differential. | 02-18-2016 |
| 20160061110 | LOW NOX TURBINE EXHAUST FUEL BURNER ASSEMBLY | 03-03-2016 |
| 20160061114 | METHOD FOR CONTROLLING A GAS TURBINE - The invention relates to a method for controlling a gas turbine, operating with an integral fuel reactivity measurement concept. In order to fast determine a safe operation range of the gas turbine with respect to flashback and blow-out, the method includes deducing the fuel composition and therefore the fuel reactivity by combined measurements of (n−1) physico-chemical properties of a fuel mixture with n>1 fuel components, for deriving the concentration of one component for each physico-chemical property of the fuel gas mixture or for determining of a ratio of the fuels with known compositions and adjusting at least one operation parameter of the gas turbine at least partially based on the determined property of the fuel gas mixture entering the combustors. With the technical solution of the present invention, by way of detecting fast changes in fuel gas, it is assured that the gas turbine may operate with varieties of fuel gas under optimized performance and in safe operation ranges. In actual applications, the present invention may improve flexibility of gas turbines and cost effectiveness of operation of the gas turbines. | 03-03-2016 |
| 20160069261 | Ultra-High Efficiency Gas Turbine (UHEGT) with Stator Internal Combustion - Provided is an internal combustion system for an ultra-high efficiency gas turbine (UHEGT) engine which includes a fuel injection system, an ignition system, a stator system, and a rotor system. The stator system includes a plurality of stators positioned radially around a central axis. The fuel injection system injects fuel within the stator system and the ignition system is located within the stator system allowing combustion to take place therein. The rotor system includes a plurality of rotors positioned radially around the central axis downstream from the stator system. The UHEGT-technology completely eliminates the combustion chambers and replaces the latter with a distributed combustion system using stator-internal combustion technology. This technology allows for an increase in the thermal efficiency of gas turbines of at least about 7% (and in many cases much more) above the thermal efficiency of the most advanced existing gas turbines. | 03-10-2016 |
| 20160069271 | Bulk Flame Temperature Regulator for Dry Low Emission Engines - Systems and methods for regulating a bulk flame temperature in a dry low emission engine are provided. According to one embodiment of the disclosure, a method may include measuring an exhaust gas temperature (EGT) and determining a target EGT. The target EGT is determined based at least in part on a compressor bleed air flow percentage and a combustor burning mode. The method may include calculating a bias based at least in part on the EGT and the target EGT and applying the bias to a bulk flame temperature schedule. The method may include regulating one or more staging valves and compressor bleeds of the DLE engine based at least in part on the bulk flame temperature schedule. The bulk flame temperature schedule is mapped to parameters of the staging valves and compressor bleeds to reduce nitric oxide, nitrogen dioxide, and carbon monoxide emissions. | 03-10-2016 |
| 20160069276 | A METHOD AND A DEVICE FOR GENERATING A COMMAND FOR THE FLOW RATE OF FUEL THAT IS TO BE INJECTED INTO A COMBUSTION CHAMBER OF A TURBINE ENGINE - During a stage (E | 03-10-2016 |
| 20160084169 | METHOD OF OPERATING A MULTI-STAGE FLAMESHEET COMBUSTOR - The present invention discloses a novel way of controlling a gas turbine engine using detected temperatures and detected turbine rotor speed. An operating system provides a series of operating modes for a gas turbine combustor through which fuel is staged to gradually increase engine power, yet harmful emissions, such as carbon monoxide, are kept within acceptable levels. | 03-24-2016 |
| 20160084503 | FUEL NOZZLE - A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, a primary air passageway centrally defined axially in the body, and a plurality of concentrically-arranged nozzle tip projections disposed at a downstream portion of the body. Each of the plurality of nozzle tip projections has a radially inwardly facing fuel filming surface communicating with respective fuel passages. The fuel filming surfaces are disposed radially outwardly of an outlet of the primary air passageway. A method for delivering fuel from a fuel nozzle of a combustor of a gas turbine engine is also presented. | 03-24-2016 |
| 20160097536 | FUEL NOZZLE - A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, an air passageway defined axially in the body, and a fuel passageway defined axially in the body radially outwardly from the air passageway. The fuel passageway has an outer wall including an exit lip at a downstream portion of the outer wall. The exit lip has a surface treatment including a swirl-inducing relief. A gas turbine engine and a method of inducing swirl in at least one of pressurised fuel and air exiting a fuel nozzle of a gas turbine engine are also presented. | 04-07-2016 |
| 20160097537 | FUEL NOZZLE - A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, an air passageway defined axially in the body, and a fuel passageway defined axially in the body radially outwardly from the air passageway. The fuel passageway has an outer wall including an exit lip at a downstream portion of the outer wall. The lip generally increases in diameter as it extends downstream. | 04-07-2016 |
| 20160097538 | FUEL NOZZLE - A fuel nozzle for a combustor of a gas turbine engine includes a body defining an axial direction and a radial direction, an air passageway defined axially in the body, and a fuel passageway defined axially in the body radially outwardly from the air passageway. The air passageway has a swirl-inducing relief defined at an exit lip of an outer wall of the air passageway. A gas turbine engine and a method of inducing swirl in an air passageway of a fuel nozzle of a gas turbine engine are also presented. | 04-07-2016 |
| 20160115839 | SYSTEM AND METHOD FOR EMISSIONS CONTROL IN GAS TURBINE SYSTEMS - A system includes an emissions control system. The emissions control system includes a processor programmed to receive one or more selective catalytic reduction (SCR) operating conditions for an SCR system. The SCR system is included in an aftertreatment system for an exhaust stream. The processor is also programmed to receive one or more gas turbine operating conditions for a gas turbine engine. The gas turbine engine is configured to direct the exhaust stream into the aftertreatment system. The processor is further programmed to derive a NH | 04-28-2016 |
| 20160123237 | VARIABLE PRESSURE AIR SUPPLY - The present disclosure relates to engine buffer systems. An engine buffer system may include a low pressure supply line and a high pressure supply line. A continuously variable valve may be coupled to and/or in fluid communication with the low pressure supply line and the high pressure supply line. The continuously variable valve may be adjusted to supply any pressure between a pressure of the low pressure supply line and a pressure of the high pressure supply line to a buffer line. | 05-05-2016 |
| 20160138469 | TOROIDAL COMBUSTION CHAMBER - A device comprising a combustion toroid for receiving combustion-induced centrifugal forces therein to continuously combust fluids located therein and an outlet for exhaust from said combustion toroid. | 05-19-2016 |
| 20160138479 | APPLICATION OF FUEL FLOW-BASED PROBABILISTIC CONTROL IN GAS TURBINE TUNING, RELATED CONTROL SYSTEMS, COMPUTER PROGRAM PRODUCTS AND METHODS - Various embodiments include a system having: at least one computing device configured to tune a set of gas turbines (GTs) by performing actions including: commanding each GT in the set of GTs to a base load level, based upon a measured ambient condition for each GT; commanding each GT in the set of GTs to adjust a respective output to match a nominal mega-watt power output value, and subsequently measuring an actual fuel flow value for each GT; and adjusting an operating condition of each GT in the set of GTs based upon a difference between the respective measured actual fuel flow value and a nominal fuel flow value at the ambient condition. | 05-19-2016 |
| 20160138807 | Gas Turbine Engine Flow Regulating - A gas turbine combustor assembly includes a fuel injector, a dome stator around the fuel injector, and a dome sleeve coupled to the dome stator. The dome sleeve defines an air inlet opening with the dome stator, and is carried to move with respect to the dome stator to change a flow area of the air inlet opening. The dome sleeve also defines a nozzle sloping downstream from the air inlet opening toward an outlet of the combustor assembly. The sloping nozzle defines an annular pinch gap adjacent an outlet of the fuel injector, and is coupled to move with the dome sleeve to change a flow area through the pinch gap. | 05-19-2016 |
| 20160146464 | COMBUSTOR WITH ANNULAR BLUFF BODY - The present invention relates to a gas turbine combustor comprising: a flow sleeve; a combustion liner located at least partially within the flow sleeve thereby creating a main passage between the flow sleeve and the combustor liner; a dome located forward of the flow sleeve and encompassing at least a part of the combustion liner, the dome having a substantially rounded head end thereby forming a turning passage between the liner and the head end; and a swirler wall aligned along a centerline of the combustor, the swirler wall projecting into the liner, wherein the swirler wall and the rounded head end are connected, wherein the connection forms an annular end face. | 05-26-2016 |
| 20160146467 | COMBUSTOR LINER - The present invention relates to a combustor liner for a gas turbine, the combustor liner having substantially cylindrical shape and comprising a first section and a second section wherein the first section is upstream of the second section with respect to the hot gas flow during operation, wherein the first section is ring shaped and comprises a rounded lip section and a trailing section, wherein an inner radius of the trailing section is increasing along a centerline of the liner in the direction of the hot gas flow during operation. The invention also relates to a combustor and a gas turbine comprising such combustion liner. | 05-26-2016 |
| 20160153364 | GAS TURBINE FUEL CONTROL SYSTEM | 06-02-2016 |
| 20160169110 | PREMIXING NOZZLE WITH INTEGRAL LIQUID EVAPORATOR | 06-16-2016 |
| 20160169111 | PIVOTING STOWABLE SPRAYBAR | 06-16-2016 |
| 20160169115 | TURBINE ENGINE CONTROL SYSTEM | 06-16-2016 |
| 20160169120 | Fuel Schedule for Robust Gas Turbine Engine Transition Between Steady States | 06-16-2016 |
| 20160177835 | GAS TURBINE ENGINE WITH ANGULARLY OFFSET TURBINE VANES | 06-23-2016 |
| 20160178207 | AXIALLY STAGED MIXER WITH DILUTION AIR INJECTION | 06-23-2016 |
| 20160195020 | MULTIPOINT FUEL INJECTION SYSTEM FOR A TURBOMACHINE AND ASSOCIATED REGULATION METHOD | 07-07-2016 |
| 20160195025 | GAS TURBINE FLAMEOUT DETECTION | 07-07-2016 |
| 20160201563 | FUEL MANAGEMENT SYSTEM FOR A TURBINE ENGINE | 07-14-2016 |
| 20160252019 | METHOD AND SYSTEM TO INCREASE GAS TURBINE ENGINE DURABILITY | 09-01-2016 |
| 20170234540 | THERMAL AND THRUST MANAGEMENT IN DYNAMIC PRESSURE EXCHANGERS | 08-17-2017 |
| 20170234541 | BURNER FOR A CAN COMBUSTOR | 08-17-2017 |
