Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A system includes a multi-tube fuel nozzle having an inlet plate and a plurality of tubes adjacent the inlet plate. The inlet plate includes a plurality of apertures, and each aperture includes an inlet feature. Each tube of the plurality of tubes is coupled to an aperture of the plurality of apertures. The multi-tube fuel nozzle includes a differential configuration of inlet features among the plurality of tubes.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A system includes a fuel nozzle. The fuel nozzle includes a center body configured to receive a first portion of air and to deliver the air to a combustion region. The fuel nozzle also includes a swirler configured to receive a second portion of air and to deliver the air to the combustion region. The swirler includes an outer shroud wall, an inner hub wall, and a swirl vane. The swirl vane includes a radial swirl profile at a downstream edge of the swirl vane. The radial swirl profile includes a region extending from the outer shroud wall to a first transition point and a second region extending from the transition point to the inner hub wall. At least one of the first and second regions is substantially straight and at least one of the first and second regions is arcuate.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: The invention relates to a method for burning a fluid fuel, in which fuel is reacted in a catalytic reaction, whereupon catalytically pre-reacted fuel continues to be burned in a secondary reaction. A swirling component is impressed onto the pre-reacted fuel, allowing the secondary reaction to be ignited in a spatially controlled manner, resulting in complete burnout. The invention further relates to a burner for burning a fluid fuel, in which the fuel outlet of a catalytic burner is disposed upstream of the fuel outlet of a primary burner in the direction of flow of the fuel within a flow channel such that the fuel is catalytically reacted. The catalytic burner is provided with a number of catalytically effective elements which are arranged such that a vortex is created in the flow channel. The invention can be applied particularly to combustion chambers of gas turbines.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A burner (1) for a gas turbine includes a duct (2) enclosing a plurality of vortex generators (3) and, downstream of them, a lance (5) provided with nozzles (6) for injecting a gaseous fuel. The burner (1) also includes a mixer (7) for diluting and mixing the gaseous fuel with air to form a mixture. The mixer (7) is connected to the nozzles (6) for feeding them with the mixture. A method includes feeding the gaseous fuel in the burner (1).

Abstract: 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.

Abstract: 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.

Abstract: A combustor includes an end cap having upstream and downstream surfaces and a cap shield surrounding the upstream and downstream surfaces. First and second sets of premixer tubes extend from the upstream surface through the downstream surface. A first fuel conduit supplies fuel to the first set of premixer tubes. A casing circumferentially surrounds the cap shield to define an annular passage, and a second fuel conduit supplies fuel through the annular passage to the second set of premixer tubes. A method for supplying fuel to a combustor includes flowing a working fluid through first and second sets of premixer tubes, flowing a first fuel into the first set of premixer tubes, and flowing a second fuel through an annular passage surrounding the end cap and into the second set of premixer tubes.

Abstract: 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 (psc1, psc3) for a predetermined time (dt) in case the combustion device is in the predefined operating stage.

Abstract: 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.

Abstract: A combustion apparatus (10) incorporating a dry low NOx (DLN) combustor (20) that uses two different fuels (30, 40) simultaneously. One fuel (30) is premixed with air (22) for a primary combustion zone (25), and a second fuel (40) supplies a diffusion pilot flame (29). Each fuel may be selected for its specialized role, maximizing overall combustion efficiency while lowering emissions. A primary fuel (30) such as natural gas (NG) may be chosen for its economy and combustion characteristics in a lean premix. A pilot fuel (40) may be chosen for having a low diffusion flame temperature and a clean burn. An oxygenated pilot fuel such as dimethyl ether (DME) has a lower flame temperature than natural gas, thereby reducing NOx from the pilot flame. The pilot fuel (40) may be produced from the primary fuel (30) by a reformer (50) associated with the combustion apparatus.

Abstract: 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.

Abstract: An apparatus and process is provided for combining fuel and combustion air to produce a mixture. The mixture is burned in a combustion chamber to produce a flame.

Abstract: 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.

Abstract: 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.

Abstract: Embodiments for controlling a gas turbine engine to minimize combustion dynamics and emissions are disclosed. Methods and an apparatus are provided for controlling the gas turbine engine where a compressor inlet temperature is measured and a turbine reference temperature is calculated in real-time and utilized to determine the most-efficient fuel splits and operating conditions for each of the fuel circuits. The fuel flow for the fuel circuits are then adjusted according to the identified fuel split.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A fuel delivery system for a gas turbine designed to efficiently transfer from one type of fuel to a separate fuel, comprising different and cooperating fuel modules, namely high hydrogen fuel gas, distillate fuel and naphtha fuel modules, each of which feeds a different liquid or gas fuel to the combustors, an atomized air delivery system for either the naphtha or distillate (or combinations thereof), an air extraction system, a plurality of distribution control valves for the high hydrogen fuel gas, distillate fuel and naphtha fuel modules, and a nitrogen purge system to purge the high hydrogen fuel gas lines to the combustors and/or flushing the naphtha lines with distillate.

Abstract: 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.

Abstract: A method and apparatus to determine a first heating value of a low BTU fuel, determine a target fuel quality level based on a state of a turbine system, control a second heating value of a high BTU fuel, and inject the high BTU fuel into the low BTU fuel to achieve the target fuel quality level.

Abstract: 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.

Abstract: A gas turbine engine that includes a compressor, a combustion stage, a turbine assembly, a fuel gas feed system for supplying fuel gas to a combustion stage of the gas turbine, an integrated fuel gas characterization system, and a buffer tank is disclosed. The integrated fuel gas characterization system may determine the heating value of the fuel and adjust either the airflow to the compressor or the fuel gas flow to the combustor to maintain a design fuel-to-air ratio.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: Clean combustion and equilibration equipment and methods are provided to progressively deliver, combust and equilibrate mixtures of fuel, oxidant and aqueous diluent in a plurality of combustion regions and in one or more equilibration regions to further progress oxidation of products of incomplete combustion, in a manner that sustains combustion while controlling temperatures and residence times sufficiently to reduce CO and NOx emissions to below 25 ppmvd, and preferably to below 3 ppmvd at 15% O2.

Abstract: A system and method for flame stabilization is provided that forestalls incipient lean blow out by improving flame stabilization. A combustor profile is selected that maintains desired levels of power output while minimizing or eliminating overboard air bleed and minimizing emissions. The selected combustor profile maintains average shaft power in a range of from approximately 50% up to full power while eliminating overboard air bleed in maintaining such power settings. Embodiments allow for a combustor to operate with acceptable emissions at lower flame temperature. Because the combustor can operate at lower bulk flame temperatures during part power operation, the usage of inefficient overboard bleed can be reduced or even eliminated.

Abstract: A method is described for the start-up of a gas turbine comprising the phases of effecting a preliminary purging cycle of the discharge duct (28) of the turbine (20), establishing a predetermined minimum value (FSR1) for the flow of gaseous fuel entering the combustor (14) for a period of time which is adequate for effecting a first attempt at ignition and effecting a first attempt at ignition, effecting an intermediate purging cycle of the discharge duct (28), interrupting the flow of gaseous fuel to the combustor (14), and progressively increasing the value (FSRn) of the flow of gaseous fuel entering the combustor (14), effecting further attempts at ignition until the mixture of air/gaseous fuel has been ignited and the consequent start-up of the turbine (20), or until a predetermined maximum value (FSRmax) of the flow of gaseous fuel has been reached.

Abstract: 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.

Abstract: 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%.

Abstract: The present application provides a dual gas fuel delivery system and a method of delivering two gas fuels. The dual gas fuel delivery system may include (a) a low energy gas delivery system comprising a low energy gas inlet, a gas split, a low energy gas primary manifold outlet, and a low energy gas secondary manifold outlet; (b) a high energy gas delivery system comprising a high energy gas inlet and a high energy gas primary manifold outlet; (c) a primary manifold; and (d) a secondary manifold, wherein the low energy gas primary manifold outlet and the high energy gas primary manifold outlet are coupled to the primary manifold, and wherein the low energy gas secondary manifold outlet is coupled to the secondary manifold.