Abstract: A method of and a plant for combusting carbonaceous fuel, the method including the steps of introducing particulate oxygen selective sorbent, such as a perovskite type material, into an adsorption reactor of the combustion plant to form a first particle bed in the adsorption reactor, fluidizing the first particle bed by an oxygen-containing first fluidizing gas to adsorb oxygen from the fluidizing gas to the sorbent, conveying oxygen-rich sorbent from the adsorption reactor to a combustion reactor of the combustion plant to form a second particle bed in the combustion reactor, fluidizing the second particle bed by an oxygen-deficient second fluidizing gas to desorb oxygen from the sorbent, so as to produce free oxygen gas, and introducing carbonaceous fuel into the combustion reactor to oxidize the fuel with the free oxygen gas.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A method of operating a gas turbine engine having a rotatable turbine shaft with an electric machine mounted to the shaft. The method includes providing supplemental acceleration and/or deceleration of the turbine shaft of the gas turbine engine through the use of the electric machine operated as an electric motor and/or an electric generator, in order to avoid an undesirable engine speed range during engine operation.

Abstract: 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.

Abstract: The present application provides a method of operating a dual gas fuel delivery system comprising selecting a manifold and fuel flow split; setting the stroke of a high-energy gas control valve based on the fuel split; measuring the primary manifold nozzle pressure ratio across a primary manifold nozzle outlet; comparing the primary manifold nozzle pressure ratio against a primary manifold specification limit; adjusting the stroke of a primary low energy gas control valve to maintain the pressure ratio across the primary manifold nozzle outlet within the primary manifold specification limit; measuring the secondary manifold nozzle pressure ratio across a secondary manifold nozzle outlet; comparing the secondary manifold nozzle pressure ratio against a secondary manifold specification limit; and adjusting the stroke of a secondary low energy gas control valve to maintain the pressure ratio across the secondary manifold nozzle outlet within the secondary manifold specification limit.

Abstract: The present invention discloses a turbine assembly (20b) driven by predetermined deflagration of anaerobic fuel. The use of anaerobic fuel enables operation without any necessity for an additional oxidant, and leads to more efficient and environmentally friendly turbine operation. In addition, the gaseous products of the deflagration can be used for any number of purposes after they have passed through the turbine, e.g. combustion of the inflammable portion can drive a second turbine stage (214, 216) or be used to heat air or water.

Abstract: An integral impeller and fuel pump for a small gas turbine engine, the impeller being a centrifugal impeller to supply compressed air to the combustor, the rotor disk of the centrifugal compressor having at least one radial fuel passage connecting a central bore of the rotor disk to a fuel nozzle located on the rear face of the rotor disk. A low pressure fuel supplied by a small pump delivers the fuel to the central bore where the fuel is collected within helical grooves and channeled into the radial fuel passages where the fuel is pressurized to a relatively high pressure due to the high rotational speed of the centrifugal compressor. Highly compressed fuel is injected into the combustor through fuel nozzles in an axial direction as the rotor disk rotates. Helical grooves in the central bore each connected to a radial fuel passage move the fuel to an opposite side of the central bore to counteract rotor imbalance.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A method and apparatus for measuring a start fuel flow to a pilot nozzle of a fuel system of a gas turbine engine using pressure differential between fuel passages leading to fuel nozzles.

Abstract: A power supply device or system for aeronautics, having a hydrocarbon supply for supplying an engine with hydrocarbon fuel and a hydrogen supply having a fuel reformer for producing hydrogen from hydrocarbon fuel from said hydrocarbon supply. The hydrogen supply is connected to a hydrogen-powered fuel cell for producing electric power and to a hydrogen injecting system for injection of hydrogen into a combustion chamber of the engine. Further, the invention relates to an aircraft having an engine that can be supplied by that power supplying device or system, and to a method for operating said engine.

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 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.

Abstract: A pulse detonation device for dividing a pulse detonation shock wave into an primary and control portion to reduce the strength of a propagating shock wave and/or change its direction. The device contains a flow separator which directs a portion of the shock wave into itself, thus reducing the shock wave's strength. In one configuration, the control region converges in cross-sectional area so as to accelerate the flow in the control region, while the primary region diverges to slow the flow in the primary region. The flow in the control region is directed, at an angle, into the flow of the primary region to impede and/or redirect the flow of the primary region.

Abstract: A protocol for assembling a fuel shut off pin valve assembly for a turbo fan engine. The protocol outlines a specific sequence of events in order to ensure failsafe incorporation of a fuel shut off valve assembly within the low pressure turbine area and specifically within the engine casing of the turbo fan.

Abstract: In a gas turbine engine control system having a first valve for regulating the flow rate of fuel in premixed combustion and a second valve for regulating that in diffusion combustion, when the combustion mode is switched from one mode to the other mode, whichever of the first and second valves is associated with the other combustion mew is opened to supply the amount of fuel required for conducting the other combustion mode and, the opening of the valve associated with the one combustion mode is gradually decreased, while that of the other combustion mode is gradually increased, thereby maintaining the total amount of fuel supplied to the engine constant, after elapse of a predetermined period. At the time of switching the combustion mode, therefore, the total amount of fuel supplied to the engine can be accurately controlled to a desired value to minimize engine speed fluctuation.

Abstract: A method of monitoring the health of a gas turbine engine 10 comprising a fuel system 30 having a fuel metering valve 36 and a fuel pump 40 for supplying fuel to a combustor apparatus 15. The method comprising the steps of: monitoring the fuel metering valve percentage open, detecting and sending a warning message when the fuel metering valve percentage open is at 97% or greater and within a predetermined time from when the first percentage open is at 97% or greater, replacing the fuel pump 40. In this way scheduling of servicing and replacement of the fuel pump may be made without disruption to the aircraft operator particularly where the operator has a fleet of aircraft.

Abstract: Aspects of the invention relate to a system and method for introducing fuel to the flow of air along a surface upstream of a combustion area. Fuel injectors can be used to deliver fuel to the air flow at or near the boundary layer of the air flow. The fuel injectors are ideally located substantially outside of the flow path. Many benefits can be realized by keeping the injectors out of the flow path including avoidance of potential excitations and smaller pressure drops in the flow. The potential of flame from the combustion area traveling upstream to the fuel injector can be minimized by creating a disturbance in the boundary layer of the flow downstream of the fuel injectors. Such disturbances can be achieved by placing a turbulence generator in the boundary layer. The turbulence generator can be a surface protrusion or a fluid cross-flow.

Abstract: In a combustor of a gas turbine which has a pilot nozzle being installed to the center of the axis of a combustor basket and a plurality of main nozzles being installed to the vicinity of the pilot nozzle and provided with a premixing tool on the outer circumference thereof, wherein, fuel being injected as air-fuel pre-mixture from the main nozzle into the interior of a transition piece forming a combustion chamber downstream of the combustor basket is ignited by diffusion flame being generated by the pilot nozzle in the transition piece so as to generate a premixed flame in the transition piece, wherein combustion is performed by a part of the plurality of main nozzles from start-up until a predetermined load rate and then performed by adding the remaining portion of the plurality of main nozzles when the predetermined load rate is exceeded.

Abstract: An aircraft engine is provided with at least one pulse detonation device, and the operational frequency of the pulse detonation device is varied over an operational range of frequencies around a mean frequency value. The pulse detonation device can be positioned upstream, downstream or adjacent to a turbine section of the engine. An additional embodiment of the present invention is an aircraft engine provided with more than one pulse detonation device, and the operational frequency of one, or more, of the pulse detonation devices is varied over an operational range of frequencies around a mean frequency value.

Abstract: A method of designing a fuel nozzle of a gas turbine engine to reduce fretting thereof during use, including establishing an initial nozzle design, determining a first natural frequency of that design and a running frequency range of the gas turbine engine, and increasing a first transverse dimension of the stem member of the nozzle across a length of a portion thereof adjacent the inlet end until the first natural frequency of the nozzle is outside the running range, while a second transverse dimension of the portion remains at least substantially unchanged across the length thereof.

Abstract: 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.

Abstract: A gas stream with a reduced oxygen concentration relative to ambient air is used to vaporize a liquid fuel or liquefied higher hydrocarbon gas, or is mixed with a vaporized gas, and the reduced oxygen vaporized fuel gas is fed to a combustion device such as a premixed or diffusion combustor. Preferably, the oxygen content of the gas stream is less than the limiting oxygen index. By mixing the fuel with a gas stream that has an appropriately reduced oxygen content, auto-ignition prior to the desired flame location in the combustor can be avoided. In some embodiments, the reduced oxygen stream is generated from an air separator or taken from the exhaust of the combustion device.

Abstract: A method for controlling the rate of transfer between operation with gas and liquid fuel types to minimize the time in undesirable operational mode, thereby preventing excessive wear and damage to gas turbine hardware. The method includes completing a fuel prefill for the oncoming fuel type through the fuel system and determining if a total fuel demand for the oncoming fuel type is greater than a predetermined flow rate for the oncoming fuel type. The method also includes selecting a fuel transfer rate and transferring from the offgoing fuel type to the oncoming fuel type at the selected fuel transfer rate. Further, the method includes determining if the offgoing fuel flowrate has decreased below a predetermined flow rate for the offgoing fuel type, selecting a final fuel transfer rate and completing the transfer from offgoing fuel type to oncoming fuel type at a selected final fuel transfer rate.

Abstract: 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.

Abstract: A combustor swirler for a gas turbine engine and method of manufacturing by metal injection molding an inner component and an outer cylindrical component. Indentations are molded in one of the inner and outer components and sealed by the engagement of the components together to form a series of fluid flow passages. The inner and outer components are molded with interlocking features for ensuring proper alignment of the components during assembly.

Abstract: A fuel manifold for a gas turbine engine having a first peripheral surface having a first channel defined therein and a second peripheral surface having a second channel defined therein, each of the first and second channels being sealingly enclosed to define a corresponding conduit.

Abstract: 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.

Abstract: A secondary fuel delivery system for delivering a secondary stream of fuel and/or diluent to a secondary combustion zone located in the transition piece of a combustion engine, downstream of the engine primary combustion region is disclosed. The system includes a manifold formed integral to, and surrounding a portion of, the transition piece, a manifold inlet port, and a collection of injection nozzles. A flowsleeve augments fuel/diluent flow velocity and improves the system cooling effectiveness. Passive cooling elements, including effusion cooling holes located within the transition boundary and thermal-stress-dissipating gaps that resist thermal stress accumulation, provide supplemental heat dissipation in key areas.

Abstract: A fuel modulation system usable in can-annular combustor systems of turbine engines for more efficiently managing fuel flow to reduce the amount of NOx while maintaining appropriate combustor flame temperatures. The fuel modulation system controls individual inline modulation valves in cooperation with a overall stage control valve to bring a turbine engine through the startup phase and to maintain operating conditions at a load set point while reducing NOx variability between baskets by using individual combustor dynamic pressure measurements. The fuel modulation system may be managed such that individual inline modulation valves remain within a predetermined range of positions relative to a nominal position to reduce NOx variability.

Abstract: A method for operating a gas turbine engine includes channeling fluid from a cooling fluid source to a combustor that includes at least one deflector and flare cone. The deflector and flare cone are coupled together and are configured to define a cooling fluid channel therebetween. The flare cone has a plurality of cooling injectors extending therethrough. The plurality of injectors are spaced circumferentially about a centerline axis of the flare cone and are coupled in flow communication with the fluid source. The plurality of injectors has a plurality of first injectors and a plurality of second injectors. The method also includes directing a portion of the fluid through the plurality of first injectors. The method further includes directing a portion of the fluid through the plurality of second injectors, wherein the first plurality of injectors facilitates cooling a portion of the deflector more than the second plurality of injectors.

Abstract: The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole.

Abstract: A method and system for controlling combustion in a gas turbine combustor (18) includes adjusting an amount of a premix fuel portion (30) of a pilot fuel (28) provided to a premix burner stage (48) of a pilot (46) of the gas turbine combustor. The amount of the premix fuel portion is adjusted from a preset premix fuel portion amount to an adjusted premix fuel portion amount to achieve a desired first operating condition of the combustor. An amount of a diffusion fuel portion (34) of the pilot fuel provided to a diffusion burner stage (50) of the pilot is then adjusted from a preset diffusion fuel portion amount to an adjusted diffusion fuel portion amount to achieve a desired second operating condition of the combustor.

Abstract: 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.

Abstract: A fuel injector for a gas turbine engine includes a stem extending along a longitudinal axis from a proximal end to a distal end. The stem includes one or more fuel tubes configured to deliver a fuel toward the distal end of the stem, and a pilot assembly coupled to the distal end of the stem. The pilot assembly includes one or more components configured to inject fuel into a combustor of the engine. The fuel injector also includes a substantially tubular premix barrel extending along the longitudinal axis from a first end to a second end. The premix barrel is circumferentially disposed about the stem and configured to couple with the combustor at the second end. The fuel injector also includes an annular premix duct formed between the premix barrel and the stem. The premix duct includes an air inlet port at the first end, and a plurality of first orifices located downstream of the air inlet port. The first orifices are configured to inject fuel from the fuel tube into the premix duct.

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

Abstract: A method for operating a gas turbine engine including a combustor assembly is provided. The method includes channeling a first fluid through a first nozzle into the combustor assembly, igniting the first fluid within the combustor assembly downstream from the first nozzle, channeling a second fluid through a second nozzle into the combustion assembly when the gas turbine engine attains a speed of greater than 85% of rated speed, igniting the second fluid within the combustor assembly downstream from the second nozzle, terminating a flow of the first fluid through the first nozzle, and channeling the second fluid to the first nozzle.

Abstract: A gas turbine engine air valve assembly has first and second valving elements. The second element is rotatable about a first axis relative to the first element. The rotation controls a flow of air through the first and second elements. An actuator is coupled by a linkage to the second element. A plurality of follower member assemblies each have an insertion portion extending through a corresponding first aperture in the first valving element and a corresponding second aperture in the second element. The insertion portion is circumferentially fixed relative to one of the first and second valving elements and circumferentially moveable relative to the other of the first and second valving elements.

Abstract: A mixer assembly for use in a combustor of a gas turbine engine, including a pilot mixer, a main mixer, and a fuel manifold positioned between the pilot mixer and the main mixer. The pilot mixer further includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing. The main mixer further includes a main housing surrounding the pilot housing and defining an annular cavity, a plurality of fuel injection ports for introducing fuel into the cavity, and a swirler arrangement including at least three swirlers positioned upstream from the plurality of fuel injection ports, wherein each swirler of the swirler arrangement has a plurality of vanes for swirling air traveling through the respective swirler to mix air and the droplets of fuel dispensed by the fuel injection ports.

Abstract: Methods and systems for decomposing ammonia in a gasifier are provided. The system includes a gasifier including a pressure vessel and an injection nozzle extending through the pressure vessel and including a nozzle tip at a distal end of the injection nozzle, the injection nozzle further includes a passage configured to direct a fluid stream including ammonia proximate the nozzle tip such that the fluid stream including ammonia facilitates reducing a temperature of at least a portion of the nozzle tip.

Abstract: A gas turbine combustor includes a liquid fuel nozzle for jetting out liquid fuel; a pre-mixture chamber wall provided with the liquid fuel nozzle at a center thereof, having a hollow conical shape gradually spreading in the direction in which the fuel is jetted out, and defining a pre-mixture chamber therein; a plurality of air inlet holes bored through the pre-mixture chamber wall and introducing the combustion air to the pre-mixture chamber such that angles at which the combustion air is introduced to the pre-mixture chamber through the air inlet holes are deflected at least toward the circumferential direction of the pre-mixture chamber wall; and a plurality of gaseous fuel nozzles disposed around the pre-mixture chamber wall in an opposing relation respectively to the plurality of air inlet holes and jetting out gaseous fuel substantially coaxially with axes of the air inlet holes.

Abstract: 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.

Abstract: A method for controlling a combustion dynamics level within a combustion device includes defining a high dynamics operating state at a first fuel split ratio. The first fuel split ratio is a ratio of an amount of fuel supplied to the combustion device through a first fuel line to a total amount of fuel supplied to the combustion device. A low dynamics operating state is defined at a second fuel split ratio different from the first fuel split ratio. The second fuel split ratio is a second ratio of an amount of fuel supplied to the combustion device through the first fuel line to a total amount of fuel supplied to the combustion device. The combustion dynamics level within the combustion device is controlled by periodically switching between the first fuel split ratio and the second fuel split ratio.

Abstract: Methods and systems for operating a gas turbine engine system include an electrical generator configured to provide electrical energy to a load, a gas turbine engine including at least one combustor that includes a plurality of fuel injection points configured to inject a fuel into the combustor at a plurality of different locations wherein the combustor configured to combust the fuel and the gas turbine engine is rotatably coupled to the generator through a shaft. The gas turbine engine system includes a control system including a plurality of sensors positioned about the gas turbine engine system and configured to measure at least one parameter associated with the sensor, a processor programmed to receive a signal indicative of a heating value of the fuel, and automatically control a fuel split between the fuel injection points on the combustor using the determined heating value.

Abstract: A combustor assembly for a gas powered turbine includes a premix section to mix a first selected volume of fuel with a selected oxidizer. The premix section includes an injector plate that includes a porosity according to selected characteristics, such as pore size, pore density, pore distribution, and other selected characteristics. Therefore, the fuel may be provided through the porous plate to the premix area in a selected uniform flux.