Abstract: An aircraft fuel cell system is provided. The system includes a fuel tank, a reactor for generating hydrogen gas from a fuel, a heating apparatus and a fuel cell. The reactor can process a synthetic fuel produced from biomass. The use, in an aircraft, of a synthetic fuel, produced from biomass, for generating a gas that contains hydrogen is also provided.

Abstract: The invention relates a method for mixing a dilution air with a hot main flow in a 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 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, wherein the hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery.

Abstract: A device and system for monitoring and controlling the flow media through a flow circuit, for example, the flow of coolant through a coolant circuit of a turbine engine fuel system, measures flow rates at the supply and return sides of the circuit and compares the measured flow rates against a threshold flow rate difference to detect the presence of a flow leak condition. Upon detecting a leak condition, the controller automatically operates a flow control valve to shut off flow to the flow circuit for repair of the leak or other low flow condition. The controller can also include components for measuring the temperate of the flow media at one or both of the supply and return sides of the circuit. The measured temperature signals can be processed to detect a high or low temperature condition. Upon detecting a low temperature condition, the controller automatically reduces the flow rate of the flow media until the condition has been corrected.

Abstract: A method for controlling a gas turbine combustion system includes supplying an ultra low calorific fuel to a combustor of the combustion system through a first fuel circuit, controlling a supply of the ultra low calorific fuel through a second fuel circuit as required to control the volumetric flow of the ultra low calorific fuel through the combustor, and combusting the ultra low calorific fuel in the combustor.

Abstract: Systems and methods for controlling fuel flow within a machine are provided. A plurality of fuel types provided to the machine and a plurality of fuel circuits associated with the machine may be identified, each of the plurality of fuel circuits adapted to be provided with one or more of the plurality of fuel types. A fuel flow parameter for calculating fuel flow may be identified, and a respective fuel flow for each of the one or more fuel types provided to each of the plurality of fuel circuits may be calculated based at least in part on the identified fuel flow parameter. Based at least in part on the calculation of the respective fuel flows, operation of one or more fuel flow control devices providing fuel to the plurality of fuel circuits may be controlled.

Abstract: This aircraft maintenance equipment is a fuel system leak test unit, designed to insure installation integrity of fuel system lines and hoses and to insure that the system is free from any debris generated during the assembly and maintenance process for an aircraft. The unit is designed to check for leaks using vacuum or pressure while monitoring for fuel leaks or bubbles through a clear tube section.

Abstract: A condition measurement apparatus is provided and includes a gas turbine engine combustor having an end cover, a liner defining a liner interior and a fuel nozzle communicative with the liner interior, the end cover being formed to separate a cold side thereof, which is a relatively low temperature environment, from a hot side thereof, which is a relatively high temperature environment in which the liner and the fuel nozzle are disposed, the combustor being formed to define a fuel flow path extending through piping disposed at the cold side of the end cover by which fuel is deliverable to the fuel nozzle, and a condition sensing device operably mounted on the piping.

Abstract: A fuel nozzle assembly for use with a turbine engine is provided. The fuel nozzle assembly includes a tube that is configured to channel at least a first type of fuel through the turbine engine. A flow member includes a first end portion and a second end portion that is removably coupled to the tube such that the flow member is severally removable from the tube to enable a plurality of different types of fuel to be channeled through the turbine engine for operation of the turbine engine.

Abstract: A gas turbine engine combustor includes an annulus with one or more circular rows of burners and a means for providing a number of equiangular spaced apart flame temperature nonuniformities around the annulus during engine operation. The number of the flame temperature nonuniformities being equal to a circumferential acoustic mode to be attenuated in the combustor (i.e three, five, or seven). Fuel lines and/or water lines in supply communication with the burners and metering orifices in a portion of the fuel lines and/or the water lines may be used to produce the flame temperature nonuniformities. The annulus of the burners may have an equal number of equiangular spaced apart first and second arcuate segments of the burners and a means for operating the burners in the first segments and operating the burners in the second segments at different first and second flame temperatures respectively.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a plurality of fuel injectors circumferentially arranged around the combustion chamber. The plurality of fuel injectors provide fluid communication into the combustion chamber, and a separate cap for each fuel injector defines a separate volume around a different fuel injector outside of the combustion chamber.

Abstract: A turbomachine including: a combustion chamber, with a fuel injection device in the combustion chamber; a supply mechanism supplying fuel to the fuel injection device; a mechanism determining instantaneous variation of fuel flow rate of the supply mechanism; and a regulation mechanism regulating the fuel flow rate of the injection device according to the instantaneous variation of the fuel flow rate of the supply mechanism determined by the determination mechanism.

Abstract: An LNG fuel system for gas turbine engine systems is disclosed that allows more efficient management of cryogenic fuels such as LNG to reduce emissions and improve engine efficiency. In one configuration, an intercooled, recuperated gas turbine engine comprises an LNG tank incorporating a liquid-to-vapor LNG fuel circuit in parallel with a vapor fuel circuit. In a second configuration, an alternate vapor fuel circuit is disclosed. In either configuration, the fuel in the liquid fuel circuit is vaporized and heated by the engine's intercooler or by both the engine's intercooler and/or a heat exchanger on the exhaust. In another configuration, both the fuel in the liquid-to-vapor LNG fuel circuit and the vapor fuel circuit are heated by a heat exchanger on the exhaust.

Abstract: Methods and systems are provided for premixing combustion fuel and air within gas turbines. In one embodiment, a combustor includes an upstream mixing panel configured to direct compressed air and combustion fuel through a premixing zone to form a fuel-air mixture. The combustor also includes a downstream mixing panel configured to mix additional combustion fuel with the fuel-air mixture to form a combustion mixture.

Abstract: A combustor may include a number of nozzles, a first fuel source with a low reactivity fuel therein, a second fuel source with a high reactivity fuel therein, and a primary valve for varying the flow of the low reactivity fuel and the high reactivity fuel delivered to the nozzles.

Abstract: A gas turbine combustion chamber includes a fuel nozzle with a cylindrical nozzle tube, in which a fluid flows, and a convexly formed nozzle cover, which is arranged downstream of the nozzle tube. The nozzle cover has a central point and a plurality of through-openings through which the fluid leaves the nozzle tube. The through-openings are arranged at different radial distances from the central point at two circular lines.

Abstract: A system provides “fail fixed” functionality and allows a user to manually manipulate fuel flow to a gas turbine engine in the unlikely event the primary control means is unavailable. The fuel metering unit includes a fuel metering valve, a flow increase valve, and a flow decrease valve. The flow increase valve and flow decrease valves are both in fluid communication with the fuel metering valve and are each adapted to selectively receive fuel flow commands from a primary fuel flow command source and from a secondary fuel flow command source. The flow increase and decrease valves are responsive to the fuel flow commands to selectively control the position of the fuel metering valve.

Abstract: The present application thus provides a combustor for use with a gas turbine engine. The combustor may include a turbine nozzle and a liner cooling system integral with the turbine nozzle. The liner cooling system may include a liner with one or more cooling features thereon and an impingement sleeve.

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: According to one aspect of the invention, an ice separating apparatus for a gas turbine engine includes a can having a first end and a tangential inlet proximate the first end of the can to receive a fluid flow within the can, wherein the fluid flow includes fuel and ice. The apparatus also includes a first conduit within the can to receive a separated fuel flow proximate a second end of the can, wherein the fluid flow separates ice from fuel to form the separated fuel flow.

Abstract: Embodiments of the disclosure can provide systems and methods for minimizing coking in gas turbine engines. According to one embodiment, there is disclosed a system for minimizing coking in a gas turbine engine. The system may include a gas turbine compartment, a fuel component disposed within the gas turbine compartment, and a thermoelectric element disposed at least partially about the fuel component. The thermoelectric element may be configured to exchange heat with the fuel component.

Abstract: A fuel metering system for supplying fuel to load includes a variable displacement piston pump having an adjustable hanger that is movable to a plurality of positions. The variable displacement piston pump is configured to receive a drive torque and, upon receipt of the drive torque, to supply fuel to the plurality of loads at a flow rate dependent on the position of the adjustable hanger. A hanger actuator is coupled to receive hanger position commands and is operable, in response thereto, to move the adjustable hanger to the commanded position.

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 system includes a turbine engine having a fuel injector. The fuel injector includes fluid ducts, each having a fuel inlet coupled to a distinct fuel source. The system includes a compressed air source that provides compressed air simultaneously to the fluid ducts, and a convergence point where combined fuel and air streams from the ducts are mixed. The fuel inlets are in a parallel flow arrangement such that no fuel from one fuel injector is present at another fuel injector.

Abstract: A combustor assembly for a gas turbine engine includes a static structure and an annular combustor extending around a central axis and located radially inwards of the static structure. The annular combustor includes an annular outer shell and an annular inner shell that define an annular combustion chamber there between. The annular combustor is free of any rigid attachments directly between the static structure and the annular outer shell. The annular outer shell includes a radially-outwardly extending flange. A stop member is rigidly connected with the static structure and is located adjacent the radially-outwardly extending flange such that axial-forward movement of the annular combustor is limited.

Abstract: A fluid delivery system in a gas turbine engine includes an electric motor, a first fluid pump system and a second fluid pump system. The electric motor operates at a variable speed. The first pump system includes a first pump driven by the motor to deliver a first fluid to the gas turbine engine. The second pump system includes a second pump driven by the motor to deliver a second fluid to the gas turbine engine, the second pump having a variable displacement.

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 fuel system including a fuel pump metering unit (FPMU) for delivering fuel to an engine manifold with an ecology valve for draining and storing fuel from the engine manifold. The ecology valve includes a housing having a piston dividing the housing into a first side in fluid communication with an output of the FPMU and a second side in fluid communication with the engine manifold. An assembly connected between the FPMU and engine manifold selectively creates a pressure differential across the first and second side of the housing when the FPMU delivers fuel to the engine manifold. In a run position, the piston moves to decrease a volume within the housing interior as a result of the pressure differential. In a drain position, the piston moves to increase the housing volume within the interior and thereby pull and store fuel from the engine manifold.

Abstract: A system for recirculating a vented fuel to a fuel delivery system of a gas turbine generally includes a vented fuel conduit in fluid communication with the fuel delivery system and a container in fluid communication with the vented fuel conduit. A pump is disposed downstream from the container and a fuel conduit is disposed downstream from the pump. The fuel conduit may provide fluid communication from the pump to the fuel delivery system.

Abstract: 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.

Abstract: A combustion swirler is provided. The combustion swirler includes multiple vanes axially extending from an annular first body portion of the combustion swirler. The combustion swirler also includes an annular second body portion enclosing the multiple vanes for directing a flow of combustion fluid. Each of the vanes comprises an aerodynamic blade body comprising a leading edge with a plurality of first tubercles and a trailing edge with a plurality of second tubercles.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a head end with a non-circular configuration, a number of fuel nozzles positioned about the head end, and a transition piece extending downstream of the head end.

Abstract: An object is to provide a low cost gas turbine combustor. The gas turbine combustor has a plurality of main nozzles annularly supported on a nozzle pipe base (11), and the main nozzles each have an oil fuel path for supplying oil fuel and a gas fuel path for supplying gas fuel. A plurality of the oil fuel paths are provided to penetrate the nozzle pipe base (11), and a plurality of lengths of upright piping (21) are provided to be connected to the oil fuel paths, respectively, and be erected on the nozzle pipe base (11). A plurality of lengths of connection piping (22) for interconnecting the lengths of the upright piping (21) in a polygonal shape at a minimum distance are provided outside the nozzle pipe base (11). An oil fuel supply section (14) is connected to one of the lengths of the connection piping (22) to distribute oil fuel to each of the oil fuel paths.

Abstract: A fuel manifold assembly for a gas turbine engine is described and includes an annular manifold body and an annular heat shield assembly mounted to and surrounding the manifold body. The heat shield assembly has an inner surface facing the manifold body and spaced apart therefrom by an inner air gap, which substantially surrounds the manifold body, defined therebetween. The heat shield assembly has an outer surface facing away from the manifold body and spaced apart from the inner surface by an outer air gap substantially surrounding the inner air gap. At least the outer air gap is formed by a double wall configuration of the heat shield assembly.

Abstract: Systems and methods for controlling fuel flow to a turbine component are provided. One or more parameters associated with a fuel flow to a turbine component may be monitored. The fuel flow may be modeled based at least in part on the one or more monitored parameters. The fuel flow may be adjusted to a target flow based at least in part on the modeling of the fuel flow.

Abstract: The fuel flow rate to a gas turbine is measured using an inert gas. The inert gas is injected into the fuel flow and the concentration of the inert gas in the fuel/inert gas mixture is later measured. The concentration of the inert gas in the fuel/inert gas mixture is then used to calculate the mass flow rate of the fuel.

Abstract: An architecture for feeding fuel to a power plant (1) of a rotorcraft, which power plant comprises a plurality of engines (2, 3) individually fed with fuel by respective assemblies (4, 5). Each assembly (4, 5) comprises a fuel feed circuit (9, 10) for feeding a safe tank (6, 7) from a fuel tank (8) that is common to the assemblies (4, 5). Together the feed circuits (9, 10) form a circuit for forced both-way transfer of fuel from either one of the assemblies to the other via the common tank (8) and an intercommunication (21) interposed between the safe tanks (6, 7), which safe tanks are fitted with respective spillage devices (17, 18, 19, 20) for transferring excess fuel to the common tank (8).

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: The present invention relates to an annular combustion chamber of a gas turbine with—relative to the engine axis—a radially outer combustion chamber wall and a radially inner combustion chamber wall, with the combustion chamber walls forming an annular combustion space, with a combustion chamber head having a plurality of fuel nozzles and air inlet openings, with the respective central axes of the fuel nozzles forming an envelope rotationally symmetrical to the engine axis, the envelope dividing the combustion chamber into an annular and radially outer area and an annular and radially inner area, with the radially outer area and the radially inner area having the same volumes.

Abstract: Embodiments of a system are disclosed that include a heat source, an endothermic process module, and a fuel source configured to supply fuel to the endothermic process module and to receive isomerized fuel from the endothermic process module. A controller includes logic instructions operable to receive information regarding temperature of fuel received by the endothermic process module, and regulate application of heat from the heat source to the fuel at the endothermic process module. The endothermic process module includes a catalyst that increases the thermal carrying capacity of the fuel by isomerizing fuel from the fuel source.

Abstract: The invention relates to a method for monitoring the operation of a fuel circuit shut-off valve comprising an LPSOV, a flow regulator, the shut-off valve, characterised in that it comprises: a step for ordering (30) the closure of the valves; a step for measuring (32) the fuel flow rate in the circuit qui carried out before the LPSOV is completely closed; and a diagnostic step (34) consisting of determining that said shut-off valve is defective if the fuel flow rate measured is not equal to zero and determining that said shut-off valve is operating correctly if the fuel flow rate measured is zero.

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: The invention relates to a device for regulating the flow rate of fuel taken from a fuel circuit of an aircraft propelled by an engine. The device comprises: a positive displacement pump receiving fuel from a fuel circuit of the aircraft and delivering fuel at a flow rate that is proportional to its speed of rotation; a differential gear having a predefined transmission ratio, a first inlet mechanically coupled to the aircraft engine, a second inlet mechanically coupled to an electric motor/generator, and an outlet mechanically coupled to the positive displacement pump to drive it in rotation; and an electronic control system for regulating the speed of rotation of the electric motor/generator as a function of a setpoint value for the flow rate of fuel to be injected.

Abstract: Embodiments of the present disclosure provide a turbine combustor having a primary fuel injection system, a first wall portion disposed about a primary combustion zone downstream from the primary fuel injection system, and a second wall portion disposed downstream from the first wall portion. The turbine combustor also has a secondary fuel injection system disposed between the first wall portion and the second wall portion, where the secondary fuel injection system is removable form the first and second wall portions.

Abstract: A combustor is provided in which a fuel and working fluid can be injected in an annulus. In one form the fuel and working fluid is circumferentially flowed within the annulus and traverses the annulus in an axial direction from one side to another side where the flow exits. The working fluid and air can be co-axially admitted to the combustor and in one form the working fluid can be swirled about the fuel dispensed from a fuel injector. The combustor can provide for a rich burning zone. In one embodiment the combustor is configured as an inter-turbine combustor having an outlet at one axial side of the combustor. A lean burn region can be created within a flow path of the turbine.

Abstract: A fuel shroud assembly (100) into which fuel (118) is injected for mixing with an air stream (120) in a fuel manifold. The shroud assembly (100) comprises a plurality of parallel fuel scoops (102) each receiving the injected fuel (118). The fuel stream (118) flows through each scoop (102), exiting at an open scoop end (114A). The air stream (120) flows between scoops, creating a shear region proximate each scoop end (114A) where the fuel exits. The shear causes mixing of the air (120) and the fuel (118) streams, wherein the degree of mixing is not dependent on the momentum ratio of the air (120) or fuel (118) streams.

Abstract: Embodiments of this invention provide a premixed pilot assembly for use with a fuel nozzle for a turbine. The level of nitrogen oxide (NOx) emitted by the pilot is reduced by mixing the fuel and air fast, at the end of the pilot nozzle, thereby avoiding significant zones of rich fuel and air mixtures. The air is mixed with the fuel through the use of openings in a first cylinder and a second cylinder, one of which carries air and one of which carries pilot fuel. The openings can be configured as necessary to obtain a desired effect on the pilot.

Abstract: A combustion chamber in which the injection system is attached so as to prevent any axial motion of the injection system, the combustion chamber includes: a baffle including a tubular portion having a first upstream end surrounded by a first radially projecting collar; and an injection system including a bowl which includes a cylindrical portion inserted into the tubular portion, the cylindrical portion including a second upstream end surrounded by a second radially projecting collar, the second collar axially bearing against the first collar, wherein the combustion chamber includes a clamp which axially clamps the first and the second collars against one another, and a fastener which retains the clamp clamped on either side of the first and second collars.

Abstract: A fuel delivery system for a gas turbine engine includes one or more subsets of combustors, and at least two fuel manifolds including a fuel manifold coupled to each combustor subset and configured to deliver fuel to only the subset of combustors during a predetermined gas turbine engine operating mode. A gas turbine engine assembly including a fuel delivery system and a method of operating a gas turbine engine is also described herein.

Abstract: An endcover for a turbine combustor adapted to support one or more combustor nozzles, includes a plate having one side which in use, faces a combustion chamber and an opposite side which, in use, faces away from the combustion chamber. At least one fuel cavity is formed in the plate; and a fuel restrictor insert is formed with at least one flow orifice located within the fuel cavity for supplying fuel to at least one combustor nozzle. The fuel restrictor insert is adjustable along a length dimension of the fuel cavity.

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.