Abstract: Method and system for transitioning from a first liquid fuel supplied to a combustor of a gas turbine system to a second liquid fuel supplied to the combustor utilizing a fuel pump configured to supply both fuels from the fuel pump during the transition without shutting down the gas turbine system. Through utilizing a fuel pump and adjusting the operation of the fuel pump during the transition, the additional costs, size, and operational complexity associated with multiple liquid fuel trains to facilitate operation of the gas turbine system with multiple liquid fuels may be avoided. The utilization of multiple liquid fuels during operation of a gas turbine system increase the operational flexibility of the gas turbine system by enabling the utilization of varied liquid fuel sources and/or local liquid fuel sources during a steady-state operation, yet utilizing standardized fuel sources for startup or shutdown procedures.

Abstract: An apparatus comprises a combustion chamber of a gas turbine engine, a primary manifold for injecting fuel into the combustion chamber via a first plurality of nozzles and a secondary manifold for injection fuel into the combustion chamber via a second plurality of nozzles. A fuel control system provides fuel flow to the primary manifold and to the secondary manifold. The fuel control system determines at least one control parameter responsive to a first point at which the fuel flow level causes the secondary manifold to start draining and a second point at which an acceleration request is made and selects one of a plurality of secondary manifold filling profiles responsive to the determined at least one control parameter.

Abstract: A gas turbine engine includes a main compressor section and a main turbine section. A cooling air supply system cools a location in at least one of the main compressor section and the main turbine section. The cooling air supply system includes a tap for tapping cooling air compressed by the main compressor section, connected for passing the cooling air through a heat exchanger and to a boost compressor, and then to the cooling location in the at least one of the main compressor section and the main turbine section. A fuel supply system has a fuel tank for delivering fuel to a fuel pump. At least one valve for selectively returning fuel downstream of the main pump back to an upstream location. At least one return turbine drives at least one fluid moving device in the air cooling system.

Abstract: A combustor assembly for a gas turbine engine includes a dome having a forward surface and an inner surface. The forward surface and the inner surface of the dome at least partially define a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner is positioned within the slot of the dome. The forward end of the liner includes an axial interface surface and a radial interface surface. The axial interface surface defines a radial gap with the inner surface of the dome and the radial interface surface defines an axial gap with the forward surface of the dome. At least one of the radial gap or the axial gap is less than about 0.150 inches during operating conditions of the combustor assembly to prevent an undesirable airflow.

Abstract: A fuel swirler has a swirler housing defining an interior chamber having a fuel outlet at a downstream end. A swirler core is mounted inside the interior chamber. The swirler core has a downstream end portion with one or more fuel channels disposed thereon. The one or more fuel channels is in fluid communication with the fuel outlet. An internal bore extends longitudinally through the swirler core. The internal bore has an inlet connectable to a source of fuel. One or more exit holes fluidly connects the internal bore to the one or more fuel channels. An annular air gap is provided radially between the swirler housing and the swirler core for thermally shielding the internal bore.

Abstract: A gas turbine plant includes a gas turbine, a liquefaction facility capable of liquefying air, and a liquefaction controller. A compressor has an intake amount adjuster capable of adjusting an intake amount into a compressor casing. The liquefaction facility includes: a bleed line capable of bleeding compressed air from the compressor; a liquefaction system capable of liquefying the compressed air, a bleed amount adjustment valve; a return air line capable of guiding return air into a flow passage through which compressed air flows in the gas turbine; and a return amount adjusting valve. The liquefaction controller opens the bleed amount adjustment valve if an opening degree of the intake amount adjuster is a first opening degree, and opens the return amount adjusting valve if the opening degree of the intake amount adjuster is a second opening degree, which is an opening degree greater than the first opening degree.

Abstract: The disclosure relates to a jet engine. The jet engine includes: an engine casing; and a core engine arranged in the engine casing, having a core engine casing, and including a compressor, a turbine, and a combustion chamber arranged between the compressor and the turbine. The compressor is internally provided with a compressor blade, the turbine is internally provided with a turbine blade, radial outer edges of the compressor blade and the turbine blade are connected to the core engine casing, and a space constructed between the engine casing and the core engine casing forms an engine-mounted fuel tank for accommodating fuel.

Abstract: A fuel injector has a body extending in an axial direction. The body includes: an axial passage formed so as to extend in the axial direction; a radial passage formed so as to communicate at one end with the axial passage and to open at another end to an outer surface of the body; and an internal passage including a first opening and a second opening open to the outer surface, and formed so as to extend inside the body from the first opening to the second opening. The first opening and the second opening are located opposite to each other across a third opening, through which the radial passage opens to the outer surface, in a circumferential direction centered on an axis of the body.

Abstract: An injector for a combustor of a gas turbine engine is provided with a plurality of air/fuel mixing tubes divided into radially outer and radially inner subsets of air/fuel mixing tubes with a first fuel manifold in fluid communication with the radially outer subset of air/fuel mixing tubes and a second fuel manifold in fluid communication with the radially inner subset of air/fuel mixing tubes. Each of the air/fuel mixing tubes of the radially outer subset of air/fuel mixing tubes includes a substantially quadrilateral cross-sectional profile, and each of the air/fuel mixing tubes of the radially inner subset of air/fuel mixing tubes includes a substantially circular cross-sectional profile.

Abstract: An architecture for supplying air to a high-pressure compressor of an auxiliary gas generator from a pressurized cabin of an aircraft, comprising: a load compressor rotationally driven by a common rotation shaft providing a mechanical coupling between the high-pressure compressor and a high-pressure turbine of the auxiliary gas generator and supplied by an outside air intake; a first regulation valve assembled at the outlet of the load compressor to control all or part of the air flow delivered by the load compressor; a second regulation valve assembled at the outlet of the pressurized cabin to control the air flow drawn from inside the pressurized cabin; a mixer receiving the outputs of the first and second regulation valves to add the air drawn from inside the pressurized cabin to all or part of the air delivered by the load compressor.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a supplemental thrust section and a duct. The supplemental thrust section includes a rotating detonation combustor. The duct includes a supplemental thrust section inlet fluidly coupled with and leading to the rotating detonation combustor. The supplemental thrust section inlet has a flow area that decreases as at least a first portion of the supplemental thrust section inlet extends towards the rotating detonation combustor.

Abstract: A gas turbine engine component includes a first surface and a second surface. The component further includes a dilution hole defined by the first surface and the second surface. The component further includes a first effusion hole and a second effusion hole each having an inlet defined by the second surface and an outlet defined by the first surface such that the outlet of the first effusion hole is located nearer to the dilution hole than the outlet of the second effusion hole.

Abstract: An embodiment of a combustor for a gas turbine engine includes a combustor case, a combustor liner disposed within the combustor case, a fuel nozzle, and a torch igniter within the combustor case. The torch igniter includes a combustion chamber, a cap defining the upstream end of the combustion chamber and configured to receive a fuel injector and a surface igniter, a tip defining the downstream end of the combustion chamber, an annular igniter wall extending from the cap to the tip and defining a radial extent of the combustion chamber, a structural wall coaxial with and surrounding the igniter wall, and an outlet passage within the tip that fluidly connects the combustion chamber to the combustor. The torch igniter is situated such that the tip is mounted through the combustor liner, the combustion chamber is within the combustor case, and the cap extends through the combustor case.

Abstract: A dry low NOX staged combustion system includes a fuel nozzle and a combustion compartment. The fuel nozzle includes a purge gas tube, a diffusion combustion fuel tube, an isolation gas tube, a premixed combustion fuel tube, a premixed combustion air tube. The purge gas tube is configured to feed a purge gas. The diffusion combustion fuel tube is fitted over the purge gas tube, and having an end provided with a diffusion combustion fuel swirler. The isolation gas tube is fitted over the diffusion combustion fuel tube. The premixed combustion fuel tube is fitted over the isolation gas tube. The premixed combustion air tube is fitted over the premixed combustion fuel tube. The combustion compartment is located downstream of the fuel nozzle.

Abstract: Systems and methods including a plurality of reformer stacks extended around the combustion chamber. The reformer stacks are distributed along a length of the combustion chamber in the axial direction and configured to provide output products to the combustion chamber.

Abstract: An electronic speed-controlled pulsed supersonic turbine engine powering automotive, drone, and electric power generation, energized by breathable, clean renewable energy airflow from 2700 psi integral air-tank energizing the engine continuously for 3 hours, replacing the toxic fossil gasoline-diesel energized internal combustion engine with carbon emissions that affect climate change. The turbine blades are turned by the pulsed impulse of supersonic airflow from sequentially energized eight manifolds of de Laval convergence-divergence-CD with sonic choking nozzle and supersonic divergence airflow impulsing turbine blades turning them within divergence shroud to atmospheric pressure with turbine nose with engine output shaft supported with bearings supported by the air-tank.

Abstract: A fuel system for an aircraft engine comprises a fuel metering unit and a separate flow divider. The flow divider has an inlet port fluidly connected to the fuel metering unit via a fuel line. A primary and a secondary fuel manifold are fluidly connected to the flow divider. The fuel metering unit and the flow divider have a fuel supply mode in which fuel is allowed to flow in a first direction through the fuel line from the fuel metering unit to the flow divider to feed the primary and secondary fuel manifolds, and an ecology mode in which fuel is allowed to flow in a second direction through the same fuel line from the flow divider towards the fuel metering unit. A same fuel line is thus used as a fuel supply line and an ecology line.

Abstract: A combustion device for an electricity production or energy cogeneration apparatus, the apparatus comprising a gas turbine supplied by the device, the device being suitable for a combustion regime of a “flameless” type, the device including: an outer tube; and a combustion tube forming a combustion zone for flameless combustion of a mixture of oxidizing air and fuel, the combustion tube being concentric with the outer tube, in communication with fuel-injection means, and with air-injection means arranged at a first end of the combustion tube, referred to as a front end, and which is closed by an end wall at a second end of the combustion tube, referred to as a rear end, the end wall being secured to the combustion tube and fluidtight, the fuel-injection means including at least a first orifice, referred to as a fuel-injection orifice, the air-injection means including at least a second orifice.

Abstract: An injection nozzle installed in a combustor of a gas turbine to inject fuel and compressed air into a combustion chamber is provided. The injection nozzle includes an inlet portion through which fuel and compressed air are introduced, an outlet portion disposed downstream of the inlet portion in a flow direction of fluid and configured to discharge the fuel and compressed air to the combustion chamber, and an intermediate portion connected obliquely to each of the inlet portion and the outlet portion, wherein the inlet portion and the intermediate portion are symmetrically formed based on an imaginary central line extending from the outlet portion along the flow direction of fluid and passing through a center of the outlet portion.

Abstract: A bleed air control system is configured to vary the air pressure at the inlet of a gas turbine engine. The bleed air control system includes a first gas turbine engine configured to provide bleed air, a second gas turbine engine acting as an auxiliary power unit, and a bleed air control system configured to selectively provide bleed air from the first gas turbine engine to the second gas turbine engine.