Abstract: A combustion chamber may include a first surface and a second surface interconnected by a wall forming a chamber having a central axis. The first surface may define an exhaust opening and the second surface defining a pilot opening, wherein the exhaust opening and the pilot opening align along the central axis. A plurality of inlet ports may be configured to deliver air to the chamber. A plurality of fuel ports may be arranged on an inside of the second surface to deliver fuel to the chamber. The air flow from the inlet ports and fuel from the fuel ports may oppose each other to create a vortex of product proximal to the second surface.

Abstract: A gas turbine engine includes an electrical system that includes a controller coupled to a first inverter/converter controller, a second inverter/converter controller, and a converter/controller that is coupled to an energy storage device. The system is configured to provide electrical power to a first electrical bus and a second electrical bus, from first and second electrical machines, under the direction of the controller. The converter controller is configured to control the amount of electrical power supplied to the first electrical bus and the second electrical bus from the energy storage system. The amount of electrical power received from the first electrical bus and the second electrical bus, and energy supplied to the energy storage system are under the direction of the controller.

Abstract: A control device for a thrust reverser of a turbojet engine nacelle, which includes movable portions actuated by actuators, includes: a first blocking control acting on a motor-actuated system; and a second blocking control acting, mechanically, on the movable portions. The second blocking control may inhibit, when activated, a deployment of the movable portions. The control device further includes a device inhibiting disengagement of the first blocking control when a mechanical control is activated.

Abstract: A control system is provided for performing staging control of a multi-stage combustor of a gas turbine engine. The fuel is fed to the combustor by a fuel supply system comprising: a plurality of fuel manifolds distributing fuel to respective stages of the combustor, a fuel metering valve operable to control the rate at which fuel passes to the fuel manifolds, and an actuating arrangement which splits the fuel flow from the fuel metering valve between the fuel manifolds. The control system includes a fuel pressure sensor arrangement which senses the fuel pressure at entry to the actuating arrangement, and/or in one or more of the fuel manifolds. The control system further includes a controller which repeatedly: calculates a fuel split between the fuel manifolds based on the sensed fuel pressure(s), and issues a command signal to the actuating arrangement to implement the calculated fuel pressure-based fuel split.

Abstract: A multi-surface blocker door system for a thrust reverser of an aircraft is provided. The system may comprise one or more first blocker doors coupled to the inner fixed structure. The first blocker doors may be configured to obstruct a fan duct area adjacent the engine. The system may also comprise one or more second blocker doors coupled to a translating sleeve. The second blocker doors may be configured to obstruct a fan duct area adjacent to an inlet of the nacelle.

Abstract: A disclosed intershaft seal assembly for a gas turbine engine includes a support fixed to a static structure. A first seal housing is supported radially outboard of the support for holding a first seal. A second seal housing is supported radially inboard of the support and supports a second seal. A first biasing member is provided between the support and the first seal housing that biases the first seal housing in a first direction away from the support. A second biasing member between the support and the second seal housing biases the second seal housing in a second direction away from the support.

Abstract: Providing mobile electric power comprising a power generation transport configured to convert hydrocarbon fuel to electricity and an inlet and exhaust transport configured to: couple to at least one side of the power generation transport such that the inlet and exhaust transport is not connected to a top side of the power generation transport, provide ventilation air and combustion air to the power generation transport, collect exhaust air from the power generation transport, and filter the exhaust air. In another embodiment, a fracturing pump transport comprising a first pump configured to pressurize and pump fracturing fluid, a second pump configured to pressurize and pump the fracturing fluid, and a dual shaft electric motor comprises a shaft and configured to receive electric power from a power source and drive in parallel, both the first pump and the second pump with the shaft.

Abstract: The present disclosure provides an integrated power generating system and method and liquefied natural gas (LNG) vaporization system and method. More particularly, heat from a CO2 containing stream from the power generating system and method can be used to heat the LNG for re-gasification as gaseous CO2 from CO2 containing stream is liquefied. The liquefied CO2 can be captured and/or recycled back to a combustor in the power generating system and method.

Abstract: A premixer assembly for mixing air and fuel for combustion includes a plurality of tubes disposed at a head end of a combustor assembly. Also included is a tube of the plurality of tubes, the tube including an inlet end and an outlet end.

Abstract: Methods are directed to operating a burner of a gas turbine. The burner includes a swirl generator and, downstream of it, a mixing tube. The swirl generator is defined by at least two walls facing one another to define a conical swirl chamber and includes nozzles arranged to inject fuel and apertures arranged to feed an oxidizer into the swirl chamber. The burner includes a lance which extends along a longitudinal axis of the swirl generator and side nozzles for ejecting a fuel within the burner. The side nozzles have their axes inclined with respect to the axis of the lance and can be positioned along the axis of the burner. During operation, an oil fuel or gaseous fuel is injected into to the burner through the lance tip and the lance side nozzles at various percentages depending on an operating mode of the gas turbine.

Abstract: A device and a method for bleeding compressor air in an engine includes at least one actuator and at least one closing element linked to the actuator for closing or partially closing a bypass duct via which compressor air can be bled off. It is provided here that the closing element is designed to be successively moved into the bypass duct, with the airflow passing through the bypass duct being settable by the position of the closing element. Furthermore, an air guiding device linked to the closing element is provided and has air guiding surfaces which adjoin the closing element downstream, the spatial alignment of the air guiding surfaces being dependent on the position of the closing element.

Abstract: A gas turbine engine includes a core engine, a first bypass passage disposed about the core engine and a second bypass passage disposed about the first bypass passage. A flow control is disposed within the second bypass for controlling bypass airflow through the second bypass. The flow control translates axially between an open position and a closed position to vary and control airflow through the second bypass passage.

Abstract: A fuel injector for a turbine engine, the injector including a body including a mechanism for admitting fuel under pressure, a stop valve for feeding a primary fuel circuit, and a metering valve mounted downstream from the stop valve for feeding a secondary fuel circuit. The injector further includes at least one leakage channel, that can be formed by a thread, that extends from a zone situated downstream from the stop valve and upstream from the metering valve to a zone situated downstream from the metering valve to generate a permanent leakage flow in the secondary circuit.

Abstract: A fuel injector for a turbine engine, the injector including a body including a primary fuel circuit and a secondary fuel circuit fed by a metering valve. A leakage channel extends from a zone in fluid flow connection with the primary circuit to a zone in fluid flow connection with the secondary circuit. The leakage channel is configured to be open in a closed position of the metering valve and to be closed by the metering valve moving.

Abstract: A combustor cap assembly includes an annular shroud and an impingement plate coupled to the shroud. The impingement plate at least partially defines a plurality of impingement cooling holes and a cooling flow return passage. A cap plate is coupled to the impingement plate. The cap plate includes an impingement side which faces a second side portion of the impingement plate where the impingement side is axially spaced from the second side portion to define an impingement air plenum therebetween. The cooling flow return passage is in fluid communication with the impingement air plenum. A fluid conduit extends from a first side portion of the impingement plate towards a first end portion of the shroud. The fluid conduit is in fluid communication with the cooling flow return passage and provides for fluid communication out of the impingement air plenum.

Abstract: A system and method of adaptively synchronizing the performance margin of a multi-engine system includes continuously, and in real-time, determining the performance margin of a first engine and the performance margin of the second engine. A difference between the performance margins of the first and second engines is calculated, and the first and second engines are controlled to attain a predetermined difference between the performance margins of the first and second engines.

Abstract: Transition duct assemblies for turbine systems and turbomachines are provided. In one embodiment, a transition duct assembly includes a plurality of transition ducts disposed in a generally annular array and comprising a first transition duct and a second transition duct. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct assembly further includes an aerodynamic structure defined by the passages of the first transition duct and the second transition duct. The aerodynamic structure includes a pressure side, a suction side, and a trailing edge, the trailing edge having a modified aerodynamic contour.

Abstract: The disclosure is directed to a swirler body for a combustor of a gas turbine engine, where the swirler body includes an annular mount face which defines at least one pocket. The disclosure is directed to a swirler assembly for a combustor of a gas turbine engine, where the swirler assembly includes a swirler first body with an annular first mount face which defines at least one first pocket and a swirler second body with an annular second mount face which abuts said annular first mount face, where said second annular mount face defines at least one second pocket. The disclosure is directed to a method of lightening a swirl assembly for a combustor of a gas turbine engine, where the method includes defining at least one pocket within an annular mount face of a swirler body.

Abstract: A combustor for a gas turbine including a casing, a flow sleeve at least partially disposed within the casing, a combustion liner at least partially disposed within the flow sleeve, a liner stop feature extending from the combustion liner, and a liner guide stop including a first end separated from a second end, the second end configured to be at least partially engaged with the liner stop feature, wherein the liner guide stop extends through the casing and the flow sleeve.

Abstract: A gas turbine engine includes a first annular portion that is stationary and adapted for partially surrounding an engine core. The first annular portion includes a fore pylon connecting portion. The gas turbine engine also includes a rail coupled to the fore pylon and extending in the aft direction from the first annular portion. The gas turbine engine also includes a second annular portion, arranged aft of the first portion and coupled to the rail. The second annular portion is movable along an engine core centerline between a closed position and at least one open position. The second annular portion is configured to engage the first annular portion in the closed position, thereby providing access to the engine core. The gas turbine engine further comprises a thrust reverser arranged in the second annular portion.