Abstract: An aircraft system includes a propulsion system structure, a variable area nozzle and a linkage system. The propulsion system structure is moveable between a first position and a second position. The variable area nozzle is fluidly coupled with a duct in the propulsion system structure. The variable area nozzle is moveable between a first configuration and a second configuration. An exit area of the variable area nozzle has a first value when the variable area nozzle is in the first configuration and a second value when the variable area nozzle is in the second configuration. The variable area nozzle includes a flap moveably connected to the propulsion system structure. The linkage system mechanically links the variable area nozzle with the propulsion system structure. The linkage system includes a driver linkage, a first bell crank, a bridge linkage, a second bell crank, a follower linkage and a crank arm.

Abstract: An ignition and combustion assist system and method comprising a plasma igniter and electronic driver unit for use with gas turbine engines operating under low air densities, reduced voltage conditions and overall pressure ratios of 3:1 to 7:1. The plasma igniter has an inner chamber housing a centrally positioned and electrically isolated electrode attached to an electrical lead, driver unit, and AC or DC power supply. The electrode features a corner positioned near an outlet end of the igniter, where a plasma arc ignites a fuel-air mixture creating a flame extending into a primary burn region of a combustor of the gas turbine. The driver unit is in two embodiments and configured with low-cost microsecond voltage wave time periods or energy-efficient nano-second pulses. The method uses the plasma igniter and the electronic driver units described herein separately with other components or together.

Abstract: A combustor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a combustion chamber, and a fuel injector assembly in communication with the combustion chamber that has a swirler body situated about a nozzle to define an injector passage that converges to a throat. The throat is defined at a distance from the combustion chamber. The nozzle includes a primary fuel injector and an array of secondary plain jet fuel injectors.

Abstract: A method for detecting blowout precursors in at least one gas turbine combustor comprising: receiving combustion dynamics acoustic data measured by an acoustic measuring device associated with the combustor in real time; performing wavelet analysis on the acoustic data using simplified Mexican Hat wavelet transform analysis; and determining the existence of a blowout precursor based at least in part on the wavelet analysis. Provided also is a system and a non-transitory computer readable medium configured to perform the method.

Abstract: Discharge grate intended to be mounted inside or at the outlet of a conduit of a discharge valve of a turbine engine of an aircraft, the discharge grate comprising an upstream face intended to receive a gas flow, a downstream face parallel to the upstream face and intended to deliver the gas flow received on the upstream face, and orifices passing through the perforated plate from the upstream face to the downstream face and intended to convey the gas flow through the perforated plate. The discharge grate comprises for each orifice of the perforated plate a tubular channel, coaxial with the orifice with which it is associated, and projecting from the downstream face of the perforated plate.

Abstract: An ammonia decomposition facility includes a heating medium line configured to flow a heating medium heated by heat generated by a gas turbine, an ammonia supply line configured to flow ammonia, an ammonia decomposition device, and an ammonia removal device. The ammonia decomposition device is configured to use heat of the heating medium from the heating medium line, thermally decompose ammonia from the ammonia supply line, and generate a decomposition gas containing hydrogen, nitrogen, and residual ammonia. The ammonia removal device is configured to remove the residual ammonia contained in the decomposition gas from the ammonia decomposition device.

Abstract: Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Abstract: A last chance screen for a fuel system includes a mesh that includes a first plurality of members extending in a first direction and a second plurality of members extending in a second direction and intersecting the first members. Openings are formed in the mesh between the first plurality of members and the second plurality of members. The first plurality of members and the second plurality of members have an airfoil shaped cross-section.

Abstract: A pressure regulating valve assembly includes: a valve having an upstream side receiving an input flow and a downstream side providing an output flow, an actuator for opening and closing the valve, including partially opening the valve, and a regulator controlling the actuator to open, close or partially open the valve. The regulator includes a sense pressure port, wherein pressure at the port is maintained constant by the regulator. A chamber has a first entry orifice, a second entry orifice and an exit orifice. The first entry orifice is connected to the upstream side, the second entry orifice is connected to the downstream side, and the exit orifice is connected to the port. The exit orifice provides that the pressure at the exit orifice lies between the pressure at the first entry orifice and the pressure at the second entry orifice.

Abstract: A fuel injector includes an outer body and an inner body. The outer body extends about an axis and has a radially inner surface and a retention groove defined in the inner surface of the outer body. The inner body is positioned within the outer body has an outer surface and a retention tab. The retention tab retains the inner body relative to the outer body by engagement of the retention tab within the retention groove. A is axially offset from the retention tab and fixes the inner body within the outer body.

Abstract: A propulsion system assembly includes a variable area inlet and an inlet duct. The variable area inlet includes an outer airflow inlet passage, an inner airflow inlet passage, an inlet structure and a center body structure. The outer airflow inlet passage is between the inlet structure and the center body structure. The inner airflow inlet passage is formed within the center body structure. The center body structure includes a valve configured to regulate air flow through the inner airflow inlet passage. The valve includes a first door configured to pivot between a closed position and an open position. The inlet duct is configured to receive air from the outer airflow inlet passage when the first door is in the closed position. The inlet duct is configured to receive air from the outer airflow inlet passage and the inner airflow inlet passage when the first door is in the open position.

Abstract: A torch igniter for a combustor of a gas turbine engine includes an igniter body and an igniter head. The igniter body is disposed within a high-pressure case of a gas turbine engine and extends primarily along a first axis, and includes an annular wall and an outlet wall. The annular wall surrounds the first axis and defines a radial extent of a combustion chamber therewithin. The outlet wall is disposed at a downstream end of the annular wall, defines a downstream extent of the combustion chamber, and includes an outlet fluidly communicating between the combustion chamber and an interior of the combustor. The igniter head is removably attached to the igniter body at an upstream end of the annular wall, wherein the igniter head defines an upstream extent of the combustion chamber, and includes an ignition source and a fuel injector.

Abstract: The gas turbine system includes: a first gas turbine element 2; a second gas turbine element 3; a single combustor 4; a first supply pipe 61 which connects the first compressor 21 to the combustor 4; a second supply pipe 62 which connects the second compressor 31 to the combustor 4; a first discharge pipe 66 and a second discharge pipe 67 which discharge a fluid discharged from the combustor 4 to the outside; and a heat exchanger 5. The heat exchanger 5 allows each of a low-temperature fluid flowing through the first supply pipe 61 and the second supply pipe 62 and a high-temperature fluid flowing through the first discharge pipe 66 and the second discharge pipe 67 to flow therethrough and exchanges heat between the low-temperature fluid and the high-temperature fluid.

Abstract: A variable exhaust nozzle for use with a gas turbine engine includes an outer shroud and an inner plug that can move relative to the outer shroud. The relative movement of the inner plug and the outer shroud changes the shape of the variable exhaust nozzle from one that converges in area to one that converges and then diverges in area.

Abstract: A direct-fired gas turbine heater comprises a gas turbine engine, a main blower that receives cold air from the ambient, a mixing plenum that receives cold air from the main blower and hot gas from the turbine and delivers warm air, an air blower plenum that that receives cold air from the main air blower and delivers air to the mixing plenum, and an air intake plenum that receives cold air from the ambient and the air blower plenum and delivers cold air to a turbine compressor, an air intake valve, and an air starter valve. The gas turbine engine comprises the compressor that receives cold air, a fuel manifold that receives combustible fuel, a combustor that receives compressed air from the compressor and fuel from the fuel manifold, a turbine that receives hot gas from the combustor, and a shaft connecting the compressor and turbine.

Abstract: A torch ignitor system includes a torch wall defining a combustion chamber therein with a flame outlet passing out of the torch wall downstream of the combustion chamber. A fuel nozzle is mounted to the torch wall to issue fuel into the combustion chamber. An ignitor is mounted to the torch wall, extending into the combustion chamber to ignite fuel issued from the fuel nozzle. A cooling passage is in thermal communication with the ignitor for cooling the ignitor with fluid passing through the cooling passage.

Abstract: A gas turbine engine includes; a compressor, a combustor, and a turbine in serial flow relationship; a heat exchanger, the heat exchanger having an inlet, an outlet, and an internal surface coated with a catalyst, the heat exchanger being located upstream of the compressor; a source of hydrocarbon fuel in fluid communication with the inlet of the heat exchanger; a source of oxygen in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

Abstract: A torch ignitor system includes a torch wall defining a combustion chamber therein with a flame outlet passing out of the torch wall downstream of the combustion chamber. A fuel injector is mounted to the torch wall to issue fuel into the combustion chamber. At least one ignitor is mounted to the torch wall, positioned to ignite fuel issued from the fuel injector. A film cooler is defined through the torch wall and is operative to issue a film of cooling air from outside the torch wall along an interior surface of the combustion chamber for cooling the torch wall.

Abstract: A bleed valve for an air plenum includes a valve body having a poppet seat, a poppet supported by the valve body and movable relative to the poppet seat, and piston slideable relative to the poppet and movable relative to the poppet between an extended position and a retracted position. Gas turbine engines and methods of bleeding fluid from gas turbine engines are also described.

Abstract: A method and a system for providing in-flight reverse thrust for an aircraft are provided. The aircraft comprises an engine having a rotor, a compressor mechanically coupled to the rotor, and a variable geometry mechanism provided upstream of the compressor and configured to modulate an amount of compression work performed by the compressor. The method comprises operating the rotor with the variable geometry mechanism in a first position, receiving a request to increase reverse thrust for the rotor, in response to the request, adjusting the variable geometry mechanism from the first position towards a second position, the variable geometry mechanism having a greater opening angle in the second position than in the first position, and operating the rotor with the variable geometry mechanism in the second position for causing an increase in the amount of compression work performed by the compressor and an increase in reverse thrust for the rotor.