Abstract: Fairing installations disclosed herein may include a damper for mitigating vibration of a cantilevered fairing disposed in a bypass duct of a gas turbine engine. The bypass duct may include a first shroud radially spaced apart from a second shroud to define a bypass passage between the first and second shrouds. The fairing may be disposed in the bypass passage and cantilevered from the first shroud. The fairing may have a secured end secured to the first shroud and a free end proximate the second shroud. The damper may be engaged with the free end of the fairing to damp movement of the free end of the fairing.

Abstract: A turbine engine is disclosed that includes a fan case surrounding a fan rotatable about an axis. A core is supported relative to the fan case by support structure arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage having a rotor blade with a blade trailing edge. The support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge. The support structure trailing edge is arranged axially forward of the blade trailing edge. In one example, a forward attachment extends from the support structure to the inlet case.

Abstract: A turbofan engine assembly including a compressor, an intermittent internal combustion engine having an inlet in fluid communication with an outlet of the compressor through at least one first passage of an intercooler, a turbine having an inlet in fluid communication with an outlet of the intermittent internal combustion engine, the turbine compounded with the intermittent internal combustion engine, a bypass duct surrounding the intermittent internal combustion engine, compressor and turbine, and a fan configured to propel air through the bypass duct and through an inlet of the compressor, wherein the intercooler is located in the bypass duct, the intercooler having at least one second passage in heat exchange relationship with the at least one first passage, the at least one second passage in fluid communication with the bypass duct.

Abstract: A combustor for a gas turbine engine includes a combustion chamber defined between an inner shell and an outer shell. The combustor further includes a bulkhead extending between the inner shell and the outer shell. The bulkhead includes a plurality of impingement cooling rings. Each impingement cooling ring of the plurality of impingement cooling rings includes a plurality of impingement cooling holes extending through the bulkhead. The combustor further includes a heat shield panel mounted to the bulkhead so as to define an impingement cooling chamber between the bulkhead and the heat shield panel. The heat shield panel further includes a radial portion between a perimeter and an opening, with respect to an opening center axis, which is free of penetrations. The plurality of impingement cooling holes of each of the plurality of impingement cooling rings are directed toward the radial portion of the heat shield panel.

Abstract: A module (4) for an aircraft turbomachine comprises an assembly of constant-volume combustion chambers, and including a first sub-assembly of first chambers succeeding each other along a given sense (76) and forming series of chambers (S1), and within each series (S1), a first ignition chamber (C1.1) located at one of both circumferential ends of the series is defined, the first ignition chamber (C1.1) being connected to the first directly consecutive chamber (C1.2) along the given sense (76) so as to supply the same with exhaust gases, and so forth up to the first chamber (C1.3) located at the other circumferential end of the series. In addition, a control device (46) is configured such that for all the first ignition chambers (C1.1), diametrically opposite two by two, the combustion cycles are simultaneously initiated. Finally, a second sub-assembly comprising second combustion chambers (C2.1-C2.3) is also provided.

Abstract: Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to the combustor by supplying fuel to the first fuel manifold via a first flow divider valve and supplying fuel to the second fuel manifold via a second flow divider valve. While supplying fuel to the combustor by supplying fuel to the first fuel manifold, the method includes stopping supplying fuel to the second fuel manifold and supplying pressurized gas to the second fuel manifold via the second flow divider valve to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated nozzles.

Abstract: A cooling system configured to reduce a temperature within a gas turbine engine in a shutdown mode of operation includes a first gas turbine engine including a compressor having a bleed port. In a first operating mode of the gas turbine engine, the compressor bleed port is configured to channel a high pressure flow of air from the compressor. During a shutdown mode of operation, the compressor bleed port is configured to channel an external flow of cooling air into the compressor. The cooling system also includes a source of cooling air and a conduit coupled in flow communication between the compressor bleed port and the source of cooling air. The source of cooling air configured to deliver a flow of cooling air into the compressor through the compressor bleed port.

Abstract: An anti-icing system for a gas turbine engine comprises a closed circuit containing a phase-change fluid, at least one heating component for boiling the phase-change fluid, the anti-icing system configured so that the phase-change fluid partially vaporizes to a vapour state when boiled by the at least one heating component. The closed circuit has an anti-icing cavity adapted to be in heat exchange with an anti-icing surface of the gas turbine engine for the phase-change fluid to release heat to the anti-icing surface and condense. A feed conduit(s) has an outlet end in fluid communication with the anti-icing cavity to feed the phase-change fluid in vapour state from heating by the at least one heating component to the anti-icing cavity, and at least one return conduit having an outlet end in fluid communication with the anti-icing cavity to direct condensed phase-change fluid from the anti-icing cavity to the at least one heating component.

Abstract: An apparatus and method of operating a translating cowl for a turbine engine. The translating cowl is moveable between a first position and a second position. The translating cowl includes a fixed cascade element located within and a blocker door that is operably coupled to die translating cowl. Hie blocker door is movable between a stowed position and a deployed position.

Abstract: A ducted heat exchanger system for a gas turbine engine includes an additive manufactured heat exchanger core with a contoured external and/or internal geometry. A method of additively manufacturing a heat exchanger for a gas turbine engine includes additively manufacturing a core of a heat exchanger to set a ratio of local surface area to flow area to control a pressure drop per unit length along the core.

Abstract: An example hybrid aircraft propulsion system includes one or more power units configured to output electrical energy onto one or more electrical busses; a plurality of propulsors; and a plurality of electrical machines, each respective electrical machine configured to drive a respective propulsor of the plurality of propulsors using electrical energy received from at least one of the one or more electrical busses.

Abstract: The invention relates to a bottom wall (118) of a gas turbomachine combustion chamber. This bottom wall comprises openings (119) for mounting combustion air supply systems and holes (128) for the passage of cooling air between at least one inlet (128a) and at least one outlet of said holes. The holes (128) extend inwardly along the bottom wall with the outlet port located closer to the opening (119) adjacent to it than the inlet hole.

Abstract: A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: determining when a flame in a combustion chamber of a gas turbine engine is extinguished, during a start-up process or re-light process or during operation; purging the combustion chamber by controlling rotation of a low pressure compressor using a first electrical machine, and controlling rotation of a high pressure compressor using a second electrical machine, the combustion chamber downstream of the low pressure compressor and high pressure compressor; and controlling rotation of the low pressure compressor using the first electrical machine, and controlling rotation of the high pressure compressor using the second electrical machine to restart the start-up process or perform the re-light process.

Abstract: A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of: the ? ? lowest ? ? frequency ? ? of ? ? either ? ? mode ? ? Fan ? ? RTW ? ? or ? ? Rotor ? ? RW ? ? at the ? ? MTO ? ? speed the ? ? MTO ? ? speed may be in the range from 15 to 50%.

Abstract: An aircraft gas turbine engine includes a fan system having a reverse traveling wave first flap mode, Fan RTW, and including a fan upstream of the engine core; a fan shaft; and a front engine structure to support the shaft and having a front engine structure nodding mode FSN including two modes at similar, but unequal, natural frequencies in orthogonal directions; and a gearbox. The engine includes a gearbox, and a gearbox output shaft to couple output of the gearbox to the fan shaft. An LP rotor system including the fan system and the gearbox output shaft has a first reverse whirl rotor dynamic mode, Rotor RW.

Abstract: A rocket engine nozzle includes an aerospike having a plurality of adjustable airfoil vanes distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The adjustable airfoil vanes and an inner perimeter of the combustion chamber define a plurality of apertures which choke an exhaust exiting the combustion chamber and cause the exhaust to expand supersonically along the adjustable airfoil vanes, creating a supersonic jet. An actuator is configured to adjust a position of each of the adjustable airfoil vane relative to each other so as to direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically over the airfoil vanes without causing a shockwave to be imparted on the supersonic jet that is created. Accordingly, performance of the rocket engine is improved over conventional systems.

Abstract: A gas turbine engine includes a bypass duct, a compressor, and a pyrotechnic gas conduit. The bypass duct conducts bypass air between an outer bypass wall and an inner bypass wall. The compressor includes an impeller to compress engine air. The pyrotechnic gas conduit is configured to conduct a gas onto the impeller to start the gas turbine engine.

Abstract: A passive flow splitting system for use in a turbine engine control system to provide split fuel flow to two fuel manifolds to supply primary and secondary fuel injectors for the particular combustion zones thereof utilizing intentionally different split ratios dependent on ascending or descending combustion fuel flow is provided. The system includes a passive fuel divider valve (FDV) that includes a primary piston and a secondary piston. The primary piston is moveable independently from the secondary piston during a portion of its stroke, and is hydro-locked to the secondary piston during another portion of its stroke. An ecology valve is also provided to purge the fuel from the primary and/or secondary manifolds during different modes of operation. A transfer valve is included to control the position of ecology piston of the ecology valve.

Abstract: A bleed valve includes a housing with an inlet coupled to an outlet by a duct, a guide tube with an orifice fixed in the housing between the inlet and the outlet, a piston, and baffle. The piston is slideably supported on the guide tube and is movable between an open and a closed position, the duct fluidly coupling the inlet and outlet in the open position, the duct fluidly separating the inlet and outlet in the closed position. The orifice fluidly couples the inlet and outlet in the open and closed positions to move piston between the open and closed positions according to differential pressure between the bleed valve inlet and outlet. The baffle is slideably supported by the guide tube to set the differential pressure at which the piston moves between the open and closed positions. Gas turbines and differential pressure adjustment methods are also described.

Abstract: The present disclosure is directed to a gas turbine engine assembly having a compressor configured to increase pressure of incoming air, a combustion chamber, at least one turbine coupled to a generator, a torsional damper, and a controller. The combustion chamber is configured to receive a pressurized air stream from the compressor. Further, fuel is injected into the pressurized air in the combustion chamber and ignited so as to raise a temperature and energy level of the pressurized air. The turbine is operatively coupled to the combustion chamber so as to receive combustion products that flow from the combustion chamber. The generator is coupled to the turbine via a shaft. Thus, the torsional damper is configured to dampen torsional oscillations of the generator. Moreover, the controller is configured to provide additional damping control to the generator.