Abstract: An exhaust assembly for a turboprop engine, the assembly having: a primary exhaust nozzle disposed at an aft end of an engine core of the turboprop engine; a secondary exhaust nozzle extending aft from the primary exhaust nozzle to an aft end of the secondary exhaust nozzle, a movable panel disposed in the aft end of the secondary exhaust nozzle and movable to decrease an area of the secondary exhaust nozzle.

Abstract: Systems and methods for operating systems are provided. For example, a system comprises a heat source for providing a flow of a hot fluid and a fuel flowpath for a flow of a fuel. The fuel flowpath includes a fuel accumulator and a heat exchanger for heat transfer between the hot fluid and fuel. The heat exchanger includes a hot fluid inlet for receipt of the hot fluid at an inlet temperature and a fuel inlet for receipt of the fuel at an inlet temperature. The hot fluid inlet temperature is greater than the fuel inlet temperature such that the fuel is heated through heat transfer with the hot fluid in the heat exchanger. The fuel accumulator accumulates at least a portion of the heated fuel. An exemplary system is selectively operated to heat and circulate the fuel through the fuel flowpath for consumption and/or accumulation in the fuel accumulator.

Abstract: The gas propulsion thrust device comprises a cone-shaped propulsion element with a rigid concave internal surface and a second convex external surface. The propulsion element is submerged in gas and aligned with a high-frequency linear actuator, which causes its reciprocal motion along the longitudinal axis, generating thrust. The device includes a thrust chamber that supports the actuator and surrounds the propulsion element, maintaining a consistent gap between the propulsion element and the chamber to direct gas from the second side to the first side. The chamber tapers around the propulsion element, and a gas directing cap is configured to direct gas around the second side towards the first side. Propulsion element is rigid and its shape remains unchanged during the operation. The reciprocal motion creates a gas pressure differential across the propulsion element, generating thrust by propelling gas away from the propulsion element in the opposite direction of propulsion.

Abstract: A propulsion system for an aircraft including a core engine connected to a forward frame and a turbine frame assembly, and an aft frame assembly is connected by a turbine casing to the core engine aft of the turbine frame assembly. A core cowl surrounds the core engine, wherein the core cowl is connected to the forward frame. A plurality of cowl mount links selectively loads the core cowl to the aft frame assembly and the turbine frame assembly, wherein the plurality of cowl mount links is each loaded by deflection of the core engine.

Abstract: A method for operating a hybrid-electric propulsion system of an aircraft is provided. The hybrid-electric propulsion system includes a gas turbine engine having a high pressure system, a low pressure system, and an electric machine coupled to one of the high pressure system or low pressure system. The method includes receiving data indicative of an actual or anticipated in-flight shutdown of the gas turbine engine; and adding power to the gas turbine engine through the electric machine in response to receiving data indicative of the actual or anticipated in-flight shutdown of the gas turbine engine.

Abstract: A power system includes a gas turbine engine including, in serial flow relationship defining a flow passage, a compressor section, a combustion section, a turbine section, and an exhaust section, the gas turbine engine generating exhaust gas that includes a baseline amount of carbon dioxide by volume within the exhaust gas. A carbon capture unit is arranged downstream of the gas turbine engine and processes the exhaust gas output by the gas turbine engine to capture the carbon dioxide contained in the output exhaust gas. A carbon dioxide reinjection system is arranged to reinject at least a portion the carbon dioxide captured from the output exhaust gas by the carbon capture unit into the flow passage of the gas turbine engine.

Abstract: A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.

Abstract: A device for guiding a main air flow (F1) for an aircraft turbine engine, the device including a first air flowing pipe of a main air flow, the first pipe having a main axis, a plurality of ejectors of a secondary air flow located within the first pipe and configured to eject a secondary air flow and force the flow of the main air flow into the first pipe, the ejectors being distributed around the main axis, and a second air flowing pipe located at the outlet of the ejectors and including one end which is connected to one end of the first pipe, wherein the second pipe includes a narrow end.

Abstract: A turbofan engine includes a first spool including a first turbine, and a first tower shaft engaged to the first spool. A second spool includes a second turbine, and a second tower shaft is engaged to the second spool. A superposition gearbox includes a sun gear, a plurality of intermediate gears engaged to the sun gear, and is supported in a carrier and a ring gear circumscribing the intermediate gears. The first tower shaft or the second tower shaft drives one of the intermediate gears. A drive motor is engaged to drive the sun gear, an inner electric motor, a stator disposed radially outside of the inner electric motor, and an outer electric motor disposed radially outside the stator. A first load on the first spool and a second load on the second spool is adjusted by operation of at least one of the inner electric motor and the outer electric motor.

Abstract: An engine support assembly for operatively connecting an engine to an aircraft, the aircraft, the aircraft having a support frame surrounded by opposing first and second skins. The engine support assembly broadly comprises first and second support structures. The first support structure is secured to the support frame and extends away from the first skin. The second support structure is releasably connected to the first support structure and releasably connected to the engine.

Abstract: A thrust reverser including a sealing membrane designed to deflect at least one portion of a secondary flow towards a deflection cascade when a movable structure of the reverser is in the retracted reverse-thrust position, the reverser including a rear frame for supporting the deflection cascade. The reverser further includes an interface device between the sealing membrane and a rear cascade structure including the rear cascade support frame, the vanes located furthest back on the deflection cascade, and a bearing member via which the cascade bears against the rear cascade support frame. Moreover, the device forms an elastic clip mounted around the rear structure, and a rear clip portion whereof is arranged behind the structure.

Abstract: An aircraft engine, comprising a thermal engine, an axial turbine having a turbine inlet, and an exhaust assembly fluidly connecting the thermal engine to the axial turbine. The exhaust assembly includes a housing and a deflector removably mounted within the housing. The deflector has circumferentially distributed vanes. The deflector is a first deflector, having a first set of geometric characteristics, that is removable from the housing and replaceable by a second deflector having a second set of geometric characteristics different from the first set of geometric characteristics. The first deflector is one of a first class of deflectors and the second deflector is one of a second class of deflectors, the first and second class of deflectors respectively defining first and second exhaust flow profiles that differ from each other.

Abstract: In one embodiment, systems and methods include using an inspection system to remove or seal foreign object debris. An inspection system comprises a borescope; wherein a plurality of optical fibers is disposed within the borescope; a mounting system; an articulation system; and a controller, wherein the controller comprises a display; wherein a proximal end of the borescope is coupled to the mounting system, wherein the mounting system is configured to secure the inspection system to an external surface, wherein the mounting system is coupled to the articulation system, wherein the articulation system is configured to actuate the borescope, wherein the articulation system is communicatively coupled to the controller.

Abstract: A powerplant is provided that includes a first gas turbine engine, a second gas turbine engine, a second engine bypass flowpath and a flow control system. The first gas turbine engine includes a first core flowpath fluidly coupled with a first inlet and a first exhaust. The first core flowpath extends sequentially through a first compressor section, a first combustor section and a first turbine section. The second gas turbine engine a second core flowpath fluidly coupled with a second inlet and a second exhaust. The second core flowpath extends sequentially through a second compressor section, a second combustor section and a second turbine section. The flow control system fluidly couples the first inlet and the first exhaust to the second core flowpath during a first mode. The flow control system fluidly couples the first inlet and the first exhaust to the second engine bypass flowpath during a second mode.

Abstract: A turbine stationary blade is a turbine stationary blade that includes: a blade body; and a shroud formed in an end portion of the blade body in a blade height direction. The shroud includes: a bottom plate in contact with a combustion gas flow passage; a peripheral wall formed in the blade height direction and along a peripheral edge of the bottom plate; a recess forming a space surrounded by the peripheral wall and the bottom plate; a plurality of partition ribs connecting the blade body and the peripheral wall, dividing the recess into a plurality of spaces, and forming a plurality of cavities; and an impingement plate dividing the space into an outer cavity formed on an outer side in the blade height direction and an inner cavity formed on an inner side of the outer cavity, and having a plurality of through holes via which the outer cavity and the inner cavity communicate with each other.

Abstract: An aircraft propulsion system includes a series of condensers where water is condensed from the exhaust gas flow. Each of the series of condensers define a passage for exhaust gas flow such that water recovered from the exhaust gas flow drains at a downward angle in a direction of gravity through each of the series of condensers. A series of water separators separate the exhaust gas flow from water condensed in a corresponding one of the series of condensers and the water is transformed into a steam flow in an evaporator system for communication to the combustor.

Abstract: An assembly for an aircraft includes an aircraft housing, an auxiliary power unit (APU), an oil system, a door actuator, and a controller. The aircraft housing forms a compartment. The aircraft housing includes an intake door. The intake door is movable between a closed position, an open position, and intermediate positions between the closed position and the open position. The APU is disposed within the compartment. The APU includes an engine. The engine includes an air inlet and a rotational assembly. The oil pump is operatively connected to the rotational assembly. The door actuator is operatively connected to the intake door. The door actuator is operable to position the intake door in the closed position, the open position, and the intermediate positions to control an ambient air flow to the air inlet. The controller is operatively connected to the door actuator.

Abstract: A powerplant is provided for an aircraft. This aircraft powerplant includes a turbine engine core, a steam system and a bypass system. The turbine engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The flowpath extends through the compressor section, the combustor section and the turbine section from an inlet into the flowpath to an exhaust from the flowpath. The steam system includes an evaporator disposed along the flowpath downstream of the turbine section. The steam system is configured to evaporate water into steam using the evaporator. The steam system is configured to introduce the steam into the flowpath upstream of the turbine section. The bypass system is configured to bleed fluid from the flowpath upstream of the turbine section to provide bleed fluid. The bypass system is configured to direct the bleed fluid into the flowpath downstream of the evaporator.

Abstract: A powerplant for an aircraft, comprising a frame, a propulsion system comprising a core enclosed in a casing and comprising a combustion chamber and a turbine, a supply pipe for conveying dihydrogen to the combustion chamber and that snakes outside the casing running along the turbine before dropping down into the combustion chamber through the casing, and a protective plate fastened to the frame and positioned between the casing and the supply pipe. With such an arrangement, if a blade of the turbine became detached it would meet the protective plate blocking its path to the dihydrogen pipe and would thus be diverted or stopped.

Abstract: A fuel injector for a gas turbine engine combustor is provided that includes a swirler, a mounting stage, and a distributor. The swirler has a shaft, a collar, a throat section, and first and second axial ends. The throat section includes an inner radial surface that defines a central passage that extends between the swirler inner bore and the collar. The collar includes a plurality of apertures extending therethrough disposed radially outside of the central passage. The mounting stage is disposed in the inner bore, and has an annular flange, a central hub, and at least one strut. The distributor has a stem attached to a head. The stem has a distal end opposite the head portion engaged with the central hub. The head portion has an end surface and a side surface. The distributor is selectively positionable relative to the throat section.