Abstract: A propulsion system for an aircraft includes a gas turbine engine and an engine power extraction system. The gas turbine engine includes a first rotational assembly. The first rotational assembly includes a propeller and a first shaft coupled to the propeller. The first shaft is configured for rotation about a rotational axis of the propulsion system. The engine power extraction system includes a fluid pump, a fluid motor, and a generator. The fluid pump is operably coupled to the first shaft. The fluid pump is in fluid communication with the fluid motor. The fluid pump is configured to pressurize and direct a hydraulic fluid to the fluid motor causing the fluid motor to rotatably drive the generator to generate an electrical power output.

Abstract: A gas turbine engine for an aircraft, comprising: a combustor; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least some fuel that has passed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the flow of fuel along the fuel return line such that a ratio of a temperature, in Kelvin, of fuel in the fuel tank to a temperature, in Kelvin, of fuel being delivered to the combustor is less than 0.56. Methods of operating a gas turbine engine are also disclosed.

Abstract: A propulsion system is provided for an aircraft. This aircraft propulsion system includes an engine housing, an open propulsor rotor, an engine core and a thrust diversion system. The open propulsor rotor is outside of the engine housing. The engine core is within the engine housing and motively coupled to the open propulsor rotor. The thrust diversion system is configured with the engine housing. The thrust diversion system includes a door configured to move radially from a stowed position to a deployed position.

Abstract: A combustor for a gas turbine includes a dome structure including a dome-side swirler opening therethrough, and a deflector connected to the dome structure. The deflector includes a deflector-side swirler opening therethrough and a deflector-dome connecting member arranged at the deflector-side swirler opening. The deflector-dome connecting member is connected with the dome-side swirler opening via a snap-fit type connection.

Abstract: A combustor for a gas turbine includes a combustor liner, an outer casing surrounding the combustor liner, an ignitor housing extending through an ignitor opening through the liner, and an ignitor disposed within the ignitor housing. The ignitor housing includes a housing wall including at least one airflow inlet passage on a downstream side of the housing wall and arranged within the outer flow passage, and at least one cooling passage on the downstream side within the housing wall. The at least one cooling passage extends along a length of the ignitor housing from the at least one airflow inlet passage and through the inner end of the ignitor housing, the at least one cooling passage being in fluid communication with the at least one airflow inlet passage.

Abstract: A dome-deflector assembly for a combustor of a gas turbine includes a dome portion having a dome-side swirler opening therethrough, and a deflector portion having a deflector-side swirler opening therethrough. The dome portion and the deflector portion are connected together to form a dome-deflector cavity therebetween. The deflector portion includes a first-side surface and a second-side surface, the second-side surface being arranged within the dome-deflector cavity, and the second-side surface including a plurality of stress-relief contours arranged in a pattern about the deflector-side swirler opening.

Abstract: Apparatuses and systems are provided herein for unducted propulsion systems. The system includes an aft housing for low drag for high subsonic sustained flight. A plurality of blades are affixed to the aft housing, wherein the housing defines a flowpath curve extending from the axial extent of the aft blade root to the aft end of the aft housing. The flowpath curve is described by an axial direction parallel to an axis of rotation and a radius from the axis of rotation. The flowpath curve includes first point having a first radius where the radius reaches a maximum aft of the aft blade root and a second point forward of the first point having a second radius where the radius stops decreasing. The ratio of the first radius to the second radius is greater than or equal to 1.081.

Abstract: A combustor wall includes a panel, a boss, a wall aperture and multiple cooling apertures. The panel extends axially along and circumferentially about an axial centerline. The panel extends radially between a first panel surface and a second panel surface. The first panel surface forms a peripheral boundary of a combustion chamber. The boss projects out from the second panel surface. The boss extends circumferentially around and forms an outer peripheral boundary of the wall aperture. The wall aperture extends along a wall aperture centerline through the combustor wall to the first panel surface. The cooling apertures are arranged circumferentially about the wall aperture. Each of the cooling apertures extends along a cooling aperture centerline through the panel and/or the boss to the wall aperture. The cooling aperture centerline of a first of the cooling apertures is angularly offset from the wall aperture centerline by a first acute angle.

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 combustor further includes a liner panel mounted to one of the inner shell and the outer shell aft of the bulkhead. The liner panel includes a first section including a first plurality of effusion holes. A first portion of the first plurality of effusion holes extends in a substantially circumferential direction. A second portion of the first plurality of effusion holes, disposed forward of the first portion, transitions from the substantially circumferential direction toward a substantially forward direction. A third portion of the first plurality of effusion holes, disposed aft of the first portion, transitions from the substantially circumferential direction to a substantially aft direction.

Abstract: A liner assembly for a combustor. The liner assembly includes a liner defining a combustion chamber of the combustor. The liner includes a looped section at a forward end of the liner. The looped section includes one or more bends such that the looped section forms a compliant joint and vibrations are dampened through the liner downstream of the looped section.

Abstract: An air turbine starter having a housing defining an interior with a primary inlet and a primary outlet to define a primary air flow path from the primary inlet to the primary outlet. A nozzle is located within the interior and has circumferentially spaced vanes. A shroud is also located within the interior and circumscribes at least a portion of the vanes. A retention mechanism constrains the axial movement of nozzle.

Abstract: A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. A combustor in the combustion section can include a combustor liner at least partially defining a combustion chamber. The combustor liner can include at least one aperture fluidly coupled to the combustion chamber.

Abstract: A system and method of providing prime reliability of a primary fuel manifold upon failure of a primary fuel metering system by providing metered fuel from a secondary fuel metering system are provided. The system utilizes a transfer electro-hydraulic servo valve (EHSV), a discharge select valve (DSV) coupled to the EHSV, and a pair of throttling valves positioned between secondary fuel metering system and the primary and secondary fuel manifolds. The DSV has a multi-land piston that switches a control pressure and a shutoff pressure between the pair of throttling valves to transfer the supply of metered fuel of the secondary fuel metering system from the secondary fuel manifold to the primary fuel manifold. The output pressurizing valve of the failed primary fuel metering system is also closed to isolate the failed primary fuel metering system from the primary fuel manifold.

Abstract: An exhaust mixer assembly for a gas turbine engine includes a core passage extending along a central engine axis for directing a core gas flow. An outer annular passage coaxially surrounds the core passage for directing a bypass gas flow. An exhaust mixer communicating with the core passage and outer annular passages has an upstream end, a downstream end and an annular wall extending therebetween. The annular wall defines a mixing plane at a trailing edge thereof at the downstream end at which the core and bypass gas flows are mixed. The annular wall has a radially inner surface and a radially outer surface. A plurality of protrusions are disposed on and extend laterally along and radially from at the radially inner surface and/or the radially outer surface. The plurality of protrusions are arranged one alongside another on the radially inner surface and/or the radially outer surface.

Abstract: Electric engine to provide thrust to fly an aircraft. Engine includes housing, air inlet, shaft, bladed rotor having a plurality of magnets secured on shaft, stator having plurality of coils positioned so as to interact with plurality of magnets and an exhaust nozzle. Powering coils causes interaction with magnets that results in bladed rotor rotating and pressurizing and accelerating air received via air inlet and expelling via exhaust nozzle to provide thrust. Engine may include generator having stator with coils secured to shaft via bearings and plurality of rotors with magnets secured to shaft to rotate with shaft. Magnets rotating past coils results in electric generation. Second hollow shaft may be mounted to shaft with bearings and generator may be located with hollow shaft. Second bladed rotor may be connected to, and rotate, second shaft. Engine may include ducts external to bladed rotor and fan to route air therein.

Abstract: An embodiment of a torch igniter for a combustor of a gas turbine engine includes a combustion chamber oriented about an axis, a cap defining an axially upstream end of the combustion chamber, a tip defining the axially downstream end of the combustion chamber, an 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, an outlet passage defined by the igniter wall within the tip, a glow plug housing configured to receive a glow plug and allow an innermost end of the glow plug to extend into the combustion chamber, and a cooling system. The cooling system includes an air inlet formed within an exterior of the structural wall, a cooling channel forming a flow path through the structural wall at the glow plug housing, and an air passage.

Abstract: A combustor section support structure can include a fuel manifold and one or more radial stud assemblies connected to the fuel manifold and configured to allow the fuel manifold to move radially relative to an engine case and to fix the fuel manifold axially relative to the engine case. Each of the one or more radial stud assemblies can include a first portion mounted to the fuel manifold and a second portion configured to mount to an engine case. The first portion and the second portion can be configured to slide relative to each other to allow relative radial movement, but to prevent axial relative movement. A system can include a fuel manifold configured to distribute fuel to one or more fuel nozzles of a turbomachine, and a fuel inlet assembly configured to connect to the fuel manifold through an engine case. The fuel inlet assembly can be configured to axially retain the fuel manifold relative to the engine case and to allow radial movement of the fuel manifold relative to the engine case.

Abstract: A gas turbine engine includes a compressor, a combustor, and a turbine. The compressor compresses gases entering the gas turbine engine. The combustor receives the compressed gases from the compressor and mixes fuel with the compressed gases. The turbine receives the hot, high pressure combustion products created by the combustor by igniting the fuel mixed with the compressed gases. The turbine extracts mechanical work from the hot, high pressure combustion products to drive the compressor and a fan, shaft, or propeller.

Abstract: Methods and apparatus for gas turbine frame flow path hardware cooling are disclosed. An example engine fan case includes an outer band and an inner band, the outer band and the inner band connected using a double-walled vane, the vane including openings to pass cooling air flow from the outer band to an airfoil of the fairing, and an end segment seal, the seal formed on an edge of the fairing using an auxetic material.

Abstract: A method includes forming a ceramic matrix composite component by infiltrating an array of ceramic-based fibers with a ceramic-based matrix; forming a plurality of cooling holes in the ceramic matrix composite component; applying a slurry of particles in a carrier fluid to the ceramic matrix composite component such that the slurry passes through the cooling holes and wicks into the ceramic matrix composite material; and processing the ceramic matrix composite component to remove the carrier fluid, thereby leaving a filler at a wall surface of the plurality of cooling holes. A component is also disclosed.