Abstract: A noise reducing airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a pressure side, a suction side, and a trailing edge sheath including an outer surface coupled to the trailing edge of the airfoil. The trailing edge sheath extends at least partially along the chord on the pressure and suction sides at each point along the span within the trailing edge sheath. The trailing edge sheath defines a fluid passageway extending along at least a portion of the span. Further, the trailing edge sheath defines at least one aperture on at least one of the pressure side, the suction side, or trailing edge fluidly coupling the fluid passageway to the outer surface.

Abstract: A turbofan engine assembly includes a core gas turbine engine with a booster compressor having a radially outer diameter, a fan case assembly, and an outlet guide vane assembly. The outlet guide vane assembly includes a plurality of outlet guide vane segments spaced circumferentially about the core gas turbine engine. Each outlet guide vane segment of the plurality of outlet guide vane segments includes a radially inner fan hub frame flange configured to couple to the core gas turbine engine, a radially outer fan case flange configured to couple to the fan case assembly, and a plurality of outlet guide vanes extending therebetween. The radially inner diameter of the outlet guide vane assembly is smaller than the radially outer diameter of the booster compressor.

Abstract: A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.

Abstract: A gas turbine engine includes a frame member extending generally along an axial direction and defining a plurality of slots spaced along a circumferential direction; and a gearbox including a ring gear, a plurality of planet gears, and a sun gear, the plurality of planet gears each positioned at least partially in a respective slot of the plurality of slots defined by the frame member such that the frame member extends through the gearbox.

Abstract: An engine includes a static frame member; a spool configured to rotate relative to the static frame member during operation of the engine; a gearbox coupled to the spool; and a gearbox coupling assembly mounting the gearbox to the static frame member using a plurality of moveable connection members, each moveable connection member including a sliding connection allowing for movement of the gearbox relative to the static frame member in a plane perpendicular to an axial direction of the engine.

Abstract: A method includes installing a first stage assembly including a first ring member and a first stage of rotor blades, the first ring member defining a first end and the first stage of rotor blades defining a second end; installing a second stage assembly including a second ring member and a second stage of rotor blades, the second ring member defining a first end and the second stage of rotor blades defining a second end, wherein installing the second stage assembly includes fitting the first end of the second ring member to the second end of the first stage of rotor blades to form a first attachment interface; and pressing the second stage assembly against the first stage assembly to fix the first attachment interface.

Abstract: A gas turbine engine includes a compressor and a turbine; a plurality of rotor blades positioned in the compressor or the turbine and arranged in a plurality of stages along an axial direction; an outer drum positioned at least partially radially outward of the plurality of rotor blades and rotatable with the plurality of rotor blades; and a casing extending over the outer drum and defining a plurality of air cavities with the outer drum, the plurality of air cavities arranged in series airflow communication along the axial direction.

Abstract: A gas turbine engine includes a turbine having a stage of turbine rotor blades. The gas turbine engine additionally includes a combustion section defining a cooling air passage for providing a cooling airflow to the stage of turbine rotor blades. The gas turbine engine additionally includes a fluidic nozzle assembly having a fluidic nozzle positioned in or immediately upstream of the cooling air passage of the combustion section. The fluidic nozzle defines an opening and is operable to induce cooling airflow through the opening to increase or decrease an amount of cooling air provided to the stage of turbine rotor blades to, e.g., increase an efficiency of the gas turbine engine.

Abstract: In one embodiment, a compressor clearance control system is provided. The compressor clearance control system includes a cooling air source, a compressor casing, and a cooling air flowpath defined through the compressor casing. The flowpath includes an inlet configured to receive cooling air from the cooling air source and an outlet configured to exhaust the cooling air to the cooling air source such that the flowpath defines a closed loop within the compressor.

Abstract: A system for mounting an engine to an aircraft includes an engine forward mount angled toward the forward end of the engine at a first angle. At least two thrust links extend between an engine aft mount to a link support connection at a second angle. The engine aft mount is angled toward the forward end of the engine at a third angle. A projection of a load vector of the engine forward mount onto a vertical plane extending through the axis of rotation of the engine and a projection of a load vector of each of the at least two thrust links onto the vertical plane intersect the axis of rotation of the engine within a first vertical plane segment extending between a forward end of a nose of a fan assembly and forward of a forward mount interface.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor.

Abstract: A gas turbine engine defines a radial direction and an axial centerline. The gas turbine engine includes a core turbine engine that defines a core inlet. The core inlet is oriented with respect to the axial centerline and positioned along the radial direction such that the area available to capture foreign object debris is minimized. In one aspect, the gas turbine engine defines a capture ratio less than about 35%, wherein the capture ratio is a ratio of an area between a splitter radius and a tangency radius to an area encompassed by the splitter radius. The splitter radius is defined as a radial distance between the axial centerline and an outer lip of a splitter of the core turbine engine. The tangency radius is defined as a radial distance between the axial centerline and a tangency point, which can be defined at an inner lip of the core inlet.

Abstract: A gas turbine engine includes a fan section having a single-stage fan, a core turbine engine, and a nacelle at least partially surrounding the fan of the fan section and the core turbine engine. The gas turbine engine also includes an outlet guide vane extending between a casing of the core turbine engine and the nacelle. The gas turbine engine is configured to define an acoustic ratio greater than or equal to 2.3, with the acoustic ratio being a ratio of an axial spacing between the trailing edge of the fan blade and the leading edge of the outlet guide vane at a radial location seventy-five percent (75%) along the span of the fan blade to the axial width of the fan blade also at the radial location seventy-five percent (75%) along the span of the fan blade.

Abstract: A gas turbine engine includes a compressor section having a high pressure compressor and a core casing surrounding the compressor section and defining an inlet. The gas turbine engine also includes a cooling system for cooling air in or to the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and a fluid line assembly in fluid communication with the fluid tank. The fluid line assembly includes an outlet positioned upstream of the high pressure compressor and downstream of the inlet defined by the core casing for injecting cooling fluid into an airflow upstream of the high pressure compressor.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, an axial centerline extended along the longitudinal direction, an upstream end and a downstream end opposite of the upstream end along the longitudinal direction, a radial direction, and a circumferential direction. The gas turbine engine includes a high speed turbine rotor coupled to a high pressure (HP) shaft and HP compressor, a low speed turbine rotor comprising an axially extended hub, and a first turbine bearing disposed radially between the low speed turbine rotor and the high speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit through the high speed turbine rotor. The low speed turbine rotor includes a rotating nozzle adjacent to the turbine cooling conduit. The first turbine bearing defines an outer air bearing and an inner air bearing. The first turbine bearing defines a stationary nozzle adjacent to the rotating nozzle of the first turbine rotor.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a low pressure (LP) turbine defining an outer flowpath. The outer flowpath defines a first outer flowpath radius at an upstream-most end of the LP turbine, a last outer flowpath radius disposed at a downstream-most end of the LP turbine, a middle outer flowpath radius disposed therebetween along the longitudinal direction. The middle outer flowpath radius is greater than the last outer flowpath radius.

Abstract: The present disclosure is directed to a method of operating a gas turbine engine with an interdigitated turbine section. The engine includes a fan rotor, an intermediate pressure compressor, a high pressure compressor, a combustion section, and a turbine section in serial flow arrangement. The turbine section includes, in serial flow arrangement, a first stage of a low speed turbine rotor, a high speed turbine rotor, a second stage of the low speed turbine rotor, an intermediate speed turbine rotor, and one or more additional stages of the low speed turbine rotor. The low speed turbine rotor is coupled to the fan rotor via a low pressure shaft. The intermediate speed turbine rotor is coupled to the intermediate pressure compressor via an intermediate pressure shaft. The high speed turbine rotor is coupled to the high pressure compressor via a high pressure shaft.

Abstract: A noise reducing airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a pressure side, a suction side, and a trailing edge sheath including an outer surface coupled to the trailing edge of the airfoil. The trailing edge sheath extends at least partially along the chord on the pressure and suction sides at each point along the span within the trailing edge sheath. The trailing edge sheath defines a fluid passageway extending along at least a portion of the span. Further, the trailing edge sheath defines at least one aperture on at least one of the pressure side, the suction side, or trailing edge fluidly coupling the fluid passageway to the outer surface.

Abstract: In one aspect, the present disclosure is directed to a cooling circuit for a gas turbine engine. The cooling circuit includes a rotor blade having a connection portion and a rotor disc having a first axial side and a second axial side. The rotor disc defines a connection slot and a cooling passage extending between the first axial side and the second axial side. The connection slot receives the connection portion to couple the rotor blade to the rotor disc. Cooling air flows through the cooling passage.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor including a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor including a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; and a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit extended within the high speed turbine rotor.