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 fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

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: 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 fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

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 fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

Abstract: A turbine assembly including a first rotor assembly with a rotatable outer drum from which one or more stages of a plurality of outer drum airfoils is extended radially inward is provided. An outer casing surrounds the outer drum of the first rotor assembly. A seal assembly is coupled to the outer casing and positioned radially outward from an upstream-most stage of the plurality of outer drum airfoils. The seal assembly is positioned in axial alignment with the upstream-most stage of the plurality of outer drum airfoils. The seal assembly separates a first plenum from a second plenum. The second plenum is formed axially aft of the first plenum and is formed by the seal assembly, the outer casing, and the outer drum of the first rotor assembly. The first plenum is positioned radially outward from the upstream-most stage of the plurality of outer drum airfoils.

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: A support assembly for a bearing of a gas turbine engine including a spring finger ring positioned radially exterior to an outer race of the bearing. The spring finger ring includes an outer ring positioned radially exterior to the outer ring, an inner ring positioned radially interior to the outer ring, and a plurality of spring fingers extending between the inner and outer rings. One or more spring fingers configured as two-sided spring fingers including a first ligament coupled to the outer ring and extending at a first circumferential angle to a first radial bumper proximate to the inner ring and a second ligament coupled to the inner ring and extending at a different second circumferential angle to a second radial bumper proximate to the outer ring. The first and second radial bumpers define first and second radial gaps between the bumpers and the inner and outer rings, respectively.

Abstract: A turbine engine including a core engine cowl including a compartment, and a cooling system positioned within the compartment. The cooling system includes a cooling fan configured to exhaust heat from the compartment, a temperature sensor configured to monitor a temperature within the compartment, and a controller coupled in communication with the cooling fan and the temperature sensor. The controller is configured to actuate the cooling fan when the temperature is greater than a threshold.

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: The present disclosure is directed to a method of control of a gas turbine engine comprising a fan section coupled to a low turbine together defining a low spool, an intermediate compressor coupled to an intermediate turbine together defining an intermediate spool, and a high compressor coupled to a high turbine together defining a high spool. The method includes providing an intermediate spool speed to low spool speed characteristic curve to a controller; providing a commanded power output to the controller; providing one or more of an environmental condition to the controller; determining, via the controller, a commanded fuel flow rate; determining, via the controller, a commanded intermediate compressor loading; and generating an actual power output of the engine, wherein the actual power output is one or more of an actual low spool speed, an actual intermediate spool speed, an actual high spool speed, and an actual engine pressure ratio.

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 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: The present disclosure is directed to a rotor thrust balanced turbine engine that may increase engine performance and efficiency while managing thrust mismatch or imbalance in a low pressure (LP) spool between a fan assembly and a turbine rotor assembly. The gas turbine engine defines 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 turbine rotor assembly and a turbine frame. The turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline. The first flowpath radius is disposed at the upstream end of the turbine rotor assembly, and wherein the second flowpath radius is disposed at the downstream end of the turbine rotor assembly.

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 fan section having a fan rotatable with a fan shaft and a turbomachinery section having a turbine and a turbomachine shaft rotatable with the turbine. A power gearbox is also provided mechanically coupled to both the fan shaft and the turbomachine shaft such that the fan shaft is rotatable by the turbomachine shaft across the power gearbox. A torque monitoring system includes a gearbox sensor operable with a gear of the power gearbox and a shaft sensor operable with at least one of the turbomachine shaft or the fan shaft. The torque monitoring system determines an angular position of the gear of the gearbox relative to at least one of the fan shaft or the turbomachine shaft using the gearbox sensor and the shaft sensor to determine a torque within the gas turbine engine.

Abstract: A support assembly for a bearing of a gas turbine engine including a spring finger ring positioned radially exterior to an outer race of the bearing. The spring finger ring includes an outer ring positioned radially exterior to the outer ring, an inner ring positioned radially interior to the outer ring, and a plurality of spring fingers extending between the inner and outer rings. One or more spring fingers configured as two-sided spring fingers including a first ligament coupled to the outer ring and extending at a first circumferential angle to a first radial bumper proximate to the inner ring and a second ligament coupled to the inner ring and extending at a different second circumferential angle to a second radial bumper proximate to the outer ring. The first and second radial bumpers define first and second radial gaps between the bumpers and the inner and outer rings, respectively.

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: A gas turbine engine includes a low pressure compressor and a low pressure turbine, with a low pressure shaft coupled to the low pressure turbine through a turbine extension and coupled to the low pressure compressor through a compressor extension. A forward bearing assembly supports the low pressure shaft within a compressor section and at a location aft of the compressor extension and an aft LP bearing assembly supports the low pressure shaft within a turbine section at a location aft of the turbine extension.

Abstract: A support assembly for a bearing of a gas turbine engine including a spring finger ring positioned radially exterior to an outer race of the bearing. The spring finger ring includes an outer ring positioned radially exterior to the outer ring, an inner ring positioned radially interior to the outer ring, and a plurality of spring fingers extending between the inner and outer rings. One or more spring fingers configured as two-sided spring fingers including a first ligament coupled to the outer ring and extending at a first circumferential angle to a first radial bumper proximate to the inner ring and a second ligament coupled to the inner ring and extending at a different second circumferential angle to a second radial bumper proximate to the outer ring. The first and second radial bumpers define first and second radial gaps between the bumpers and the inner and outer rings, respectively.

Abstract: The present disclosure is directed to a method of turbine section thermal management for a gas turbine engine. The engine includes a first turbine bearing defining an outer air bearing disposed radially adjacent to a low speed turbine rotor hub of a low speed turbine rotor and an inner bearing disposed radially adjacent to a high pressure (HP) shaft coupled to a high speed turbine rotor, wherein a first manifold is in fluid communication from a pressure plenum of a combustion section to the first turbine bearing, and wherein a second manifold is in fluid communication from the first turbine bearing to a pressure regulating valve and an outer diameter secondary flowpath of the turbine section, and wherein a third manifold is in fluid communication from the pressure plenum of the combustion section to the pressure regulating valve.