Abstract: A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

Abstract: A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A turbofan gas turbine engine includes a fan section having a fan, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture including an epicyclic gear train, a turbine section including a low pressure turbine and a high pressure turbine, the fan driven by the low pressure turbine through the geared architecture, and a power density between 4.84 lbf/in3 and 5.5 lbf/in3.

Abstract: A geared architecture for a gas turbine engine includes a central gear supported for rotation about the axis, a plurality of intermediate gears engaged with the central gear and a ring gear circumscribing the intermediate gears. A first flexible coupling is provided between an input shaft driven by a turbine section and the sun gear. The geared architecture provides a power density comprising a power measured in horsepower (HP) related to a weight of the geared architecture within a defined range that benefits overall engine weight and efficiency.

Abstract: A gas turbine engine according to the present disclosure includes, among other things, a propulsor section including a propulsor having a plurality of blades, the plurality of blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine that drives a first compressor and a second turbine that drives the propulsor section. A second inboard boundary of a core flowpath has a radius R1 defined at a first stage of a second compressor and has a radius R2 defined at a splitter rim that guides flow into the core flowpath.

Abstract: A turbine engine according to an example of the present disclosure includes, among other things, a fan shaft, at least one tapered bearing mounted on the fan shaft, the fan shaft including at least one passage extending in a direction having at least a radial component, and adjacent the at least one tapered bearing, a fan mounted for rotation on the at least one tapered bearing. An epicyclic gear train is coupled to drive the fan, the epicyclic gear train including a carrier supporting intermediate gears that mesh with a sun gear, and a ring gear surrounding and meshing with the intermediate gears, wherein the epicyclic gear train defines a gear reduction ratio of greater than or equal to 2.3. A turbine section is coupled to drive the fan through the epicyclic gear train, the turbine section having a fan drive turbine that includes a pressure ratio that is greater than 5. The fan includes a pressure ratio that is less than 1.45, and the fan has a bypass ratio of greater than ten (10).

Abstract: A propulsion system includes an electric fan propulsion motor with a plurality of propulsion motor windings. The propulsion system also includes a means for controlling a flow rate of a working fluid through a cryogenic working fluid flow control assembly to the propulsion motor windings. The propulsion system further includes a controller operable to control supplying a pre-cooling flow of the working fluid from a cryogenic liquid reservoir through the cryogenic working fluid flow control assembly to the propulsion motor windings.

Abstract: An example gas turbine engine includes a propulsor assembly consisting of a fan module and a fan drive turbine module, an epicyclic gear train, a high spool and a low spool. A weight of the propulsor assembly is less than 40% of a total weight of a gas turbine engine. The high spool includes an outer shaft, a high pressure turbine and a high pressure compressor. The low spool includes an inner shaft, a low pressure turbine and a low pressure compressor. The inner shaft drives the propulsor through the gear train to drive the propulsor. A weight of the propulsor is greater than a weight of the low pressure turbine.

Abstract: A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a high spool, and a low spool including a low pressure turbine that drives the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.

Abstract: A gas turbine engine according to the present disclosure includes, among other things, a fan section having a plurality of fan blades, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section, a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a low pressure turbine and a high pressure turbine. The low pressure turbine drives the low pressure compressor and the gear arrangement to drive the propulsor. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

Abstract: A bowed rotor start mitigation system for a gas turbine engine of an aircraft is provided. The bowed rotor start mitigation system includes a motoring system and a controller coupled to the motoring system and an aircraft communication bus. The controller is configured to determine at least one inferred engine operating thermal parameter based on at least one aircraft-based parameter received on the aircraft communication bus, where the at least one inferred engine operating thermal parameter is based on data describing a history of the aircraft before an engine shutdown. The motoring system is controlled to drive rotation of a starting spool of the gas turbine engine below an engine idle speed based on determining that the at least one inferred engine operating thermal parameter is within a preselected range.

Abstract: A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

Abstract: A gas turbine engine comprises a compressor section and a turbine section, the compressor section having a last compressor stage. High pressure cooling air is tapped from a location downstream of the last compressor stage and passed through a heat exchanger. Lower pressure air passes across the heat exchanger to cool the high pressure cooling air. A housing surrounds the compressor section and the turbine section and there being a space radially outwardly of the housing, and a mixing chamber received in the space radially outwardly of the housing, the mixing chamber receiving the high pressure cooling air downstream of the heat exchanger, and further receiving air at a temperature higher than a temperature of the high pressure cooling air downstream of the heat exchanger. Mixed air from the mixing chamber is returned into the housing and utilized to cool at least the turbine section.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

Abstract: An example gas turbine engine includes a propulsor assembly including at least a fan module and a fan drive turbine module; a gas generator assembly including at least a compressor section, a combustor in fluid communication with the compressor section, and a turbine in fluid communication with the combustor; and a geared architecture driven by the fan drive turbine module for rotating a fan of the fan module. A weight of the fan module and the fan drive turbine module is less than about 40% of a total weight of a gas turbine engine.

Abstract: A method is provided that includes providing a first turbine engine and providing a second turbine engine. The first turbine engine is configured with a first thrust rating. The first turbine engine includes a first engine rotating assembly and a first engine case structure housing at least the first engine rotating assembly. The second turbine engine is configured with a second thrust rating that is different than the first thrust rating. The second turbine engine includes a second engine rotating assembly and a second engine case structure housing at least the second engine rotating assembly. The first engine case structure and the second engine case structure have at least substantially common configurations. The first turbine engine and the second turbine engine are provided by a common entity.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and a fan duct annulus area outboard of the low pressure compressor section inlet, and a fan drive turbine section. The fan drive turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35, and is less than 0.55.