Abstract: A gas turbine engine includes a propulsor. A speed reduction device is drivingly connected to the propulsor. A compressor section with a high pressure compressor has between 8 and 13 stages and a pressure ratio of at least 16:1 and less than 35:1. A turbine section including a transition duct is located between a high pressure turbine and a low pressure turbine including fewer support struts than vanes in a first vane row of the low pressure turbine. The first vane row of the low pressure turbine is located downstream of the transition duct and downstream of a first row of blades in the low pressure turbine. The first row of blades in the low pressure turbine are immediately downstream of the support struts.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan, a compressor section, a turbine section including a fan drive turbine and a second turbine, an epicyclic gear system with a gear reduction, first and second bearings, and a fan drive shaft interconnecting the gear system and the fan. The fan drive turbine has a first exit area at a first exit point and is rotatable at a first speed. The second turbine has a second exit area at a second exit point and is rotatable at a second speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area.

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 turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

Abstract: A turbine engine includes 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 turbofan engine includes a fan section that drives air along a bypass flow path in a bypass duct. An epicyclic gear system in driving engagement with the fan shaft and has a gear mesh lateral stiffness and a gear mesh transverse stiffness. A gear system input to the gear system defines a gear system input lateral stiffness and a gear system input transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness. A first performance quantity is defined as the product of a first speed squared and a first area and a second performance quantity is defined as the product of a second speed squared and a second area. A performance quantity ratio of a first performance quantity to a second performance quantity is between 0.5 and 1.5.

Abstract: A gas turbine engine includes a plurality of fan blades rotatable about an axis, wherein each of the plurality of fan blades includes a leading edge. The gas turbine engine also includes turbine section includes an aft most turbine blade having a trailing edge and a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis. A center of gravity of the gas turbine engine is located a first axial distance from the trailing edge of the aft most turbine blade that is between about 35% and about 75% of a total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.

Abstract: An engine-driven cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for condensing a chilled air stream into liquid air for an aircraft use. The first air cycle machine includes a plurality of components operably coupled to a gearbox of a gas turbine engine and configured to produce a cooling air stream based on a first engine bleed source of the gas turbine engine. The second air cycle machine is operable to output the chilled air stream at a cryogenic temperature based on a second engine bleed source cooled by the cooling air stream of the first air cycle machine.

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 includes a propulsor section including a propulsor, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture, a turbine section including a low pressure turbine and a high pressure turbine, and a power density of greater than or equal to 4.75 and less than or equal to 5.5 lbf/in3, wherein the power density is a ratio of a thrust provided by the engine to a volume of the turbine 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: An aircraft propulsion system is disclosed and includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan, and an outer housing surrounding the fan to define a bypass flow path, a compressor, a turbine section including a fan drive turbine and a second turbine, an epicyclic gear system with a gear reduction, the fan drive turbine driving the fan through the gear system, and the gear system straddle-mounted by first and second bearings. The fan drive turbine has a first exit area at a first exit point and is rotatable at a first speed. The second turbine has a second exit area at a second exit point and is rotatable at a second speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area.

Abstract: A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section.

Abstract: A turbine engine includes 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 propulsor section includes a propulsor having a plurality of blades rotatable about an engine longitudinal axis. A compressor section includes a low pressure compressor and a high pressure compressor. A turbine section includes a low pressure turbine that drives the propulsor through an epicyclic gear arrangement, and includes a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor. The power ratio is defined by the second power input divided by the first power input. The power ratio is equal to, or greater than, 1.0 and less than, or equal to, 1.4.

Abstract: A gas turbine engine includes, among other things, a propulsor section including a propulsor hub, the hub including a hub diameter supporting a plurality of propulsor blades. A compressor section includes a first compressor and a second compressor. A turbine section includes a first turbine and a second turbine. A geared architecture interconnects the first turbine and the propulsor hub. The geared architecture includes a gear volume. A compressor inlet passage is disposed annularly about the geared architecture.

Abstract: A heat exchanger (HEX) for cooling air in a gas turbine engine is provided. An adjustable damper is provided. The adjustable damper may be for damping a movement of the HEX relative to the gas turbine engine. An adjustable damper may comprise: a first tube; a second tube located at least partially within the first tube; a housing coupled to the second tube; a moveable member, the moveable member comprising a contacting surface in contact with the second tube; an adjusting member adjustably coupled to the housing; and a spring member located between the moveable member and the adjusting member, the spring member configured to at least one of compress or decompress in response to adjusting member moving relative to the housing.

Abstract: A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

Abstract: A cryogenic cooling system for an aircraft includes a first air cycle machine, a second air cycle machine, and a means for collecting liquid air. The first air cycle machine is operable to output a cooling air stream based on a first air source. The second air cycle machine is operable to output a chilled air stream at a cryogenic temperature based on a second air source cooled by the cooling air stream of the first air cycle machine. An output of the second air cycle machine is provided to the means for collecting liquid air.