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 includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

Abstract: A turbofan gas turbine engine includes, among other things, a fan section including a fan hub and an outer housing, the fan hub including a hub diameter supporting a plurality of fan blades, a turbine section including a fan drive turbine, and a geared architecture that interconnects the fan drive turbine and the fan hub, the geared architecture including a gear volume.

Abstract: A heat exchanger (HEX) for cooling air in a gas turbine engine is provided. The HEX may comprise a central manifold comprising an inlet portion, a first outlet portion, and a second outlet portion. The HEX may further comprise a plurality of tubes coupled to the central manifold, the plurality of tubes comprising at least a first tube, a second tube, a third tube, and a fourth tube, a shroud at least partially encasing said plurality of tubes, and a cooling air flow path defined by at least one of the shroud, the plurality of tubes, and an outer surface of the central manifold, wherein the cooling air flow path is orthogonal to said plurality of tubes.

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 high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure 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: 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 compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas pat

Abstract: An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes an engine bleed tap coupled to a fan-air source or a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine and a motorized compressor in fluid communication with the engine bleed tap. The engine bleed control system also includes a controller operable to selectively drive the motorized compressor to boost a bleed air pressure as pressure augmented bleed air and control delivery of the pressure augmented bleed air to an aircraft use.

Abstract: A seal assembly for a gas turbine engine includes first and second split-ring seals. The first split-ring seal includes circumferentially separated first and second ends, laterally separated first and second edges, a first sealing lobe adjacent the first edge, and a first integral clip portion adjacent the second edge. The first sealing lobe has a curved surface extending laterally outward from the first edge. The second split-ring seal is slidably received and nested in the first seal. The second split-ring seal includes circumferentially separated third and fourth ends and laterally separated third and fourth edges.

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 gas turbine engine includes a fan bypass ratio greater than 12. A speed reduction device includes a gear ratio of at least 2.6. A turbine section includes a transition duct that is located between a high pressure turbine and a low pressure turbine and includes fewer support struts than vanes in a first vane row of the low pressure turbine.

Abstract: An improved stator assembly is disclosed. The stator assembly may comprise an exit guide vane, an OD ring, and an ID ring. The exit guide vane may couple at one end to the OD ring and at an opposite end to the ID ring. The exit guide vane may comprise a leading edge opposite of a trailing edge. The OD ring and the ID ring may couple to a diffuser assembly of a gas turbine engine. The stator assembly may further comprise an aft OD seal, a forward OD seal, an ID seal, and a diffuser assembly seal to reduce airflow leaks around the stator assembly.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

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 heat exchanger array includes a first row of heat exchangers, a second row of heat exchangers, and side curtains. The first row heat exchangers are spaced apart to define first gaps. The second row heat exchangers are spaced apart to define second gaps and are positioned downstream of and staggered from the first row heat exchangers such that the second row heat exchangers are aligned with the first gaps and the first row heat exchangers are aligned with the second gaps. Each side curtain is in close proximity to a first row heat exchanger and a second row heat exchanger. The side curtains define a neck region upstream of and aligned with each first row heat exchanger and each second row heat exchanger. Each neck region has a neck area that is less than a frontal area of the heat exchanger with which it is aligned.

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 ground-based cryogenic cooling system includes a means for cooling an airflow and producing chilled air responsive to a power supply. A liquid air condensate pump system is operable to condense the chilled air into liquid air and urge the liquid air through a feeder line. A cryogenic cartridge includes a coupling interface configured to detachably establish fluid communication with the feeder line and a cryogenic liquid reservoir configured to store the liquid air under pressure. The cryogenic cartridge can be coupled to a cryogenic liquid distribution system on an aircraft. The liquid air can be selectively released from the cryogenic cartridge through the cryogenic liquid distribution system for an aircraft use.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor supported on a propulsor shaft, a turbine section including a turbine shaft, a compressor section, and an epicyclic gear train interconnecting the propulsor shaft and the turbine shaft. The gear train includes a sun gear coupled to the turbine shaft, intermediary gears arranged circumferentially about and meshing with the sun gear, a carrier supporting the intermediary gears, and a ring gear including first and second portions each having an inner periphery with teeth intermeshing with the intermediate gears. The first and second portions have axially opposed faces abutting one another at a radial interface. The first and second portions have grooves at the radial interface that form a hole that expels oil through the ring gear.

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

Abstract: A turbofan engine may comprise an inlet and a fan case coupled to the inlet. An engine case may be coupled to the fan case via a vane extending between the fan case and the engine case. A strut apparatus may extend from the fan case and limit deflection of the fan case. The strut apparatus may comprise a first end proximate the fan case, and a second end coupled to at least one of the engine case or a structure for mounting the turbofan engine to an aircraft.

Abstract: A gas turbine engine includes a bypass ratio greater than about ten (10). A fan is supported on a fan shaft and has a plurality of fan blades. A gear system is connected to the fan shaft. A plurality of planetary gears and a first set of opposed angled ring gear teeth are separated from a second set of opposed angled ring gear teeth. A lubricant flow path is located axially between the first set of opposed angled ring gear teeth and the second set of opposed angled ring gear teeth. A gear system support is relative to a fixed housing facilitating segregation of vibrations. An annular channel is axially aligned with the lubricant flow path. A low pressure turbine has an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1.