Abstract: A gas turbine engine is provided that includes compressor and combustor sections, inner and outer casings, an annular diffuser, an inner diffuser casing, a heat exchanger, and an HPT stator vane stage. An annular combustor is disposed radially inward of the outer casing and has inner and outer radial wall structures. The outer casing and the combustor outer radial wall structure define a diffuser OD flow path. The annular diffuser directs diffuser gas towards the combustor section. The inner diffuser casing is disposed radially inward of the annular combustor and spaced apart from the combustor inner radial wall structure. The inner casing is disposed radially inward of and spaced apart from the inner diffuser casing. The inner diffuser casing and the inner casing define an ICF passage. The heat exchanger is configured to produce intercooler gas. Intercooler gas is directed through the ICF passage and into the HPT stator vanes.

Abstract: A gas turbine engine is provided having an axial centerline, a compressor section, a turbine section, and a combustor section. The turbine section has a turbine first vane assembly that includes an annular first vane inner radial support. The combustor section has an outer casing, an annular combustor, and a unitary inner diffuser structure. The annular combustor has an inner radial flange. The unitary inner diffuser structure includes a compressor discharge, an inner diffuser case, and a tangential onboard injector (TOBI) inseparably attached to one another. The unitary inner diffuser structure further includes an outer radial flange and a TOBI connection flange. The annular FV inner radial support, the combustor inner radial flange, and the TOBI connection flange are secured to one another.

Abstract: A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

Abstract: A gas turbine engine includes a compressor section having a downstream most location and more upstream locations. A compressor housing surrounds the compressor rotors. A combustor is downstream of the compressor and a turbine section is downstream of the combustor. At least one bleed inlet passage bleeds air compressed from the compressor section, and delivers the air for use on an associated aircraft. A baffle plate is intermediate the at least one bleed inlet passage and a bleed exit port from the gas turbine engine. The baffle plate has a plurality of holes for allowing the air to flow from the bleed inlet passage to the bleed exit port, with an area of the holes increasing from a location adjacent the at least one bleed inlet passage and in a circumferential direction away from the at least one bleed inlet passage. An aircraft air supply system and an aircraft are also disclosed.

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: A bleed air cooling system for a gas turbine engine includes one or more bleed ports located at one or more axial locations of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. One or more safety sensors are configured to sense operational characteristics of the gas turbine engine, and a controller is operably connected to the one or more safety sensors and configured to evaluate the sensed operational characteristics for anomalies in operation of the bleed air cooling system.

Abstract: A turbofan engine includes a propulsor section that has a propulsor shaft in driving engagement with a propulsor. An epicyclic gear system has a gear mesh lateral stiffness and a gear mesh transverse stiffness. A gear system input 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 turbine section rotates at a first speed, and a second turbine rotates at a second speed that is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area of the first turbine, a second performance quantity is defined as the product of the second speed squared and the second area of the second turbine, and a performance quantity ratio is between 0.5 and 1.5.

Abstract: A gas turbine engine includes a compressor section including a first compressor, a turbine section including a first turbine and a second turbine, a first shaft and a second shaft, the first shaft interconnecting the first turbine and the second compressor, and a geared architecture. The first shaft is supported on a first bearing in an overhung manner. A performance ratio is between 0.5 and 1.5.

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: 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 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: An environmental control system for an aircraft includes a higher pressure tap associated with a higher compression location in a main compressor section. The higher pressure tap leads into a turbine section of a turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive a compressor section of the turbocompressor. A combined outlet receives airflow from a turbine outlet and a compressor outlet intermixing airflow and passing the mixed airflow downstream to be delivered to an aircraft system. A buffer air outlet communicates airflow to an engine buffer air system.

Abstract: A gas turbine engine includes, among other things, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, an inner shaft that interconnects the low pressure compressor and the low pressure turbine, an outer shaft that interconnects the high pressure compressor and the high pressure turbine, a bearing structure that supports at least one of the inner and outer shafts, and a buffer system. The buffer system prepares buffer air, communicates the buffer air to the bearing structure to pressurize a bearing compartment, and then from the bearing structure along at least one of the shafts.

Abstract: An environmental control system for an aircraft includes a higher pressure tap associated with a higher compression location in a main compressor section. The higher pressure tap leads into a turbine section of a turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive a compressor section of the turbocompressor. A combined outlet receives airflow from a turbine outlet and a compressor outlet intermixing airflow and passing the mixed airflow downstream to be delivered to an aircraft system. A buffer air outlet communicates airflow to an engine buffer air system.

Abstract: A bleed air cooling system for a gas turbine engine includes one or more bleed flowpaths operably connected to a bleed outlet to divert a bleed airflow from a gas turbine engine flowpath. Each bleed flowpath includes two or more bleed ports to divert a bleed airflow from a gas turbine engine flowpath, and a bleed duct in fluid communication with the bleed ports and configured to convey the bleed airflow from the two or more bleed ports to the bleed outlet. A valve is located at each bleed port of and is configured to move between an opened position and a closed position, and one or more sensors are located along the bleed flowpath to sense one or more conditions of the bleed air cooling system. The valve at a particular bleed port is moved to the opened position based on the sensed one or more conditions.

Abstract: A gas turbine engine includes a buffer system that communicates a buffer cooling air to at least one bearing structure and at least one shaft of the gas turbine engine. The buffer system includes a first bleed air supply and a conditioning device that conditions the first bleed air supply to render the first buffer supply air at an acceptable temperature to pressurize the at least one bearing structure and cool the at least one shaft.

Abstract: A gas turbine engine includes, among other things, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, an inner shaft that interconnects the low pressure compressor and the low pressure turbine, an outer shaft that interconnects the high pressure compressor and the high pressure turbine, a bearing structure that supports at least one of the inner and outer shafts, and a buffer system. The buffer system prepares buffer air, communicates the buffer air to the bearing structure to pressurize a bearing compartment, and then from the bearing structure along at least one of the shafts.

Abstract: An assembly is provided for rotational equipment. The assembly includes a circumferentially segmented stator and a rotor radially within the stator. The assembly also includes a seal assembly configured for substantially sealing a gap radially between the stator and the rotor. The seal assembly includes a carrier and a non-contact seal seated with the carrier. The carrier includes a plurality of discrete carrier segments circumferentially arranged around the non-contact seal.

Abstract: A cooling system includes an inlet, a passageway, and a feed passage. The inlet is defined by a strut of a diffuser case that extends between a first case portion and a second case portion. The inlet extends from a leading surface towards a trailing surface along a first axis. The passageway is defined by the strut and extends from the inlet towards the second case portion along a second axis that is disposed transverse to the first axis. The feed passage is defined between a body of a tangential onboard injection system that extends between and is connected to an inner vane platform of a guide vane and the second case portion. The feed passage extends along an axis that is disposed parallel to the first axis.