A gas turbine engine includes a fan that has fan blades wherein the fan delivers airflow to a bypass duct. A gearbox is defined along an engine axis. A low spool is arranged aft of the gearbox and coupled to drive the gearbox. A front center body assembly is defined around the engine axis. A flexible support supports the gearbox relative to the front center body assembly. A bearing package is mounted to the front center body assembly and the low spool. A front wall is mounted to the front center body assembly. The front wall is removable from the front center body assembly to access at least one of the gearbox or the bearing package. The low spool includes a low pressure compressor hub that provides a spline for engaging the bearing package.

Aircraft turbine engine (10), comprising:—at least one first turbine rotor (22a) connected to a first turbine shaft (36) and comprising rotor blades located in a stream (V),—an annular exhaust casing (28) supporting at least one bearing (60, 62) for guiding the rotation of the first shaft and comprising arms (28a) located in the stream, downstream of the blades, and—a mechanical planetary gearbox (42) which comprises a sun gear (44), a ring gear (40) and a planet carrier (46), the gearbox being at least partially surrounded by the exhaust casing and one of the rotatable elements chosen from the sun gear and the ring gear being connected to the first shaft, characterised in that the exhaust casing carries a device (70) for recovering and discharging oil projected centrifugally by the reduction gear.

Aircraft turbine engine (10), comprising:—at least one first turbine rotor (22a) connected to a first turbine shaft (36) and comprising rotor blades located in a stream (V),—an annular exhaust casing (28) supporting at least one bearing (60, 62) for guiding the rotation of the first shaft and comprising arms (28a) located in the stream, downstream of the blades, and—a mechanical planetary gearbox (42) which comprises a sun gear (44), a ring gear (40) and a planet carrier (46), the gearbox being at least partially surrounded by the exhaust casing and one of the rotatable elements chosen from the sun gear and the ring gear being connected to the first shaft, characterised in that the exhaust casing carries a device (70) for recovering and discharging oil projected centrifugally by the reduction gear.

An auxiliary power unit, in particular for an aircraft, includes a gas generator and an electric generator. The electric generator has an input shaft that is connected to an output shaft of the gas generator. The auxiliary power unit also has an accessories housing directly connected to a rotor shaft of the electric generator, wherein the electric generator is interposed between the gas generator and the accessories housing.

A roller bearing arrangement for a gas turbine engine. The roller bearing arrangement includes a fan shaft, and a stub shaft connected to the fan shaft. The roller bearing arrangement further includes a plurality of roller bearing elements positioned between a first axial bearing surface created on a radially outer surface of the stub shaft and a second axial bearing surface of a static structure, the roller bearing arrangement further including a first snubber positioned between the radially outer surface of the fan shaft and a radially inner surface of the stub shaft, the first snubber being spaced apart from the radially inner surface of the stub shaft or the radially outer surface of the fan shaft so as to limit a radial movement range of the stub shaft.

A roller bearing arrangement for a gas turbine engine. The roller bearing arrangement includes a fan shaft, and a stub shaft connected to the fan shaft. The roller bearing arrangement further includes a plurality of roller bearing elements positioned between a first axial bearing surface created on a radially outer surface of the stub shaft and a second axial bearing surface of a static structure, the roller bearing arrangement further including a first snubber positioned between the radially outer surface of the fan shaft and a radially inner surface of the stub shaft, the first snubber being spaced apart from the radially inner surface of the stub shaft or the radially outer surface of the fan shaft so as to limit a radial movement range of the stub shaft.

A gas turbine engine according to an example of the present disclosure may include, among other things, a fan section including a fan having a plurality of fan blades and including an outer housing surrounding the fan blades to establish a bypass duct, a geared architecture, a first spool including a first shaft that interconnects a first compressor and a fan drive turbine, the fan drive turbine driving the fan through the geared architecture. The gas turbine engine is rated to provide an amount of thrust at ground idle, and the gas turbine engine is rated to provide an amount of thrust at maximum takeoff. A thrust ratio is defined as a ratio of the amount of thrust at ground idle divided by the amount of thrust at maximum takeoff. The thrust ratio can be less than or equal to 0.050.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor includes a pressure ratio of greater than 6.5 and less than 11.5. A ratio of a product of a pressure ratio of the fan with a pressure ratio of the low pressure compressor pressure to the pressure ratio of the high pressure compressor is greater than 0.35 and less than 0.90. An exhaust gas exit temperature is greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor includes a pressure ratio of greater than 6.5 and less than 11.5. A ratio of a product of a pressure ratio of the fan with a pressure ratio of the low pressure compressor pressure to the pressure ratio of the high pressure compressor is greater than 0.35 and less than 0.90. An exhaust gas exit temperature is greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.

A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.