An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating structure, a stationary structure, a bearing and a gearbox. The rotating structure is configured to rotate about a rotational axis. The rotating structure includes a shaft and a bladed rotor connected to the shaft. The stationary structure circumscribes the rotating structure. The bearing rotatably mounts the shaft to the stationary structure. The gearbox is disposed radially inboard of the bearing.

A gas turbine engine includes a propulsor and a fan drive turbine. The fan drive turbine drives the propulsor through a geared architecture. The geared architecture includes a sun gear, a ring gear, and intermediate gears supported on journal support pins. The sun gear engages the intermediate gears and the intermediate gears engages the ring gear. The journal support pins include a titanium body and an outer surface outside of the titanium body that has a surface hardness that is harder than the titanium body. The outer surface is provided by a steel sleeve. Oil supply holes extend from a central bore in the titanium body through the steel sleeve. At least one pin extends through the steel sleeve to secure the steel sleeve to the titanium body.

A gas turbine engine includes a propulsor and a fan drive turbine. The fan drive turbine drives the propulsor through a geared architecture. The geared architecture includes a sun gear, a ring gear, and intermediate gears supported on journal support pins. The sun gear engages the intermediate gears and the intermediate gears engages the ring gear. The journal support pins include a titanium body and an outer surface outside of the titanium body that has a surface hardness that is harder than the titanium body. The outer surface is provided by a steel sleeve. Oil supply holes extend from a central bore in the titanium body through the steel sleeve. At least one pin extends through the steel sleeve to secure the steel sleeve to the titanium body.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

An assembly is provided for a turbine engine. This turbine engine assembly includes a first rotating structure, a turbine engine apparatus, a rotating coupler and a seal assembly. The first rotating structure is configured to rotate about a rotational axis. The turbine engine apparatus includes an electric machine and a second rotating structure. The electric machine includes an electric machine rotor and an electric machine stator. The second rotating structure is configured to rotate about the rotational axis and is coupled to the electric machine rotor. The rotating coupler is coupled to the first rotating structure by a first connection. The rotating coupler is coupled to the second rotating structure by a second connection. The seal assembly includes a rotating seal land and a stationary seal element. The rotating seal land is mounted onto the rotating coupler. The stationary seal element sealingly engages the rotating seal land.

The present disclosure relates to gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox comprising: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S.sub.tip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M.sub.rel, wherein M.sub.rel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and S.sub.tip is from −1 to 0.1.