Gas turbine engine shaft bearing configuration
A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supporting the shaft relative to the inlet case. The second bearing is arranged radially outward from the shaft.
1. A gas turbine engine comprising:
a core housing including an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path, the inlet case includes an inlet case portion defining the inlet case flow path;
a bearing support portion removably secured to the inlet case portion;
a shaft supporting a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path, the shaft includes a main shaft, a flex shaft axially forward of the main shaft, a hub removably secured to the main shaft and splined to the flex shaft at a spline end of the flex shaft, and the compressor section includes a rotor mounted to the hub;
a geared architecture coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture, wherein the geared architecture includes a sun gear, the sun gear supported by and cantilevered on the shaft;
a first bearing supporting the shaft relative to the intermediate case; and
a second bearing supporting the shaft relative to the inlet case, and the second bearing is arranged radially outward from the shaft and axially forward of the spline end of the flex shaft, the second bearing mounted to the bearing support portion, the hub supporting the second bearing,
wherein the hub comprises a first portion extending axially between and rotationally coupling the main shaft and the flex shaft, and a second portion supporting the second bearing and extending radially outward from the first portion and axially past the spline end of the flex shaft.
2. The gas turbine engine according to claim 1 , wherein the first bearing is supported on the first portion of the hub.
3. The gas turbine engine according to claim 1 , wherein the first bearing is mounted to the hub axially rearward of the spline end of the flex shaft.
4. The gas turbine engine according to claim 1 , wherein:
the intermediate case includes an intermediate case portion defining the intermediate case flow path;
the bearing support portion is a second bearing support portion; and
a first bearing support portion is removably secured to the intermediate case portion, the first bearing mounted to the first bearing support portion.
5. The gas turbine engine according to claim 1 , wherein:
the compressor section is fluidly connected to the fan, the compressor section comprising a first pressure compressor and a second pressure compressor upstream from the first pressure compressor, the second pressure compressor including multiple compressor stages;
the first and second bearings are arranged radially inward of and axially overlapping with at least some of the multiple compressor stages; and
wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6).
6. The gas turbine engine according to claim 5 , further comprising:
a combustor fluidly connected to the compressor section;
a turbine section fluidly connected to the combustor, the turbine section comprising:
a high pressure turbine;
a low pressure turbine; and
wherein the gas turbine engine includes at least one of a low Fan Pressure Ratio of less than 1.45 across a fan blade alone and a low pressure turbine pressure ratio that is greater than 5.
7. The gas turbine engine of claim 5 , wherein the multiple compressor stages include a variable stator vane array, rotatable compressor blades, and a fixed stator vane array.
8. The gas turbine engine of claim 1 , wherein the shaft terminates in an end that supports the sun gear.
9. The gas turbine engine of claim 8 , wherein the first and second bearings are arranged on one axial side of the sun gear, and the sun gear includes another axial side opposite the one axial side that is without direct bearing support.
10. The gas turbine engine of claim 1 , wherein the first bearing is a ball bearing, and the second bearing is a roller bearing.
11. The gas turbine engine of claim 1 , wherein the first and second bearings are arranged in separate sealed lubrication compartments.
12. The gas turbine engine of claim 1 , further comprising a lubrication compartment, the second bearing and the geared architecture arranged in the lubrication compartment.
13. The gas turbine engine of claim 1 , wherein the flex shaft includes a bellow and the second portion axially overlaps the bellow.
14. The gas turbine engine of claim 13 , wherein the second portion comprises a flange that is spaced radially outward from the shaft.
15. The gas turbine engine of claim 1 , wherein the second portion comprises a flange that is spaced radially outward from the shaft.
16. The gas turbine engine of claim 1 , wherein the geared architecture includes a torque frame supporting multiple circumferentially arranged star gears intermeshing with the sun gear, the torque frame secured to the inlet case and the bearing support portion.
17. The gas turbine engine of claim 16 , wherein the second bearing is arranged radially between the torque frame and the flex shaft.
18. The gas turbine engine of claim 16 , wherein the torque frame is secured to the inlet case portion and the bearing support portion at a joint.