LPC flowpath shape with gas turbine engine shaft bearing configuration

• US 9,038,366 B2
• Filed: 10/30/2013
• Issued: 05/26/2015
• Est. Priority Date: 01/31/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Merry, Brian D., Suciu, Gabriel L., Brilliant, Lisa I., Rose, Becky E., Dong, Yuan, Balamucki, Stanley J.
• Primary Examiner(s): SUNG, GERALD LUTHER
• Application Number:
  US14/067,354
• Publication Number:
  US 20140248129A1
• Time in Patent Office: 573 Days
• Field of Search: 602/261, 602/68, 607/95
• US Class Current: 60/226.1
• CPC Class Codes: F01D 25/162 Bearing supports
• CPC Class Codes: F01D 5/143 Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
• CPC Class Codes: F02C 3/13 having variable working fluid interconnections between turbines or compressors or stages of different rotors controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075
• CPC Class Codes: F02C 7/36 Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user F02C3/107 - F02C3/13 and F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H
• CPC Class Codes: F02C 9/18 by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages F02C3/113 takes precedence
• CPC Class Codes: F02K 3/04 the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

Claims

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 flowpath, wherein the inlet case includes a first inlet case portion defining the inlet case flow path, and a bearing support portion is removably secured to the inlet case portion;

a fan arranged fluidly upstream of the core housing;

a shaft providing a rotational axis;

a geared architecture coupled to the shaft, and the fan coupled to and rotationally driven by the geared architecture;

a hub operatively supported by the shaft;

a rotor connected to the hub and supporting a compressor section arranged fluidly downstream from the inlet case, the compressor section being arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path, the core flowpath having an inner diameter and an outer diameter, wherein the at least a portion of the inner diameter has an increasing slope angle relative to the rotational axis;

a first bearing mounted to the hub and supporting the shaft relative to one of the intermediate case and the inlet case, the first bearing mounted to the bearing support portion; and

a second bearing mounted to the hub and supporting the shaft relative to the other of the intermediate case and the inlet case.

2. The gas turbine engine according to claim 1, wherein the outer diameter has an outer diameter slope angle relative to the rotational axis along a fluid flow direction of the core flow path of between 0 degrees and 15 degrees.

3. The gas turbine engine according to claim 1, wherein the fan is connected to the shaft through the geared architecture, wherein the compressor section is a low pressure compressor.

4. The gas turbine engine according to claim 2, wherein the outer diameter slope angle decreases relative to the rotational axis.

5. The gas turbine engine according to claim 2, wherein the outer diameter slope angle is in the range of 0 degrees to 10 degrees.

6. The gas turbine engine according to claim 5, wherein the outer diameter slope angle is in the range of 5 degrees to 7 degrees.

7. The gas turbine engine according to claim 6, wherein the outer diameter slope angle is 6 degrees.

8. The gas turbine engine according to claim 4, wherein the fan is connected to the shaft through the geared architecture, wherein the compressor section is a low pressure compressor.

9. The gas turbine engine according to claim 8, wherein the low pressure compressor comprises at least one variable vane.

10. The gas turbine engine according to claim 8, wherein the low pressure compressor further comprises an exit guide vane, wherein the exit guide vane is located in a low pressure compressor outlet section of the core flow path.

11. The gas turbine engine according to claim 10, wherein the low pressure compressor further comprises a low pressure bleed located between a low pressure compressor rotor and the exit guide vane.

12. The gas turbine engine according to claim 11, wherein the low pressure bleed further comprises a bleed trailing edge, and wherein the bleed trailing edge extends into the core flow path beyond the outer diameter of the core flow path.

13. The gas turbine engine according to claim 8, wherein the low pressure compressor is a multi-stage compressor.

14. The gas turbine engine according to claim 1, wherein the intermediate case includes an intermediate case portion defining the intermediate case flow path, and a second bearing support portion removably secured to the intermediate case portion, the second mounted to the second bearing support portion.

15. The gas turbine engine according to claim 14, wherein the other bearing is a ball bearing.

16. The gas turbine engine according to claim 1, wherein the first and second bearings are arranged in separate sealed lubrication compartments.

17. The gas turbine engine according to claim 1, wherein the shaft includes a main shaft and a flex shaft, the flex shaft secured to the main shaft at a first end and including a second end opposite the first end, wherein the geared architecture includes a sun gear supported on the second end.

18. The gas turbine engine according to claim 17, wherein the hub is secured to the main shaft, and the compressor section includes a rotor mounted to the hub.

19. The gas turbine engine according to claim 18, 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.

20. The gas turbine engine according to claim 1, wherein the rotor supports multiple compressor stages, and the first bearing is axially aligned with and radially inward of one of the compressor stages.

21. The gas turbine engine according to claim 20, wherein the compressor section includes a variable vane array.

22. The gas turbine engine according to claim 1, comprising a lubrication compartment, wherein the geared architecture is arranged in the lubrication compartment.