Gas turbine engine architecture with low pressure compressor hub between high and low rotor thrust bearings
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-007/06
F02C-003/107
F02K-003/06
F01D-025/16
출원번호
US-0304053
(2011-11-23)
등록번호
US-9784181
(2017-10-10)
발명자
/ 주소
Davis, Todd A.
Reinhardt, Gregory E.
DiBenedetto, Enzo
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Carlson, Gaskey & Olds, P.C.
인용정보
피인용 횟수 :
0인용 특허 :
14
초록
A method for servicing a gas turbine engine includes providing access from a forward section of the gas turbine engine to a gearbox contained within a bearing compartment.
대표청구항▼
1. A gas turbine engine comprising: a low pressure compressor along an axis;a first bearing system which at least partially supports an inner shaft along said axis;a second bearing system which at least partially supports an outer shaft along said axis, the second bearing system positioned upstream
1. A gas turbine engine comprising: a low pressure compressor along an axis;a first bearing system which at least partially supports an inner shaft along said axis;a second bearing system which at least partially supports an outer shaft along said axis, the second bearing system positioned upstream from inlet stage vanes of a high pressure compressor; anda low pressure compressor hub mounted to said inner shaft, said low pressure compressor hub extends forwardly to said low pressure compressor between said first bearing system and said second bearing system, wherein said low pressure compressor hub supports a plurality of stages of said low pressure compressor, wherein at least one of the plurality of stages is positioned axially between the first bearing system and the second bearing system, said inner shaft to drive a fan through a geared architecture,wherein said low pressure compressor hub includes a tubular hub for supporting the inner circumference portion of said first bearing system and a frustro-conical web extending from a rear portion of said tubular hub to said low pressure compressor across the space between said first bearing system and said second bearing system. 2. The gas turbine engine as recited in claim 1, wherein said frustro-conical web extends at least partially around said first bearing system. 3. The gas turbine engine as recited in claim 2, wherein said low pressure compressor is radially outboard of said first bearing system. 4. The gas turbine engine as recited in claim 1, wherein said low pressure compressor hub is angled toward said low pressure compressor. 5. The gas turbine engine as recited in claim 1, wherein low pressure compressor hub is mounted to a second stage disk of said low pressure compressor. 6. The gas turbine engine as recited in claim 5, wherein said low pressure compressor includes three stages. 7. The gas turbine engine as recited in claim 1, wherein said first bearing system is mounted to a front center body case structure, said front center body case structure defines a core flow path for a core airflow. 8. The gas turbine engine as recited in claim 7, wherein said second bearing system is mounted to an intermediate case structure, said intermediate case structure mounted to said front center body case structure to continue said core flow path. 9. A gas turbine engine comprising: a front center body case structure;a geared architecture at least partially supported by said front center body case structure;a first bearing system mounted to said front center body case structure to rotationally support an inner shaft;a coupling shaft mounted to said inner shaft and said geared architecture, said coupling shaft at least partially supported by said first bearing support; anda low pressure compressor hub mounted to said inner shaft, said low pressure compressor hub extends forwardly relative to a general direction of flow through the gas turbine engine to a low pressure compressor between said first bearing system and a second bearing system which at least partially supports an outer shaft, the second bearing system positioned upstream from inlet stage vanes of a high pressure compressor, wherein at least a portion of said coupling shaft axially overlaps at least a portion of a compressor rotor supported by said low pressure compressor hub, wherein said low pressure compressor includes a plurality of stages and at least one of the plurality of stages is positioned axially between the first bearing system and the second bearing system;wherein said low pressure compressor hub includes a tubular hub for supporting the inner circumference portion of said first bearing system and a frustro-conical web extending from a rear portion of said tubular hub to said low pressure compressor across the space between said first bearing system and said second bearing system. 10. The gas turbine engine as recited in claim 9, wherein said inner shaft drives a fan through said geared architecture. 11. The gas turbine engine as recited in claim 9, wherein said outer shaft at least partially surrounds said inner shaft, said outer shaft drives a high pressure compressor. 12. The gas turbine engine as recited in claim 9, wherein said low pressure compressor includes three stages, said low pressure compressor hub mounted to a low pressure compressor rotor at a position that is axially aligned with a second stage disk of said low pressure compressor. 13. A gas turbine engine comprising: a front center body case structure along an engine axis, said front center body case structure defines a core flow path;a low pressure compressor along said core flow path;an intermediate case structure mounted aft of said front center body case structure along said engine axis;a first bearing system mounted to said front center body case structure to at least partially support an inner shaft for rotation about said engine axis;a second bearing system mounted to said intermediate case structure to at least partially support an outer shaft for rotation about said engine axis, said second bearing system upstream from an inlet stage of a high pressure compressor; anda low pressure compressor hub mounted to said inner shaft, said low pressure compressor hub is attached to a low pressure compressor rotor that extends forwardly from said low pressure compressor rotor to said low pressure compressor between said first bearing system and said second bearing system, said low pressure compressor hub extending from a position that is radially inboard said first bearing system to a position that is radially outboard of said first bearing system, wherein said low pressure compressor includes a plurality of stages and at least one of the plurality of stages is positioned axially between the first bearing system and the second bearing system, said inner shaft to drive said fan through a geared architecture,wherein said low pressure compressor hub includes a tubular hub for supporting the inner circumference portion of said first bearing system and a frustro-conical web extending from a rear portion of said tubular hub to said low pressure compressor across the space between said first bearing system and said second bearing system. 14. The gas turbine engine as recited in claim 13, wherein said front center body case structure is downstream of a fan. 15. The gas turbine engine as recited in claim 14, wherein said geared architecture is at least partially supported by said front center body case structure. 16. The gas turbine engine as recited in claim 2, wherein said frustro-conical web supports rotating blades of a stage of said low pressure compressor that is axially rearward said first bearing system. 17. The gas turbine engine as recited in claim 1, wherein said low pressure compressor hub is connected to a low pressure compressor rotor that extends from a position that is spaced from and axially rearward of said first bearing system to a position that is spaced from and axially forward of said first bearing system. 18. The gas turbine engine as recited in claim 9, wherein said low pressure compressor hub is connected to a low pressure compressor rotor that extends from a position that is spaced from and axially rearward of said first bearing system to a position that is spaced from and axially forward of said first bearing system. 19. The gas turbine engine as recited in claim 13, wherein the position that is radially inboard said first bearing system is radially spaced from said first bearing system and the position that is radially outboard of said first bearing system is radially spaced from said first bearing system, wherein said low pressure compressor hub is connected to a low pressure compressor rotor that extends from a position that is spaced from and axially rearward of said first bearing system to a position that is spaced from and axially forward of said first bearing system. 20. The gas turbine engine as recited in claim 16, wherein said frustro-conical web further supports rotating blades of a stage of said low pressure compressor that is axially forward said first bearing system. 21. The gas turbine engine as recited in claim 1, wherein the geared architecture is upstream from the first bearing system. 22. The gas turbine engine as recited in claim 9, wherein the geared architecture is upstream from the first bearing system. 23. The gas turbine engine as recited in claim 13, wherein the geared architecture is upstream from the first bearing system.
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이 특허에 인용된 특허 (14)
Sheridan William G. (Southington CT) Pagluica Gino J. (Manchester CT), Coupling system for a planetary gear train.
Wiley ; III Walter H. (Palm Beach Gardens FL) Aaron ; Jr. Charles D. (Palm Beach Gardens FL) Carlson Russell L. (North Palm Beach FL) Davis ; III Charles L. (Palm Beach Gardens FL) Marmol Ronald A. (, Fluid damper for thrust bearing.
Barbic John R. (Tequesta FL) Nichol Kurt L. (Estill Springs TN) Hibner David H. (Ashford CT) Szafir David R. (Ellington CT), Variable stiffness oil film damper.
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