Geared turbofan engine with optimized diffuser case flange location
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-003/04
F01D-025/24
F02C-006/08
F02C-003/10
F02C-007/36
출원번호
US-0427664
(2013-03-14)
등록번호
US-9970323
(2018-05-15)
국제출원번호
PCT/US2013/031301
(2013-03-14)
국제공개번호
WO2014/058466
(2014-04-17)
발명자
/ 주소
Schwarz, Frederick M.
Suciu, Gabriel L.
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Carlson, Gaskey & Olds, P.C.
인용정보
피인용 횟수 :
0인용 특허 :
16
초록▼
A gas turbine engine has a compressor section, a compressor case substantially surrounding the compressor section, and a diffuser case attached at an attachment interface to the compressor case. The attachment interface is between a forward and an aft end of the compressor. A geared turbofan is also
A gas turbine engine has a compressor section, a compressor case substantially surrounding the compressor section, and a diffuser case attached at an attachment interface to the compressor case. The attachment interface is between a forward and an aft end of the compressor. A geared turbofan is also disclosed.
대표청구항▼
1. A gas turbine engine comprising: a compressor section;a compressor case substantially surrounding the compressor section;a diffuser case having a first end and a second end, the diffuser case being attached at an attachment interface at the first end to the compressor case, the attachment interfa
1. A gas turbine engine comprising: a compressor section;a compressor case substantially surrounding the compressor section;a diffuser case having a first end and a second end, the diffuser case being attached at an attachment interface at the first end to the compressor case, the attachment interface being between a forward and an aft end of the compressor section, and wherein the attachment interface comprises an external flange connection;a support structure for an outer flowpath static structure that has an internal flange connection to the diffuser case that is positioned axially aft of the external flange connection; andat least one bleed formed in the diffuser case, wherein the bleed is positioned aft of the external flange connection and forward of the internal flange connection. 2. The gas turbine engine according to claim 1 wherein the compressor section includes a first compressor operating at a first pressure and a second compressor operating at a second pressure that is higher than the first pressure, and wherein the first compressor defines the forward end and the second compressor defines the aft end of the compressor section. 3. The gas turbine engine according to claim 2 wherein the second compressor comprises a plurality of stages, and wherein the diffuser case is configured to surround at least one stage of the second compressor. 4. The gas turbine engine according to claim 2 wherein the second compressor comprises a plurality of stages, and wherein the diffuser case is configured to surround at least two stages of the second compressor. 5. The gas turbine engine according to claim 1 wherein the second end of the diffuser case is attached to a turbine case at a turbine attachment interface that is positioned aft of the internal flange connection, and wherein the at least one bleed is positioned adjacent the first end of the diffuser case. 6. The gas turbine engine according to claim 1 wherein the at least one bleed comprises an air supply boss configured to supply bleed air for cabin pressurization under a predetermined condition, and wherein the air supply boss defines an internal air passage that is in fluid communication with a high pressure compressor flow path. 7. The gas turbine engine according to claim 6 wherein the predetermined condition comprises a high altitude and low engine power condition. 8. The gas turbine engine according to claim 1 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that include an aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the aft-most disk. 9. The gas turbine engine according to claim 1 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that includes a second-to-aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the second-to-aftmost disk. 10. The gas turbine engine according to claim 1 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that includes a third-to-aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the third-to-aftmost disk. 11. A geared turbofan engine comprising: a fan including a plurality of fan blades rotatable about an axis;a core engine including a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor;a compressor case supporting the compressor section;a diffuser case having a first end and a second end, the diffuser case being attached at an attachment interface at the first end to the compressor case, the attachment interface being between a forward and an aft end of the compressor section, and wherein the attachment interface comprises an external flange connection;a support structure for an outer flowpath static structure that has an internal flange connection to the diffuser case that is positioned axially aft of the external flange connection; andat least one bleed formed in the diffuser case, wherein the bleed is positioned aft of the external flange connection and forward of the internal flange connection; anda geared architecture driven by the turbine section for rotating the fan about the axis. 12. The geared turbofan engine according to claim 11 wherein the compressor section includes a first compressor operating at a first pressure and a second compressor operating at a second pressure that is higher than the first pressure, and wherein the first compressor defines the forward end and the second compressor defines the aft end of the compressor section. 13. The geared turbofan engine according to claim 12 wherein the second compressor comprises a plurality of stages, and wherein the diffuser case is configured to surround at least one stage of the second compressor. 14. The geared turbofan engine according to claim 11 wherein the second end of the diffuser case is attached to a turbine case at a turbine attachment interface that is positioned aft of the internal flange connection, and wherein the at least one bleed is positioned adjacent the first end of the diffuser case. 15. The geared turbofan engine according to claim 11 wherein the at least one bleed comprises an air supply boss configured to supply bleed air for cabin pressurization under a predetermined condition, and wherein the air supply boss defines an internal air passage that is in fluid communication with a high pressure compressor flow path. 16. The geared turbofan engine according to claim 15 wherein the predetermined condition comprises a high altitude and low engine power condition. 17. The geared turbofan engine according to claim 11 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that includes an aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the aft-most disk. 18. The geared turbofan engine according to claim 11 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that includes a second-to-aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the second-to-aftmost disk. 19. The geared turbofan engine according to claim 11 wherein the attachment interface defines a parting line between the compressor case and the diffuser case, and wherein the compressor section has a plurality of disks that includes a third-to-aftmost disk that is defined by a disk center plane, and wherein the parting line is at or forward of the center plane of the third-to-aftmost disk. 20. The geared turbofan engine according to claim 11 wherein the compressor case includes a compressor mount flange at an aft compressor case end and the diffuser case includes a diffuser mount flange at the first end of the diffuser case, and wherein the compressor and diffuser mount flanges are in direct abutting engagement with each other at a parting line and are secured together by at least one fastener to form the attachment interface, and wherein the parting line is axially forward of an aftmost stage of the compressor section. 21. The gas turbine engine according to claim 1 wherein the compressor case includes a compressor mount flange at an aft compressor case end and the diffuser case includes a diffuser mount flange at the first end of the diffuser case, and wherein the compressor and diffuser mount flanges are in direct abutting engagement with each other at a parting line and are secured together by at least one fastener to form the attachment interface, and wherein the parting line is axially forward of an aftmost stage of the compressor section.
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