[특허]Turbine section architecture for gas turbine engine

Pratt & Whitney Canada Corp.

검색어에 아래의 연산자를 사용하시면 더 정확한 검색결과를 얻을 수 있습니다.
검색연산자

검색도움말
검색연산자	기능	검색시 예
()	우선순위가 가장 높은 연산자	예1) (나노 (기계 \| machine))
공백	두 개의 검색어(식)을 모두 포함하고 있는 문서 검색	예1) (나노 기계) 예2) 나노 장영실
\|	두 개의 검색어(식) 중 하나 이상 포함하고 있는 문서 검색	예1) (줄기세포 \| 면역) 예2) 줄기세포 \| 장영실
!	NOT 이후에 있는 검색어가 포함된 문서는 제외	예1) (황금 !백금) 예2) !image
*	검색어의 *란에 0개 이상의 임의의 문자가 포함된 문서 검색	예) semi*
""	따옴표 내의 구문과 완전히 일치하는 문서만 검색	예) "Transform and Quantization"

과학기술 지식인프라 ScienceON

연합인증

연합인증 가입 기관의 연구자들은 소속기관의 인증정보(ID와 암호)를 이용해 다른 대학, 연구기관, 서비스 공급자의 다양한 온라인 자원과 연구 데이터를 이용할 수 있습니다.

이는 여행자가 자국에서 발행 받은 여권으로 세계 각국을 자유롭게 여행할 수 있는 것과 같습니다.

연합인증으로 이용이 가능한 서비스는 NTIS, DataON, Edison, Kafe, Webinar 등이 있습니다.

한번의 인증절차만으로 연합인증 가입 서비스에 추가 로그인 없이 이용이 가능합니다.

다만, 연합인증을 위해서는 최초 1회만 인증 절차가 필요합니다. (회원이 아닐 경우 회원 가입이 필요합니다.)

연합인증 절차는 다음과 같습니다.

최초이용시에는
ScienceON에 로그인 → 연합인증 서비스 접속 → 로그인 (본인 확인 또는 회원가입) → 서비스 이용

그 이후에는
ScienceON 로그인 → 연합인증 서비스 접속 → 서비스 이용

연합인증을 활용하시면 KISTI가 제공하는 다양한 서비스를 편리하게 이용하실 수 있습니다.

특허 상세정보

MyON담기

내보내기 페이스북 공유 트위터 공유 카카오톡 공유

Turbine section architecture for gas turbine engine

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F01D-025/14
미국특허분류(USC)	415/009; 415/149.4; 416/002; 416/198R
출원번호	US-0858319 (2010-08-17)
등록번호	US-8734085 (2014-05-27)
발명자 / 주소	Lefebvre, Guy Durocher, Eric
출원인 / 주소	Pratt & Whitney Canada Corp.
대리인 / 주소	Norton Rose Fulbright Canada LLP
인용정보	피인용 횟수 : 0 인용 특허 : 19

초록
▼

A turbine section of a gas turbine engine is described, which includes an aftmost rotor extending, an aftmost stator located upstream of the aftmost rotor, and a second to last aftmost rotor located upstream of the aftmost stator. The second to last aftmost rotor being spaced apart a first axial distance from the aftmost stator. A turbine exhaust case is located downstream of the aftmost rotor and includes an inner radial wall and an outer radial wall defining a main gas path duct downstream of the aftmost rotor. The turbine exhaust case being axially sp...

대표
청구항
▼

1. A turbine section of a gas turbine engine defining a main gas path extending therethrough, the turbine section comprising: an aftmost rotor extending radially through the main gas path;an aftmost stator extending radially through the main gas path and located upstream of the aftmost rotor;a second to last aftmost rotor extending radially through the main gas path and located upstream of the aftmost stator, the second to last aftmost rotor being spaced apart a first axial distance from the aftmost stator;a turbine exhaust case located downstream of the aftmost rotor, the turbine exhaust case including an inner radial wall and an outer radial wall defining a main gas path duct downstream of the aftmost rotor, the main gas path duct defining a portion of the main gas path, the turbine exhaust case being axially spaced apart from the aftmost rotor a second axial distance, the second axial distance being greater than the first axial distance; andat least one deformable duct member extending from at least the outer radial wall of the turbine exhaust case, the deformable duct member projecting in an upstream direction towards the aftmost rotor over at least a portion of an axial gap therebetween, the deformable duct member being structurally weaker in an axial direction than in a radial direction, and being substantially frangible when a predetermined maximum axial load is applied thereto, the predetermined maximum axial load being insufficient to deform the turbine exhaust case, such that the deformable duct member is able to absorb an impact force of the aftmost rotor. 2. The turbine section as defined in claim 1, wherein a third axial distance between the aftmost rotor and the aftmost stator is smaller than the second axial distance. 3. The turbine section as defined in claim 1, wherein the deformable duct member at least partially axially overlaps an axially extending aftmost portion of the aftmost rotor. 4. The turbine section as defined in claim 3, wherein the axially extending aftmost portion of the aftmost rotor is a rearwardly extending blade root at a radially inner end of the aftmost rotor. 5. The turbine section as defined in claim 1, wherein the deformable duct member includes at least one of an inner deformable duct member extending from the upstream end of the inner radial wall of the turbine exhaust case and an outer deformable duct member extending from the upstream end of the outer radial wall of the turbine exhaust case. 6. The turbine section as defined in claim 1, wherein the deformable duct member extends substantially axially from the turbine exhaust case. 7. A gas turbine engine comprising: a compressor section, a combustor downstream of the compressor section, a turbine section downstream of the combustor and including a plurality of rotors and a plurality of stators alternately disposed along a main gas path of the gas turbine engine, the plurality of rotors including an aftmost rotor; anda turbine exhaust case disposed downstream of the turbine section, the turbine exhaust case including: an annular outer radial wall located downstream of an aftmost turbine rotor of the turbine section, an annular inner radial wall located downstream of the aftmost turbine rotor of the turbine section, the outer and inner radial walls defining a main gas path duct downstream of the aftmost turbine rotor, the main gas path duct defining therebetween a portion of the main gas path;an outer deformable member extending in an upstream direction from an upstream end of the outer radial wall, the outer deformable member at least partially sealing a first axial gap between the outer radial wall of the turbine exhaust case and the aftmost turbine rotor;an inner deformable member extending in an upstream direction from an upstream end of the inner radial wall, the inner deformable member at least partially sealing a second axial gap between the inner radial wall of the turbine exhaust case and the aftmost turbine rotor; andwherein the outer deformable member and the inner deformable member being structurally weaker in an axial direction than in a radial direction, and the outer and inner deformable members being substantially frangible when a predetermined maximum axial load is applied thereto, the predetermined maximum axial load being insufficient to deform the outer and inner radial walls from which the outer and inner deformable members project, such that the outer and inner deformable members are able to absorb an impact force of the aftmost turbine rotor. 8. The gas turbine engine as defined in claim 7, wherein an aftmost stator is located upstream of the aftmost rotor and a second to last aftmost rotor is located upstream of the aftmost stator, the second to last aftmost rotor being spaced apart a first axial distance from the aftmost stator, and at least one of the outer radial wall and the inner radial wall being spaced apart from the aftmost rotor a second axial distance, the second axial distance being greater than the first axial distance. 9. The gas turbine engine as defined in claim 8, wherein a third axial distance between the aftmost rotor and the aftmost stator is smaller than the second axial distance. 10. The gas turbine engine as defined in claim 7, wherein at least one of the inner deformable members and the outer deformable members axially overlaps a portion of an axially extending aftmost portion of the aftmost rotor. 11. The gas turbine engine as defined in claim 10, wherein the axially extending aftmost portion of the aftmost rotor is a reawardly extending blade root at a radially inner end of the aftmost rotor. 12. The gas turbine engine as defined in claim 7, wherein an aftmost stator is located upstream of the almost rotor and a second to last aftmost rotor is located upstream of the aftmost stator and is spaced apart and separated therefrom by an axial distance, and wherein the axial distance is smaller than both an axial length of the first axial gap and an axial length of the second axial gap. 13. The gas turbine engine as defined in claim 7, wherein at least one of the inner and outer deformable members has a radially extending portion. 14. The gas turbine engine as defined in claim 13, wherein the outer deformable member has the radially extending portion at an upstream end thereof, and the inner deformable portion has the radially extending portion at a downstream end therefore proximate the inner radial wall of the turbine exhaust case.

이 특허에 인용된 특허 (19)

IPC 상위 출원인

내보내기

상세정보
내보내기 구분	파일저장 인쇄 메일전송
구성항목	관리번호, 국가코드, 자료구분, 상태, 출원번호, 출원일자, 공개번호, 공개일자, 등록번호, 등록일자, 발명명칭(한글), 발명명칭(영문), 출원인(한글), 출원인(영문), 출원인코드, 대표IPC 관리번호, 국가코드, 자료구분, 상태, 출원번호, 출원일자, 공개번호, 공개일자, 공고번호, 공고일자, 등록번호, 등록일자, 발명명칭(한글), 발명명칭(영문), 출원인(한글), 출원인(영문), 출원인코드, 대표출원인, 출원인국적, 출원인주소, 발명자, 발명자E, 발명자코드, 발명자주소, 발명자 우편번호, 발명자국적, 대표IPC, IPC코드, 요약, 미국특허분류, 대리인주소, 대리인코드, 대리인(한글), 대리인(영문), 국제공개일자, 국제공개번호, 국제출원일자, 국제출원번호, 우선권, 우선권주장일, 우선권국가, 우선권출원번호, 원출원일자, 원출원번호, 지정국, Citing Patents, Cited Patents
저장형식	Text(ASCII format) Excel format PIAS분석(.xls)
메일정보	받는사람 (필수) @ 보내는사람 (선택) @ 제목 내용 KISTI 검색결과 이메일 서비스
안내	> 총 건의 자료가 검색되었습니다. > 다운받으실 자료의 인덱스를 입력하세요. (1-10,000) > 검색결과의 순서대로 최대 10,000건 까지 다운로드가 가능합니다. > 데이타가 많을 경우 속도가 느려질 수 있습니다.(최대 2~3분 소요) > 다운로드 파일은 UTF-8 형태로 저장됩니다. 파일의 내용이 제대로 보이지 않을실 때는 웹브라우저 상단의 보기 -> 인코딩 -> 자동선택 여부를 확인하십시오 > ~ > Text(ASCII format) Excel format

과학기술 지식인프라 ScienceON

ScienceON Chatbot

연합인증

특허 상세정보

Turbine section architecture for gas turbine engine

초록
▼

대표
청구항
▼

이 특허에 인용된 특허 (19)

IPC 상위 출원인

과제정보

상세정보조회

원문조회

IPC 상세정보(version : 201201)

연구과제 성과물(0)

참여연구원의 다른 문헌(0)

처리중입니다.

잠시만 기다려주세요.

ScienceON Chatbot

연합인증

특허 상세정보

Turbine section architecture for gas turbine engine

초록 ▼

대표청구항 ▼

이 특허에 인용된 특허 (19)

IPC 상위 출원인

과제정보

상세정보조회

원문조회

IPC 상세정보(version : 201201)

연구과제 성과물(0)

참여연구원의 다른 문헌(0)

처리중입니다.

잠시만 기다려주세요.

초록
▼

대표
청구항
▼