[특허]Gas turbine engine with high speed low pressure turbine section

Gas turbine engine with high speed low pressure turbine section 원문보기

IPC분류정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02C-003/113 F02C-007/36 F02C-009/16 F02K-003/06 F02C-003/06 F02K-003/072
출원번호	US-0682848 (2017-08-22)
등록번호	US-10240526 (2019-03-26)
발명자 / 주소	Suciu, Gabriel L. Schwarz, Frederick M. Ackermann, William K. Kupratis, Daniel Bernard
출원인 / 주소	UNITED TECHNOLOGIES CORPORATION
대리인 / 주소	Carlson, Gaskey & Olds, P.C.
인용정보	피인용 횟수 : 0 인용 특허 : 45

초록 ▼

A compressor section is in fluid communication with a fan, which includes a first compressor section and a second compressor section. A turbine section includes a first turbine section driving the fan and the first compressor section and a second turbine section driving the second compressor section and the second compressor rotor. A first performance quantity is defined as a product of the first speed squared and the first area. A second performance quantity is defined as a product of the second speed squared and the second exit area. A performance ratio of the first performance quantity to the second performance quantity is between about 0.2 and about 0.8. A gear reduction is included between the first turbine section and the first compressor section.

대표청구항 ▼

1. A gas turbine engine comprising: a fan including fan blades, and an outer housing that surrounds said fan to define a bypass flow path;a bypass ratio greater than 10 and less than 22 at a cruise condition;a compressor section in fluid communication with said fan, said compressor section including a first compressor section and a second compressor section;a combustion section in fluid communication with said compressor section;a turbine section in fluid communication with said combustion section;wherein the turbine section includes a first turbine section driving said fan and said first compressor section and a second turbine section driving said second compressor section and a second compressor rotor;wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed;wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is higher than said first speed;wherein a first performance quantity is defined as a product of the first speed squared and said first area;wherein a second performance quantity is defined as a product of the second speed squared and said second exit area;wherein a performance ratio of said first performance quantity to said second performance quantity is between 0.2 and 0.8; andwherein a gear reduction is included between said first turbine section and said first compressor section, such that said first compressor section and said fan rotate at a lower speed than said first turbine section. 2. The gas turbine engine as set forth in claim 1, wherein said gear reduction is a planetary gear reduction. 3. The gas turbine engine as set forth in claim 2, wherein a gear ratio of said gear reduction is greater than 2.0 and less than 5.0. 4. The gas turbine engine as set forth in claim 3, wherein said gear ratio is greater than 2.5. 5. The gas turbine engine as set forth in claim 1, wherein said performance ratio is less than or equal to 0.5. 6. The gas turbine engine as set forth in claim 5, wherein said second speed is between three and four times said first speed. 7. The gas turbine engine as set forth in claim 6, wherein a gear ratio of said gear reduction is greater than 2.0 and less than 5.0. 8. The gas turbine engine as set forth in claim 7, wherein said gear ratio is greater than 2.5. 9. The gas turbine engine as set forth in claim 1, wherein said second speed is greater than twice said first speed and less than three times said first speed. 10. The gas turbine engine as set forth in claim 9, wherein said fan has 26 or fewer fan blades and said first turbine section has at least three stages and has up to six stages. 11. The gas turbine engine as set forth in claim 10, wherein said gear reduction is a planetary gear reduction. 12. The gas turbine engine as set forth in claim 11, wherein a low fan pressure ratio is less than 1.35 across the fan blades alone at said cruise condition. 13. A gas turbine engine comprising: a fan including fan blades, and an outer housing that surrounds said fan to define a bypass flow path;a first turbine section;a second turbine section;wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed, said first turbine section having at least 3 stages;wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than said first speed, said second turbine section having 2 or fewer stages;wherein a first performance quantity is defined as a product of said first speed squared and said first area;wherein a second performance quantity is defined as a product of said second speed squared and said second area;wherein a ratio of said first performance quantity to said second performance quantity is between 0.5 and 1.5;wherein a gear reduction is included between said fan and a low spool driven by said first turbine section such that said fan rotates at a lower speed than said first turbine section; andsaid first and second turbine sections are designed to rotate in opposed directions relative to each other.

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Gas turbine engine with high speed low pressure turbine section 원문보기

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Gas turbine engine with high speed low pressure turbine section 원문보기

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