[특허]Geared turbofan gas turbine engine architecture

United Technologies Corporation

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특허 상세정보

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Geared turbofan gas turbine engine architecture

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02C-007/36 F02K-003/062 F02K-003/06 F02K-003/065 F02C-003/113 F02K-003/04 F02C-003/107
출원번호	US-0662505 (2015-03-19)
등록번호	US-9222417 (2015-12-29)
발명자 / 주소	Kupratis, Daniel Bernard Schwarz, Frederick M.
출원인 / 주소	United Technologies Corporation
대리인 / 주소	Carlson, Gaskey & Olds, P.C.
인용정보	피인용 횟수 : 0 인용 특허 : 36

초록
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A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The fan drive turbine includes a plurality of turbine rotors. A fan includes a plurality of blades rotatable about an axis and a ratio between the number of fan blades and the number of fan drive turbine rotors is between about 2.5 and about 8.5. A speed change system is driven by the fan drive turbine for rotating the fan about the axis. The fan drive turbine has a first exit area at a first exit point and rotates at a first speed. The second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area. A performance ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5.

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1. A gas turbine engine comprising: a combustor;a turbine section in fluid communication with the combustor, the turbine section including a fan drive turbine and a second turbine section, the fan drive turbine including a plurality of turbine rotors;a fan including a plurality of blades rotatable about an axis and a ratio between a number of fan blades and a number of fan drive turbine rotors is between about 2.5 and about 8.5; anda speed change system driven by the fan drive turbine for rotating the fan about the axis;wherein the fan drive turbine has a first exit area at a first exit point and rotates at a first speed, the second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than the first speed, wherein a first performance quantity is defined as a product of the first speed squared and the first exit area, a second performance quantity is defined as a product of the second speed squared and the second exit area, and a performance ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5. 2. The gas turbine engine as recited in claim 1, wherein a first bearing assembly and a second bearing assembly comprise roller bearings. 3. The gas turbine engine as recited in claim 2, including a compressor section that comprises a first compressor driven by the fan drive turbine through a first shaft and a second compressor section driven by the second turbine section through a second shaft, wherein the first bearing supports an aft portion of the first shaft and the second bearing supports an aft portion of the second shaft. 4. The gas turbine engine as recited in claim 3, wherein a forward portion of each of the first and second shafts are supported by a thrust bearing assembly. 5. The gas turbine engine as recited in claim 1, wherein the performance ratio is above or equal to about 0.8. 6. The gas turbine engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan and the fan drive turbine both rotate in a first direction about the axis and the second turbine section rotates in a second direction opposite the first direction. 7. The gas turbine engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan, the fan drive turbine, and the second turbine section all rotate in a first direction about the axis. 8. The gas turbine engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan and the second turbine section both rotate in a first direction about the axis and the fan drive turbine rotates in a second direction opposite the first direction. 9. The gas turbine engine as recited in claim 1, wherein the speed change system comprises a gearbox, and wherein the fan is rotatable in a first direction whereas the fan drive turbine; and the second turbine section rotate in a second direction opposite the first direction about the axis. 10. The gas turbine engine as recited in claim 1, wherein a fan pressure ratio across the fan is less than about 1.5. 11. The gas turbine engine as recited in claim 1, wherein the fan has 26 or fewer blades. 12. The gas turbine engine as recited in claim 1, wherein the fan drive turbine has between about 3 and 6 stages. 13. The gas turbine engine as recited in claim 1, further comprising: a thrust in pounds force (lbf) produced by the gas turbine engine at a sea level take-off condition; anda volume (in3) of the turbine section,wherein the thrust divided by the turbine section volume is greater than about 1.5 lbf/in3 and less than or equal to about 5.5 lbf/in3. 14. The gas turbine engine as recited in claim 1, wherein the second turbine includes at least two stages and performs at a first pressure ratio and the fan drive turbine includes more than two stages and performs at a second pressure ratio less than the first pressure ratio. 15. The gas turbine engine as recited in claim 1, wherein the fan drive turbine includes a first aft rotor attached to a first shaft and the second turbine section includes a second aft rotor attached to a second shaft and a first bearing assembly and a second bearing assembly are disposed axially aft of the combustor with the first bearing assembly being disposed axially aft of a first connection between the first aft rotor and the first shaft, and the second bearing assembly being disposed axially forward of a second connection between the second aft rotor and the second shaft. 16. The gas turbine engine as recited in claim 1, wherein the turbine section includes three turbine sections, the fan drive turbine driving the fan, the second turbine section and a third turbine section, wherein the second turbine section and the third turbine section each drive a compressor rotor of a compressor section. 17. The gas turbine engine as recited in claim 1, wherein the speed change system is positioned intermediate a compressor rotor and a shaft driven by the fan drive turbine.

이 특허에 인용된 특허 (36)

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Geared turbofan gas turbine engine architecture

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Geared turbofan gas turbine engine architecture

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이 특허에 인용된 특허 (36)

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