IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0159760
(2014-01-21)
|
등록번호 |
US-8834099
(2014-09-16)
|
발명자
/ 주소 |
- Topol, David A.
- Morin, Bruce L.
|
출원인 / 주소 |
- United Technoloiies Corporation
|
대리인 / 주소 |
Carlson, Gaskey & Olds, PC
|
인용정보 |
피인용 횟수 :
12 인용 특허 :
19 |
초록
▼
A gas turbine engine comprises a fan and a turbine section having a first turbine rotor. The first turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor. The compressor rotor has a number of com
A gas turbine engine comprises a fan and a turbine section having a first turbine rotor. The first turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor: (the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz; andthe rotational speed being in revolutions per minute. A compressor module and a method of designing a gas turbine engine are also disclosed.
대표청구항
▼
1. A gas turbine engine comprising: a fan and a turbine section having a first turbine rotor, the first turbine rotor to drive a compressor rotor;a gear reduction effecting a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor;the compressor rotor having a num
1. A gas turbine engine comprising: a fan and a turbine section having a first turbine rotor, the first turbine rotor to drive a compressor rotor;a gear reduction effecting a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor;the compressor rotor having a number of compressor blades in at least one of a plurality of rows of the compressor rotor, and the blades operating at least some of the time at a rotational speed, and the number of compressor blades in the at least one row and the rotational speed being such that the following formula holds true for the at least one row of the compressor rotor: (the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz; andthe rotational speed being in revolutions per minute. 2. The gas turbine engine as set forth in claim 1, wherein the formula results in a number greater than or equal to about 6000 Hz. 3. The gas turbine engine as set forth in claim 2, wherein the gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 4. The gas turbine engine as set forth in claim 1, wherein the formula holds true for the majority of the blade rows of the compressor rotor. 5. The gas turbine engine as set forth in claim 1, wherein the gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 6. The gas turbine engine as set forth in claim 1, wherein the gear reduction has a gear ratio of greater than about 2.3. 7. The gas turbine engine as set forth in claim 6, wherein the gear reduction has a gear ratio of greater than about 2.5. 8. The gas turbine engine as set forth in claim 1, wherein the fan delivers air into a bypass duct, and a portion of air into the compressor rotor, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor rotor, and the bypass ratio being greater than about 6. 9. The gas turbine engine as set forth in claim 8, wherein the bypass ratio is greater than about 10. 10. The gas turbine engine as set forth in claim 9, wherein the formula results in a number greater than or equal to about 6000 Hz. 11. The gas turbine engine as set forth in claim 1, wherein the rotational speed being an approach speed. 12. The gas turbine engine as set forth in claim 1, wherein the turbine section including a higher pressure turbine rotor and the first turbine rotor, and the fan drive turbine rotor being the first turbine rotor. 13. The gas turbine engine as set forth in claim 12, wherein the compressor rotor is a lower pressure compressor rotor, and the higher pressure turbine rotor driving a higher pressure compressor rotor. 14. The gas turbine engine as set forth in claim 12, wherein the gear reduction is intermediate the first turbine rotor and the compressor rotor. 15. The gas turbine engine as set forth in claim 12, wherein the gear reduction is intermediate the compressor rotor and the fan. 16. The gas turbine engine as set forth in claim 1, wherein the turbine section includes three turbine rotors, with the first turbine rotor and the fan drive turbine rotor being distinct rotors. 17. A method of designing a gas turbine engine comprising the steps of: including a first turbine rotor to drive a compressor rotor and a fan drive turbine rotor to drive a fan through a gear reduction, and selecting a number of blades in at least one row of the compressor rotor, in combination with a rotational speed of the compressor rotor, such that the following formula holds true for the at least one row of the compressor rotor: (the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz; andthe rotational speed being in revolutions per minute. 18. The method as set forth in claim 17, wherein the formula results in a number greater than or equal to about 6000 Hz. 19. The method as set forth in claim 17, wherein the gas turbine engine is rated to produce about 15,000 pounds of thrust or more. 20. The method as set forth in claim 17, wherein the rotational speed is an approach speed. 21. The method as set forth in claim 17, wherein the turbine section including a higher pressure turbine rotor and the first turbine rotor, and the fan drive turbine rotor being the first turbine rotor. 22. The method as set forth in claim 21, wherein the compressor rotor is a lower pressure compressor rotor, and the higher pressure turbine rotor driving a higher pressure compressor rotor. 23. The method as set forth in claim 22, wherein the gear reduction is intermediate the first turbine rotor and the compressor rotor. 24. The method as set forth in claim 22, wherein the gear reduction is intermediate the compressor rotor and the fan. 25. The method as set forth in claim 17, wherein the turbine section includes three turbine rotors, with the first turbine rotor and the fan drive turbine rotor being distinct rotors. 26. A compressor module comprising: a compressor rotor having at least a first blade row that includes a number of blades, the first blade row being capable of rotating at a rotational speed, so that when measuring the rotational speed in revolutions per minute: (the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz. 27. The compressor module as set forth in claim 26, wherein the rotational speed is an approach speed. 28. The compressor module as set forth in claim 27, wherein the formula results in a number greater than or equal to about 6000 Hz. 29. The compressor module as set forth in claim 26, wherein the formula results in a number greater than or equal to about 6000 Hz. 30. A method of designing a gas turbine engine comprising: providing a fan and a turbine section having a first turbine rotor, the first turbine rotor to drive a compressor rotor;providing a gear reduction effecting a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor;the compressor rotor having a number of compressor blades in at least one of a plurality of rows of the compressor rotor, and the blades operating at least some of the time at a rotational speed, and the number of compressor blades in the at least one row and the rotational speed being such that the following formula holds true for the at least one row of the compressor rotor: (the number of blades×the rotational speed)/(60 seconds/minute)≧5500 Hz; andthe rotational speed being in revolutions per minute.
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