Efficient, low pressure ratio propulsor for gas turbine engines
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02K-003/06
F01D-017/14
F02K-001/06
출원번호
US-0252811
(2016-08-31)
등록번호
US-10233868
(2019-03-19)
발명자
/ 주소
Gallagher, Edward J.
Monzon, Byron R.
출원인 / 주소
UNITED TECHNOLOGIES CORPORATION
대리인 / 주소
Carlson, Gaskey & Olds, P.C.
인용정보
피인용 횟수 :
0인용 특허 :
33
초록▼
A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A
A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A first turbine is coupled with the first shaft such that rotation of the first turbine is configured to drive the fan, through the first shaft and gear assembly, at a lower speed than the first shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 (N) of the fan blades, a solidity value (R) that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0.
대표청구항▼
1. A gas turbine engine comprising: a gear assembly;a bypass flow passage and a core flow passage, the bypass flow passage including an inlet;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first shaft, the first shaft coupled through t
1. A gas turbine engine comprising: a gear assembly;a bypass flow passage and a core flow passage, the bypass flow passage including an inlet;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first shaft, the first shaft coupled through the gear assembly with the fan;wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades, the number (N) being 12, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0. 2. The gas turbine engine as recited in claim 1, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 3. The gas turbine engine as recited in claim 1, wherein the bypass flow passage includes an outlet, the inlet and the outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being approximately 1.3 to 1.55. 4. The gas turbine engine as recited in claim 3, wherein the design pressure ratio is between 1.3 and 1.4. 5. The gas turbine engine as recited in claim 4, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 6. The gas turbine engine as recited in claim 5, wherein the ratio of N/R is from 10.9 to 12.0. 7. The gas turbine engine as recited in claim 1, wherein the ratio of N/R is from 10.9 to 12.0. 8. The gas turbine engine as recited in claim 7, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 9. The gas turbine engine as recited in claim 8, wherein the bypass flow passage includes an outlet, the inlet and the outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being approximately 1.3 to 1.55. 10. The gas turbine engine as recited in claim 9, wherein the design pressure ratio is between 1.3 and 1.4. 11. A gas turbine engine comprising: a gear assembly;a bypass flow passage and a core flow passage, the bypass flow passage including an inlet an an outlet, the inlet and the outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being approximately 1.3 to 1.4;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first shaft, the first shaft coupled through the gear assembly with the fan;wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades, the number (N) being 16, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.2, and a ratio of N/R that is from 13.3 to 16.0. 12. The gas turbine engine as recited in claim 11, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 13. The gas turbine engine as recited in claim 12, further comprising a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor. 14. The gas turbine engine as recited in claim 11, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 15. The gas turbine engine as recited in claim 14, wherein the ratio of N/R is from 14.5 to 16.0. 16. The gas turbine engine as recited in claim 11, wherein the ratio of N/R is from 14.5 to 16.0. 17. The gas turbine engine as recited in claim 16, wherein the solidity value (R) at the tips of the fan blades is from 1.0 to 1.1. 18. The gas turbine engine as recited in claim 17, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 19. A gas turbine engine comprising: a gear assembly;a bypass flow passage and a core flow passage, the bypass flow passage including an inlet;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first shaft, the first shaft coupled through the gear assembly with the fan;wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades, the number (N) being 14, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.2, and a ratio of N/R that is from 11.7 to 14.0. 20. The gas turbine engine as recited in claim 19, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 21. The gas turbine engine as recited in claim 19, wherein the bypass flow passage includes an outlet, the inlet and the outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being approximately 1.3 to 1.55. 22. The gas turbine engine as recited in claim 21, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 23. The gas turbine engine as recited in claim 22, wherein the ratio of N/R is from 12.7 to 14.0. 24. The gas turbine engine as recited in claim 19, wherein the ratio of N/R is from 12.7 to 14.0. 25. The gas turbine engine as recited in claim 24, further comprising a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine. 26. The gas turbine engine as recited in claim 25, wherein the bypass flow passage includes an outlet, the inlet and the outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being between 1.3 and 1.4. 27. A gas turbine engine comprising: a bypass flow passage and a core flow passage, the bypass flow passage including an inlet and an outlet which define a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio being between 1.3 and 1.4;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first, the first shaft coupled with the fan; anda second turbine coupled with a second shaft;wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 12 to 16, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 16.0. 28. The gas turbine engine as recited in claim 27, further comprising a gear assembly, wherein the first turbine is coupled with the fan through the first shaft and gear assembly, and further comprising a variable area nozzle, wherein the design pressure ratio is achieved in operation with the visible area nozzle fully open.
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