Efficient, low pressure ratio propulsor for gas turbine engines

Patent metadata

Field	Value
Publication number	US-2017191449-A1
Application number	US-201615252689-A
Country	US
Kind code	A1
Filing date	Aug 31, 2016
Priority date	Jul 5, 2011
Publication date	Jul 6, 2017
Grant date	—

Abstract

A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged at the inlet. A first turbine is coupled with a first shaft such that rotation of the first turbine will drive the fan. A first compressor is coupled with the first shaft and includes three stages, and a second turbine is coupled with a second shaft and includes two stages. The fan includes a row of 16 (N) fan blades that has a solidity value (R) that is from 1.0 to 1.2 and a ratio of N/R that is from 13.3 to 16.0.

First claim

What is claimed is:

1 . A gas turbine engine comprising:
a bypass flow passage and a core flow passage, the bypass flow passage including an inlet and defining a bypass ratio in a range of approximately 8.5 to 13.5 with regard to flow through the bypass flow passage and flow through the core flow passage; a fan arranged at the inlet of the bypass flow passage; a first shaft and a second shaft mounted for rotation about an engine central longitudinal axis; a first turbine coupled with the first shaft such that rotation of the first turbine is configured to drive the fan; a first compressor coupled with the first shaft, the first compressor including three stages; and a second turbine coupled with the second shaft, the second turbine including two stages; 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) that is from 1.0 to 1.2, and a ratio of N/R that is from 13.3 to 16.0.

2 . 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.

3 . The gas turbine engine as recited in claim 2 , wherein the solidity value (R) is from 1.0 to 1.1.

4 . The gas turbine engine as recited in claim 3 , further including a gear assembly, wherein the first turbine is configured to drive the fan, through the first shaft and the gear assembly, at a lower speed than the first shaft.

5 . The gas turbine engine as recited in claim 3 , wherein the ratio of N/R is from 14.5 to 16.0.

6 . The gas turbine engine as recited in claim 3 , wherein the design pressure ratio is approximately 1.55.

7 . The gas turbine engine as recited in claim 1 , wherein the solidity value (R) is from 1.1 to 1.2.

8 . The gas turbine engine as recited in claim 7 , wherein the ratio of N/R is from 13.3 to 14.5.

9 . The gas turbine engine as recited in claim 8 , wherein the design pressure ratio is approximately 1.55.

10 . The gas turbine engine as recited in claim 9 , wherein the first shaft and the second shaft are concentric and rotate about the engine central longitudinal axis, the first shaft being an inner shaft and the second shaft being an outer shaft.

11 . The gas turbine engine as recited in claim 1 , wherein each of the fan blades is fixed in position between the hub and the tip.

12 . A gas turbine engine comprising:
a bypass flow passage and a core flow passage, the bypass flow passage defining a bypass ratio in a range of approximately 8.5 to 13.5 with regard to flow through the bypass flow passage and flow through the core flow passage; a fan arranged in the bypass flow passage; a first shaft and a second shaft mounted for rotation about an engine central longitudinal axis; a first turbine coupled with the first shaft such that rotation of the first turbine is configured to drive the fan; a second turbine coupled with the second shaft, the second turbine including two stages; and a first compressor coupled with the first shaft, the first compressor including three stages; 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, and a solidity value (R) that is from 1.0 to 1.2.

13 . The gas turbine engine as recited in claim 12 , wherein the first shaft and the second shaft are concentric and rotate about the engine central longitudinal axis, the first shaft being an inner shaft and the second shaft being an outer shaft.

14 . The gas turbine engine as recited in claim 13 , further comprising a ratio of N/R that is from 13.3 to 16.0.

15 . The gas turbine engine as recited in claim 14 , wherein the solidity value (R) is from 1.0 to 1.1.

16 . The gas turbine engine as recited in claim 15 , further including a gear assembly, wherein the first turbine is configured to drive the fan, through the first shaft and the gear assembly, at a lower speed than the first shaft.

17 . The gas turbine engine as recited in claim 15 , wherein the ratio of N/R is from 14.5 to 16.0.

18 . The gas turbine engine as recited in claim 17 , wherein the bypass flow passage further includes an inlet and an outlet, and the fan is located at the inlet, and the inlet and the outlet 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 approximately 1.3 to 1.55.

19 . The gas turbine engine as recited in claim 18 , wherein the outlet has a cross-sectional area, and further comprising a variable area nozzle, the variable area nozzle being operative to change the cross-sectional area of the outlet to thereby control the design pressure ratio, and wherein the variable area nozzle is configured to achieve the design pressure ratio with the variable area nozzle fully open.

20 . The gas turbine engine as recited in claim 18 , wherein the design pressure ratio is approximately 1.55.

21 . The gas turbine engine as recited in claim 14 , wherein the solidity value (R) is from 1.1 to 1.2.

22 . The gas turbine engine as recited in claim 21 , wherein the ratio of N/R is from 13.3 to 14.5.

23 . The gas turbine engine as recited in claim 22 , wherein the bypass flow passage further includes an inlet and an outlet, and the fan is located at the inlet, and the inlet and the outlet 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 approximately 1.3 to 1.55.

24 . The gas turbine engine as recited in claim 23 , wherein the design pressure ratio is approximately 1.55.

25 . A propulsor for use in a gas turbine engine, the propulsor comprising:
a rotor including a row of fan blades extending radially outwardly from a hub; wherein the row of fan blades includes a number (N) of said fan blades that is 16, a solidity value as a ratio (R) from 1.0 to 1.2, and a ratio of N/R from 13.3 to 16.0.

26 . The propulsor as recited in claim 25 , wherein the solidity value (R) is from 1.1 to 1.2.

27 . The propulsor as recited in claim 26 , wherein the ratio of N/R is from 13.3 to 14.5.

28 . The propulsor as recited in claim 25 , wherein the solidity value (R) is from 1.0 to 1.1.

29 . The propulsor as recited in claim 28 , wherein the ratio of N/R is from 14.5 to 16.0.

30 . The propulsor as recited in claim 25 , wherein each of the fan blades is fixed in position between the hub and the tip.

Assignees

• United Technologies Corp

Inventors

• Gallagher Edward J
• Monzon Byron R

Classifications

• F02K1/06

  Varying effective area of jet pipe or nozzle (by using fluid jets to influence the jet flow F02K1/30) · CPC title

• F05D2220/327

  to drive shrouded, high solidity propeller · CPC title

• F05D2260/4031

  as in toothed gearing · CPC title

• F02K3/06 Primary

  with front fan · CPC title

• F01D17/14

  varying effective cross-sectional area of nozzles or guide conduits · CPC title