Efficient, low pressure ratio propulsor for gas turbine engines
A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.
1. A gas turbine engine comprising:
a gear assembly;
a bypass flow passage;
a fan located upstream of the bypass flow passage;
a first shaft and a second shaft;
a first turbine coupled through the gear assembly to the fan;
a first compressor coupled with the first shaft; and
a second turbine coupled with the 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 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 16.0.
2. The gas turbine engine as recited in claim 1 , wherein the number of fan blades is 20 and N/R is 15.4.
3. The gas turbine engine as recited in claim 1 , wherein the number of fan blades is 18 or 20.
4. The gas turbine engine as recited in claim 3 , wherein the bypass flow passage includes an inlet and an 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 is from 1.2 to 1.4.
5. The gas turbine engine as recited in claim 3 , wherein the bypass flow passage includes an inlet and an 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 is from 1.3 to 1.4.
6. The gas turbine engine as recited in claim 1 , further comprising a case surrounding the fan, and the fan is a single fan stage.
7. The gas turbine engine as recited in claim 6 , wherein each of the fan blades is fixed in position between the hub and the tip.
8. The gas turbine engine as recited in claim 7 , wherein the fan blades are formed of a fiber-reinforced polymer matrix material.
9. The gas turbine engine as recited in claim 6 , wherein the first compressor is a 3-stage compressor.
10. The gas turbine engine as recited in claim 7 , wherein the first turbine is a 5-stage turbine.
11. The gas turbine engine as recited in claim 10 , further comprising a second turbine coupled with the second shaft, and the second turbine is a 2-stage turbine.
12. A gas turbine engine comprising:
a gear assembly;
a bypass flow passage;
a fan located upstream of the bypass flow passage;
a first shaft and a second shaft;
a first turbine coupled through the gear assembly to the fan;
a first compressor coupled with the first shaft; and
a second turbine coupled with the 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, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 13.8 to 16.4.
13. The gas turbine engine as recited in claim 12 , wherein the bypass flow passage includes an inlet and an 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 is from 1.2 to 1.4.
14. The gas turbine engine as recited in claim 12 , wherein each of the fan blades is fixed in position between the hub and the tip, further comprising a case surrounding the fan, the fan is a single fan stage, and the fan blades are formed of a fiber-reinforced polymer matrix material.
15. The gas turbine engine as recited in claim 14 , wherein the first compressor is a 3-stage compressor and the first turbine is a 5-stage turbine, further comprising a second turbine coupled with the second shaft, and the second turbine is a 2-stage turbine.
16. A gas turbine engine comprising:
a gear assembly;
a bypass flow passage;
a fan located upstream of the bypass flow passage;
a first shaft and a second shaft;
a first turbine coupled through the gear assembly to the fan;
a first compressor coupled with the first shaft; and
a second turbine coupled with the 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 18 or 20 and a solidity value (R) at tips of the fan blades, where for 20 fan blades the solidity value (R) is 1.3, 1.1, or 1.0 such that a ratio of N/R is, respectively, 15.4, 18.2, or 20, or where for 18 fan blades the solidity value (R) is 1.3, 1.2, or 1.1 such that the ratio of N/R is, respectively, 13.8, 15, or 16.4.
17. The gas turbine engine as recited in claim 16 , wherein the number of fan blades is 20.
18. The gas turbine engine as recited in claim 16 , wherein the number of fan blades is 18.
19. A gas turbine engine comprising:
a bypass flow passage;
a fan located upstream of the bypass flow passage;
a first turbine;
a first compressor; and
means for coupling an extraction of energy from the first turbine to the first compressor and 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 that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 16.
20. The gas turbine engine as recited in claim 19 , wherein the means for coupling an extraction of energy comprises a first shaft, wherein the first turbine is coupled with the first shaft, and the first shaft is coupled with the first compressor and the fan is coupled to a gear assembly which is coupled to the first shaft.