Efficient, low pressure ratio propulsor for gas turbine engines
A gas turbine engine includes a gear assembly and a bypass flow passage that includes an inlet and an outlet that define a design pressure ratio between 1.3 and 1.55. 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, through the first shaft and the gear assembly, at a lower speed than the first shaft. The fan includes a row of fan blades. The row includes 12-16 (N) fan blades, a solidity value (R) that is from 1.0 to 1.3, and a ratio of N/R that is from 10.0 to 16.
1. A gas turbine engine comprising:
a gear assembly;
a bypass 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 1.55;
a fan arranged within the bypass flow passage;
a first shaft and a second shaft;
a first turbine coupled with the fan through the first shaft and the gear assembly; and
a second turbine coupled with the second shaft;
wherein the fan includes a hub and a row of fan blades that extend radially outwardly 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 1.3, and a ratio of N/R that is from 10.0 to 16.
2. The gas turbine engine as recited in claim 1 , wherein the second turbine is a 2-stage turbine.
3. The gas turbine engine as recited in claim 2 , further comprising a first compressor located between the first turbine and the gear assembly.
4. The gas turbine engine as recited in claim 3 , wherein the first compressor is a 3-stage compressor and is coupled with the first shaft.
5. The gas turbine engine as recited in claim 3 , wherein the number (N) of fan blades is 16, and the solidity value (R) at the tips of the fan blades is from 1.0 to 1.2.
6. The gas turbine engine as recited in claim 3 , wherein the number (N) of fan blades is 14, and the solidity value (R) at the tips of the fan blades is from 1.0 to 1.1.
7. The gas turbine engine as recited in claim 2 , further comprising a variable area nozzle, and wherein each fan blade is fixed in position between the hub and the tip.
8. The gas turbine engine as recited in claim 7 , wherein the design pressure ratio is defined with the variable area nozzle fully open.
9. The gas turbine engine as recited in claim 6 , wherein the fan blades include a carbon-fiber reinforced polymer matrix material.
10. The gas turbine engine as recited in claim 9 , wherein the fan blades further include a three-dimensional fiber structure.
11. The gas turbine engine as recited in claim 10 , wherein the carbon-fiber has an average diameter of 1-100 micrometers.
12. The gas turbine engine as recited in claim 9 , further comprising a case surrounding the fan, the case including a carbon-fiber reinforced polymer matrix material.
13. The gas turbine engine as recited in claim 12 , wherein the case further includes glass fiber, aramid fiber, or combinations thereof.
14. The gas turbine engine as recited in claim 13 , wherein the carbon-fiber reinforced polymer matrix material of the fan blades and the carbon-fiber reinforced polymer matrix material of the case each include a polymer and a fiber, and the carbon-fiber reinforced polymer matrix material of the fan blades is different from the carbon-fiber reinforced polymer matrix material of the case in one or more of the kinds of polymers of the matrices, or the kinds of fibers.
15. The gas turbine engine as recited in claim 3 , wherein the number (N) of fan blades is 12, and the solidity value (R) at the tips of the fan blades is from 1.0 to 1.1.
16. 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;
a first turbine coupled with the fan through the first shaft and the gear assembly; and
a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine;
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 14 to 16, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.3, and a ratio of N/R that is from 11.7 to 16.
17. The gas turbine engine as recited in claim 16 , wherein the number (N) of fan blades is 16, the solidity value (R) at the tips of the fan blades is from 1.0 to 1.2.
18. The gas turbine engine as recited in claim 17 , wherein the bypass flow passage further includes an outlet, 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, and the design pressure ratio is between 1.3 and 1.55.
19. The gas turbine engine as recited in claim 18 , wherein the design pressure ratio is between 1.3 and 1.4.
20. The gas turbine engine as recited in claim 16 , further comprising a first compressor located between the first turbine and the gear assembly, and wherein the number (N) of fan blades is 14, and the solidity value (R) at the tips of the fan blades is from 1.0 to 1.1.
21. The gas turbine engine as recited in claim 20 , further comprising a case surrounding the fan, wherein:
the case includes a carbon-fiber reinforced polymer matrix material;
the fan blades include a carbon-fiber reinforced polymer matrix material with the carbon- fiber having an average diameter of 1-100 micrometers; and
the carbon-fiber reinforced polymer matrix material of the fan blades and the carbon-fiber reinforced polymer matrix material of the case each include a polymer and a fiber, and the carbon- fiber reinforced polymer matrix material of the fan blades is different from the carbon-fiber reinforced polymer matrix material of the case in one or more of the kinds of polymers of the matrices, or the kinds of fibers.
22. A gas turbine engine comprising:
a gear assembly;
a bypass flow passage and a core flow passage;
a fan arranged within the bypass flow passage;
a first shaft and a second shaft;
a first turbine coupled with the fan through the first shaft and the gear assembly; and
a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine;
wherein the fan includes a hub and a row of fan blades that extend radially outwardly from the hub, and the row includes a number (N) of the fan blades that is from 12 to 14 and a solidity value (R) at tips of the fan blades that is from 1.0 to 1.2.
23. The gas turbine engine as recited in claim 22 , wherein a ratio of N/R is from 10.9 to 14.0.
24. The gas turbine engine as recited in claim 23 , wherein the number (N) of fan blades is 12.
25. The gas turbine engine as recited in claim 23 , wherein the number (N) of fan blades is 14, and the ratio of N/R is from 12.7 to 14.0.