Efficient, low pressure ratio propulsor for gas turbine engines

Abstract

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 has a bypass ratio of greater than 8.5. 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.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a gear assembly;   a bypass flow passage;   a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5;   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, wherein the first compressor is a 3-stage compressor;   a second turbine coupled with the second shaft; and   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 20.   
     
     
         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 , 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, and the design pressure ratio is less than 1.55. 
     
     
         4 . The gas turbine engine as recited in  claim 2 , 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, and the design pressure ratio is from 1.3 to 1.4. 
     
     
         5 . The gas turbine engine as recited in  claim 3 , further comprising a case surrounding the fan, and the fan is a single fan stage. 
     
     
         6 . The gas turbine engine as recited in  claim 5 , wherein each of the fan blades is fixed in position between the hub and the tip. 
     
     
         7 . The gas turbine engine as recited in  claim 6 , wherein the fan blades are formed of a fiber-reinforced polymer matrix material. 
     
     
         8 . The gas turbine engine as recited in  claim 6 , wherein the first turbine is a 5-stage turbine. 
     
     
         9 . A gas turbine engine comprising:
 a gear assembly;   a bypass flow passage;   a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5;   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;   a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine; and   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 20.   
     
     
         10 . The gas turbine engine as recited in  claim 9 , 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, and the design pressure ratio is less than 1.55. 
     
     
         11 . The gas turbine engine as recited in  claim 9 , 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, and the design pressure ratio is from 1.3 to 1.4. 
     
     
         12 . The gas turbine engine as recited in  claim 10 , further comprising a case surrounding the fan, and the fan is a single fan stage. 
     
     
         13 . 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. 
     
     
         14 . The gas turbine engine as recited in  claim 13 , wherein 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 turbine is a 5-stage turbine. 
     
     
         16 . The gas turbine engine as recited in  claim 13 , wherein the first compressor is a 3-stage compressor. 
     
     
         17 . A gas turbine engine comprising:
 a gear assembly;   a bypass flow passage;   a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5;   a first shaft and a second shaft;   a first turbine coupled through the gear assembly to the fan, wherein the first turbine is a 5-stage turbine;   a first compressor coupled with the first shaft;   a second turbine coupled with the second shaft; and   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 20.   
     
     
         18 . The gas turbine engine as recited in  claim 17 , wherein the first compressor is a 3-stage compressor. 
     
     
         19 . The gas turbine engine as recited in  claim 18 , wherein the second turbine is a 2-stage turbine. 
     
     
         20 . The gas turbine engine as recited in  claim 19 , 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, and the design pressure ratio is less than 1.55. 
     
     
         21 . The gas turbine engine as recited in  claim 19 , 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, and the design pressure ratio is from 1.3 to 1.4. 
     
     
         22 . The gas turbine engine as recited in  claim 20 , further comprising a case surrounding the fan, and the fan is a single fan stage. 
     
     
         23 . The gas turbine engine as recited in  claim 22 , wherein each of the fan blades is fixed in position between the hub and the tip. 
     
     
         24 . The gas turbine engine as recited in  23 , wherein the fan blades are formed of a fiber-reinforced polymer matrix material.