Efficient, low pressure ratio propulsor for gas turbine engines

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 within the bypass flow passage;
• a first shaft and a second shaft, wherein the first shaft and the second shaft are concentric, the first shaft being an inner shaft and the second shaft being an outer shaft;
• a first turbine coupled with the first shaft, the first shaft coupled with the fan, wherein the first turbine is a 5-stage turbine;
• a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor;
• 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, the number (N) being 18, 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 15.0 to 18.0.

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, and 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 being approximately 1.3 to 1.55;
• a fan arranged within the bypass flow passage;
• a first shaft and a second shaft, wherein the first shaft and the second shaft are concentric, the first shaft being an inner shaft and the second shaft being an outer shaft;
• a first turbine coupled with the first shaft, the first shaft coupled with the fan;
• a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor;
• 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, the number (N) being 18, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.1, and a ratio of N/R that is from 16.4 to 18.0.

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, and 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 being approximately 1.55;
• a fan within the bypass flow passage;
• a first shaft and a second shaft, wherein the first shaft and the second shaft are concentric, the first shaft being an inner shaft and the second shaft being an outer shaft;
• a first turbine coupled with the first shaft, the first shaft coupled with the fan;
• a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor;
• 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, the number (N) being 18, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.1, and a ratio of N/R that is from 16.4 to 18.0.