Turbofan arrangement with blade channel variations

1. A fan section for a gas turbine engine comprising:

a rotor hub defining an axis;

an array of airfoils circumferentially spaced about the rotor hub, each of the airfoils including pressure and suction sides between a leading edge and a trailing edge and extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, facing pressure and suction sides of adjacent airfoils defining a channel in a chordwise direction having a width between the facing pressure and suction sides at a given span position of the adjacent airfoils;

wherein the width at each pressure side location along the channel is defined as a minimum distance to a location along the suction side, the width diverging without converging along the channel for at least some of the span positions, and the width converging and diverging along the channel for at least some span positions greater than 5% span and less than half of the span positions; and

wherein the array of airfoils has a solidity defined by a ratio of an airfoil chord over a circumferential pitch, the solidity being between 1.6 and 2.5 for each of the span positions in which the width converges along the channel.

2. The fan section as set forth in claim 1 , wherein the width diverges without converging along the channel at span positions greater than 16% span.

3. The fan section as set forth in claim 2 , wherein the width converges and diverges along the channel at span positions greater than or equal to 10% span.

4. The fan section as set forth in claim 1 , wherein the width diverges without converging at span positions from 100% span to less than or equal to 90% span.

5. The fan section as set forth in claim 4 , wherein the solidity at the tip of each of the array of airfoils is less than or equal to 1.2.

6. The fan section as set forth in claim 4 , wherein the width converges and diverges for less than or equal to 20% of the span positions.

7. The fan section as set forth in claim 4 , wherein a ratio of the width to the solidity at each span position is greater than or equal to 0.50.

8. The fan section as set forth in claim 7 , wherein the solidity at each span position is greater than or equal to 0.8.

9. The fan section as set forth in claim 8 , wherein the array of airfoils includes 20 or fewer airfoils.

10. The fan section as set forth in claim 4 , wherein flow through the channel at span positions where the width converges and diverges along the channel corresponds to a leading edge relative mach number less than or equal to 0.8 Mach at cruise.

11. The fan section as set forth in claim 1 , wherein a stagger angle of each of the array of airfoils relative to the axis is less than or equal to 16 degrees at each of the span positions in which the width converges and diverges along the channel.

12. The fan section as set forth in claim 11 , wherein the width converges and diverges for less than or equal to 20% of the span positions, and the width diverges without converging at span positions from 100% span to less than or equal to 80% span.

13. The fan section as set forth in claim 12 , wherein the width converges along the channel at a location spaced a distance from an inlet of the channel, the distance being greater than a radius defined by the leading edge at the same span position.

14. The fan section as set forth in claim 1 , wherein the solidity at the 0% span position is greater than or equal to 2.3.

15. A gas turbine engine comprising:

a combustor section arranged between a compressor section and a turbine section;

a fan section having a rotor hub and an array of airfoils circumferentially spaced about the rotor hub defining an axis;

wherein each of the array of airfoils includes pressure and suction sides and extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, facing pressure and suction sides of adjacent airfoils defining a channel in a chordwise direction having a width between the facing pressure and suction sides at a given span position of the adjacent airfoils; and

wherein the width at each pressure side location along the channel is defined as a minimum distance to a location along the suction side, the width converging along the channel for at least some span positions greater than 5% span and less than half of the span positions; and

wherein a stagger angle of each airfoil of the array of airfoils relative to the axis is less than or equal to 16 degrees at span positions converging and diverging along the channel.

16. The gas turbine engine as set forth in claim 15 , wherein the width converges and diverges for less than or equal to 20% of the span positions.

17. The gas turbine engine as set forth in claim 15 , wherein the width diverges without converging at span positions from 100% span to less than or equal to 90% span.

18. The gas turbine engine as set forth in claim 17 , wherein each of the array of airfoils has a solidity defined by a ratio of an airfoil chord over a circumferential pitch, and a ratio of the width to the solidity at each span position is greater than or equal to 0.50.

19. The gas turbine engine as set forth in claim 18 , wherein the solidity at each span position is greater than or equal to 0.8.

20. The gas turbine engine as set forth in claim 19 , wherein the array of airfoils includes 20 or fewer airfoils.