Blade for a gas turbine engine

1. A blade for a gas turbine engine comprising an asymmetrical blade root, wherein

the blade root is asymmetrical about a mid-plane that bisects the blade root perpendicular to a longitudinal direction of a slot to which the blade root is configured to fit,

a depth of the blade root is the same on opposite ends of the blade root in the longitudinal direction, and

a portion of a lower surface includes a tapered increase of a depth of the blade root along the longitudinal direction.

2. The blade as claimed in claim 1 , wherein the blade root is asymmetrical about a mid-plane parallel with the longitudinal direction of the slot to which the blade root is configured to fit.

3. The blade as claimed in claim 1 , wherein the blade root is asymmetrical about at least a mid-plane parallel with and including an engine axis.

4. The blade as claimed in claim 1 , wherein the blade root is asymmetrical about at least a mid-plane perpendicular to an engine axis.

5. The blade as claimed in claim 1 , wherein the blade root has an outer profile which is asymmetrical about the mid-plane.

6. The blade as claimed in claim 1 , wherein the lower surface is asymmetrical about the mid-plane.

7. The blade as claimed in claim 1 , wherein the blade root has first and second corresponding regions on opposite sides of the mid-plane which have different average root depths.

8. The blade as claimed in claim 7 , wherein

the first region is configured to experience a higher stress than the second region; and

the first region has an average root depth that is greater than that of the second region.

9. The blade as claimed in claim 1 , wherein there is an asymmetrical variation in blade root depth of at least 5%, at least 7%, at least 10%, at least 12% or at least 15%.

10. The blade as claimed in claim 1 , wherein the blade is a compressor blade, a turbine blade or a fan blade.

11. A rotor disc assembly comprising a disc and at least one blade as claimed in claim 1 .

12. A gas turbine engine comprising the blade as claimed in claim 1 .

13. A gas turbine engine comprising the rotor disc assembly as claimed in claim 11 .

14. The blade according to claim 1 , wherein the tapered increase is continuous.

15. The blade according to claim 1 , wherein the tapered increase is continuous on opposite sides of the mid-plane.