Compressor rotor airfoil

A compressor blade for a gas turbine engine, the compressor blade comprising:

• an airfoil having opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending along a span in a spanwise direction from a root to a tip;
• a leading edge dihedral angle defined at a point on the leading edge between a tangent to the airfoil and a vertical in a radial direction relative to an axial axis of the gas turbine engine, a positive dihedral angle leans the airfoil in a direction of rotation;
• and
• first, second and third regions of the airfoil, the first region extending from the root toward the tip, the third region extending from the tip toward the root, and the second region extending between the first and third regions, the leading edge dihedral angle generally decreases within the first region in the spanwise direction, generally increases within the second region in the spanwise direction, and generally decreases within the third region in the spanwise direction, the leading edge dihedral angle remains positive along the span from the root to the tip.

A method of designing a compressor blade for a gas turbine engine, the compressor blade rotatable about an axial axis in a direction of rotation, the method comprising:

• reducing a rub angle between the compressor blade and a casing surrounding the compressor blade by:
• increasing a leading edge dihedral angle over a region of the compressor blade in a spanwise direction from a root to a tip, the leading edge dihedral angle defined at a point on a leading edge between a tangent to the compressor blade and a vertical in a radial direction relative to the axial axis;
• and
• decreasing the leading edge dihedral angle from the region to the tip of the compressor blade in the spanwise direction.