Compressor blade for gas turbine engine

摘要

A compressor blade for a gas turbine engine includes a root configured to engage a hub of the gas turbine engine, and an airfoil radially extending a distance from the root to a tip. The airfoil includes a suction side, a pressure side, a leading edge connecting the suction and pressure sides, and a trailing edge connecting the suction and pressure sides opposite the leading edge. The distance that the airfoil extends from the root to the tip may be divided into a plurality of radially adjacent regions. At least one, but not all, of the plurality of radially adjacent regions may have a substantially constant thickness.

主权项

1. A compressor blade for a gas turbine engine, the compressor blade comprising:
a root configured to engage a hub of the gas turbine engine; and an airfoil having a suction side, a pressure side, a leading edge connecting the suction and pressure sides, and a trailing edge connecting the suction and pressure sides opposite the leading edge, a base of the airfoil being disposed adjacent to the root, the airfoil extending a distance along a radial direction from the base to a tip of the airfoil, wherein the distance that the airfoil extends from the base to the tip is divided into a first plurality of radially adjacent regions, a second plurality of radially adjacent regions includes all regions of the first plurality of radially adjacent regions except for a region including the tip and a region including the base, at least one region, but not all regions, of the second plurality of radially adjacent regions has a thickness that varies by less than 5 percent throughout the at least one region, the thickness of the at least one region of the second plurality of radially adjacent regions extends along a line perpendicular to a chord of the airfoil at a general midpoint of the airfoil, and a thickness of the airfoil decreases by 55-65% throughout a span of the airfoil.