This disclosure relates to an airfoil for a gas turbine engine. In particular, the disclosure relates to an impingement cooling configuration for the airfoil.
Airfoils, particularly those used in a hot section of a gas turbine engine, incorporate internal cooling features. One type of configuration utilizes a radially extending cooling passage that communicates cooling fluid through holes in an impingement wall within the airfoil to impinge upon an interior surface of a leading edge wall. The leading edge wall provides the leading edge of the airfoil.
Current airfoil manufacturing techniques limit possible impingement cooling configurations. Typically, the airfoil is cast within a mold having at least first and second portions secured to one another to define an exterior airfoil surface. The core structure used to form the impingement holes and cooling passages must be retained between the mold portions, which limit the location and configuration of the impingement holes. Typically, the impingement wall can only be formed with a single row of holes extending in a radial direction and centered along an airfoil thickness direction.
USPTO Backgrounds
USPTO Backgrounds