Air foil with staggered cooling hole configuration

A gas turbine engine comprising:

• a compressor section;
• a combustor fluidly connected to the compressor section;
• a turbine section fluidly connected to the combustor, the turbine section comprising:
• a high pressure turbine coupled to a high pressure compressor of the compressor section via a shaft;
• a low pressure turbine;
• and
• wherein the high pressure turbine includes a turbine disk with a plurality of turbine blades secured thereto each of the plurality of turbine blades, comprising:
• an airfoil, the airfoil having a leading edge, a pressure side, a suction side and a trailing edge;
• a plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities;
• the leading edge cavity extending towards the suction side;
• a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage;
• and
• a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects the leading edge cavity at a point forward of a line parallel to a pull angle or edge of the leading edge feed passage and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects a vertex of the leading edge feed passage.

A method for forming an airfoil of a turbine blade, comprising:

• forming a plurality of internal cooling cavities in the airfoil, the plurality of internal cooling cavities including a leading edge cavity, a leading edge feed passage, pressure side cooling passages, suction side cooling passages and main body cavities;
• the leading edge cavity extending towards the suction side, the airfoil having a leading edge, a pressure side, a suction side and a trailing edge;
• forming a first crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage;
• and
• forming a second crossover row of cooling passages providing fluid communication between the leading edge cavity and the leading edge feed passage, a centerline of the first crossover row of cooling passages is located closer to the pressure side than a centerline of the second crossover row of cooling passages and the centerline of the second crossover row of cooling passages is located closer to the suction side than the centerline of the first crossover row of cooling passages, and wherein the second crossover row of cooling passages are radially staggered relative to the first crossover row of cooling passages and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects the leading edge cavity at a point forward of a line parallel to a pull angle or edge of the leading edge feed passage and the centerline of the first crossover row of cooling passages and the centerline of the second crossover row of cooling passages intersects a vertex of the leading edge feed passage.