Machining method for cooling hole of turbine blade of heavy duty gas turbine with thermal barrier coating of double-ceramic-layer structure

The invention provides a method for processing a cooling hole of a turbine blade of a heavy duty gas turbine with a thermal barrier coating with a double-ceramic-layer structure, which comprises the following steps of: firstly, spraying the thermal barrier coating on the turbine blade by adopting a high velocity oxy-fuel spraying and atmosphere plasma spraying process by adopting a processing method of firstly coating and then punching; the thermal barrier coating structure of the second ceramic layer/the first ceramic layer/the bonding layer/the turbine blade base body is prepared; and then high-efficiency, high-quality and high-precision machining of the turbine blade air film cooling holes is achieved through the combined machining mode of long-pulse laser machining of round holes and ultrafast laser machining of forming holes, the machined cooling holes are uniform in hole diameter and good in consistency, and good industrial application prospects are achieved. 本发明提出一种带有双陶瓷层结构热障涂层的重型燃气轮机透平叶片冷却孔加工方法，通过采用"先