The invention belongs to the technical field of
heat transfer and cooling of hot end components of
gas turbines, and relates to an air film cooling composite hole structure for a
turbine blade. The
turbine blade comprises an air film hole plate and a
thermal barrier coating arranged on the air film hole plate, a plurality of composite air film holes are formed in the air film hole plate, and a
dumbbell-shaped groove is formed in the
thermal barrier coating; and each composite air film hole comprises a spanwise expansion section and a straight hole section which communicate with each other, the straight hole sections communicate with an inner
cooling channel of the
turbine blade, and outlets of the spanwise expansion sections communicate with the
dumbbell-shaped groove. The guiding effect of the
dumbbell-shaped groove enables cooling air flow to be diffused in the groove in the transverse direction and the longitudinal direction, the
jet flow momentum is weakened, and the spanwise covering capacity is greatly improved. In addition, the dumbbell-shaped groove changes the rotating direction of a
kidney-shaped vortex pair, and the vortex pair develops towards the two sides of the groove under the
compression action of main flow, so that the
cold air jet flow is better attached to the wall surface; and the air film holes comprise the spanwise expansion sections at the outlets, so that the
jet flow momentum can be further reduced, the
diffusion of an air film in the groove is accelerated, and an excellent
cooling effect is achieved.