Experimental investigations of showerhead film cooling on the leading edge of a turbine blade

The turbine components of a gas turbine need extensive cooling to withstand the high heat loads generated by the flow of the hot combustion gases. The leading edge of a turbine blade is one of the areas that faces the hottest gas flow conditions and is thus one of the most critical areas to be cooled. The stagnation region of turbine airfoils is generally cooled with showerhead cooling scheme, which consist in injecting coolant onto the stagnation region through closely spaced rows of film cooling holes. This provides convective cooling inside the holes and produces a layer of coolant film on the external surface. The objective of this study is to investigate experimentally the influence of mainstream incidence on showerhead film cooling. The film cooling effectiveness and heat transfer coefficient were thus measured on a series of symmetric blunt bodies with different showerhead configurations simulating the leading edge of turbine airfoils. Tests have been performed at 0° on-design and 5° off-design mainstream incidences at an approach Mach number of 0.26. Besides the mainstream incidence angle, the investigated parameters were the blowing ratio, the wedge angle, the row location, the spanwise pitch between the holes and the hole shape. In addition to the film cooling measurements, baseline heat transfer tests were performed on test models without film cooling holes in order to evaluate the film cooling performances. To characterize the mainflow conditions and the velocity profile on the test models, aerodynamic measurements have also been carried out. The heat transfer measurements have been performed using the transient liquid crystal technique. The transient experiments were generated by a pre-conditioning system and a rapid insertion mechanism. The data analysis was adapted to measurements on highly curved surfaces using the approximate analytical solution of the heat conduction into a solid cylinder with convective boundary conditions on the surface. Time-varying adiabatic wall temperatures were accounted for using Duhamel's superposition theorem. The first tested showerhead geometry was a triangular blunt body representative of a turbine rotor blade showerhead with a circular leading edge, a 21.1° wedge angle and two film cooling rows. Because of jet lift-off, the film cooling performances of this two-row configuration were relatively low and were almost unaffected by the increase of blowing ratio from 1.5 to 3.9. The heat transfer coefficient was found to be greatly increased at the row location in comparison to the case without film cooling. At 5° off-design incidence angle, the stagnation line moves passed the pressure side row. The two rows were then located on the same side of the stagnation line and were both blowing towards the suction side. On the suction side the film cooling protection was thus very good due to film accumulation but the pressure side was left almost unprotected resulting in highly asymmetric film cooling performances. Two-dimensional FEM calculations of the metal temperature in this second geometry have been performed at engine representative conditions for design and off design mainstream incidences. It was found that the contribution of the film cooling to cool the leading edge region was rather small in comparison to the contribution of the internal convective cooling. Showerhead film cooling was however required to effectively cool the stagnation region and to obtain a uniform surface temperature. The second tested showerhead geometry was similar to the first geometry but with an increased wedge angle of 55° and cooling rows located further downstream. The film cooling performances were found to decrease when the blowing ratio was increased from 1.2 to 3.0 because of jet lift-off. At middle blowing ratio (2.0), the film cooling effectiveness had similar values as for the first geometry. However, the heat transfer increase due to the presence of the film cooling jets was much higher for the second geometry because of higher local mainstream velocities at the injection locations. The resulting film cooling performances were thus lower for the second geometry than for the first geometry. Nevertheless, the larger wedge angle and the cooling rows located further downstream assured that the stagnation line stayed in-between the two cooling rows even at 5° off-design incidence. The second geometry had thus the advantage to be much less sensitive to incidence angle and, unlike the first geometry, provided very similar film cooling performances at 5° and at 0° incidence. The use of shaped holes was found to greatly improve the film cooling performances of the second geometry due to a better lateral coverage and a reduced tendency of jet lift-off