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Experiments were performed to determine the heat transfer to the suction surface of two film cooled first stage turbine vanes using two-row discrete hole cooling. Evaluation of the heat transfer to the film cooled vanes was made in a turbulent environment downstream of an aircraft turbojet engine combustor at moderately high primary gas stream temperatures and pressures. Results of these experiments provided quantitative data on the adiabatic film effectiveness for a number of blowing rates, and show decreased effectiveness levels when compared to measured values for a continuous slot in a wind tunnel environment. In addition the external heat transfer coefficients were obtained based on the difference between the film cooled adiabatic wall temperature and the wall temperature. The presence of film cooling holes increased the heat-transfer coefficients even without blowing. Further increases in the heat-transfer coefficients were measured in the presence of blowing with the largest increases in the region nearest the ejection.

A method of superposition which predicts the cumulative effect of film cooling has been described in the literature by Sellers. This method predicts the magnitude and extent of film cooling downstream of multiple injection stations by accounting for the effectiveness level from each upstream station individually. Sellers compared the method with results from multiple slots and good agreement was demonstrated. Because of the simplicity of the method it was proposed as a method to predict the cumulative effect of film cooling from rows of holes. To substantiate this, experiments were conducted with film cooling schemes on flat plates and airfoils. The results of these experiments demonstrate that the method is satisfactory for design of film cooled airfoils in the required range of geometries and coolant flows.

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