Abstract To keep metal temperature at allowable levels in modern gas turbine engines, both internal and external (film) cooling is necessary, which is represented by an overall cooling effectiveness. Previous studies measured overall cooling effectiveness for first-stage vanes in low-speed experiments, where it is appropriate to use the mainstream gas temperature as the hot fluid reference temperature. For better application to gas turbine operating conditions, the recovery temperature is applied. This arises due to the viscous dissipation within the boundary layer and accounts for local Mach number variation and the boundary layer state. To understand the influence of Mach number on overall effectiveness, a two-cavity fed metallic film-cooled vane with cylindrical showerhead, pressure side, and suction side film cooling rows was experimentally tested in a high-speed linear cascade at exit Mach numbers ranging from 0.7 to 1.1 and varying blowing ratios. Vane surface temperature measurements were obtained with infrared thermography. In addition to studies with both cavities flowing, a study of the superposition of overall cooling was performed for individual cavities. This study showed that cooling effectiveness did not significantly change with Mach number and that the principle of superposition for overall effectiveness can be applied to transonic film-cooled vanes. Additionally, the appropriate reference temperature for overall effectiveness was explored. While utilizing the recovery temperature as the reference temperature resulted in the collapse of overall effectiveness at subsonic Mach number, the inlet total temperature is a better reference temperature in cases where shocks are present.