Abstract <p>To mitigate the adverse effects of aerodynamic heating on hypersonic vehicles, a tangential supersonic cooling film is typically used. This study investigates the impacts of the cooling film’s Mach number (<i>Mj</i>) on the flow field structure. Two supersonic cooling film configurations, with the Mach numbers of 2.0 and 2.3, were designed and tested in a supersonic wind tunnel at the freestream Mach number M = 3.8. The flow field structure was obtained using nanotracer-based planar laser scattering (NPLS), and the wall pressure were derived using an experimentally validated numerical simulation method. The results demonstrate that in the mixing layer at <i>Mj</i> = 2.0 instability between the freestream and the supersonic cooling film develops earlier than that at <i>Mj</i> = 2.3, occurring under an identical ratio of the static pressure (RSP) conditions. On convex surfaces, as the radius of curvature decreases, the influence of the cooling film’s M on Δ<i>P</i>/<i>P</i>_in diminishes; conversely, on concave surfaces, as the radius of curvature decreases, the influence of the cooling film’s M on Δ<i>P</i>/<i>P</i>_in increases. Beyond <i>x =</i> 240 mm, the development over curved surfaces becomes pronounced, and the static pressure of the supersonic cooling film has minimum impact on the wall pressure. Variation in the wall pressure is affected by both the coverage length and the curvature of the supersonic cooling film, and for the cooling film the higher <i>Mj</i> achieves a longer coverage length.</p>