This research addresses the impact of shock wave-boundary layer interactions (SWBLI) at greater angles of attack over a NACA 64-110 airfoil at supersonic speeds. Aerodynamic performance is largely determined by the intricate relationship between shock waves and the boundary layer, especially in supersonic flight regimes. Higher angles of attack cause more intense shock forms in the flow over the airfoil, which can cause noticeable thickening of the boundary layer, separation of the flow, and possible reattachment zones. The airfoil’s lift, drag, and overall stability are all greatly impacted by these events. To describe the nature of SWBLI under these circumstances, computational and experimental evaluations were carried out, with a focus on the boundary layer separation, shock-wave behavior, and pressure distribution. The findings demonstrate that raising the angle of attack causes the shock wave to become stronger and exacerbates boundary layer separation, which in turn causes a notable decrease in aerodynamic efficiency. In particular, for high-speed aircraft or fighter aircraft operating at crucial angles of attack, the research findings enhance our understanding of SWBLI for airfoil design and optimization in supersonic applications.