Laminar-to-turbulent boundary layer transition leads to increased friction and aerodynamic heating in hypersonic vehicles. This study proposes a novel methodology to control transition in hypersonic wall-bounded flows by introducing spanwise non-uniform temperature surface
distributions. The proposed control mechanism modulates the spanwise uniform boundary layer flow through the generation of spanwise-periodic streamwise-elongated streaks. Firstly, Direct Numerical Simulations (DNS) are performed at Mach 6 with imposed spanwise nonuniform
wall temperature variations. The wall-prescribed temperature conditions successfully generate streaks of sufficient amplitude that suppress transitional instabilities. Secondly, a wind tunnel experimental implementation of a steady-state temperature control strategy is proposed, incorporating local cooling below a wall of finite thickness. A thermal model is employed to predict the surface temperature with varying cooling temperature inputs below the finite-thickness wall at Mach 6. The thermal analysis validates the methodology to achieve spanwise non-uniform temperature distributions and emphasizes the sensitivity of temperature input and wall-thickness variations. Lastly, to guide future experimental ground tests, restricted to short test duration, prototype models are proposed in order to regulate the local surface
temperature by altering the local heat flux and diffusion within the wall. Specifically, dissimilar materials and thin wall structures are employed to achieve varying levels of heat diffusion and heat flux, respectively. Implementing these approaches generates a higher sensitivity to local surface temperature, thus enabling the passive achievement of spanwise non-uniform temperature distributions, as demonstrated by the thermal model.