Determination of the heat fluxes into the walls for compressible gas flows is an important part of the experimental work. It's critical for specific conditions under low Reynolds numbers and moderate accelerations corresponded to not only supersonic nozzles but also in many technological flow situations including swept wings and turbine blades. This is explained by possibility of change of the flow regime−both laminar-turbulent transition and laminarization should be predicted. Nowadays modern approaches based on infrared thermography with high spatial resolution allow experimental recording of surface temperatures with significant longitudinal gradients of heat flux characterized to such conditions. However, temperature distribution is insufficient to accurately describe flow evolution. Heat transfer parameters as heat flux, heat transfer coefficients and adiabatic wall temperatures should be reconstructed. In present paper the inverse heat transfer procedure is used to calculate the change of wall heat flux in time from the measured surface temperatures on the side wall of slot supersonic channel. Further, the transient heat transfer method was implemented for recovering the heat transfer coefficients and adiabatic wall temperature. The experimental channel slot geometry allowed to study process of laminarization. During the measurements no active heating or cooling of the gas or channel was applied, the surface temperature variation was relatively small with respect to the adiabatic wall temperature and driving temperature difference did not affect the boundary layer characteristics. The laminarization zones, level of partial laminarization were identified as function of heat transfer rate reduction scaled to full developed turbulence values.