In order to enter orbit, a launch vehicle requires a substantial energy input, approximately 32000 kJ. Upon re-entry back into the planet’s atmosphere at hypersonic velocities, this energy is converted into kinetic energy, which results in significant aerodynamic heating due to atmospheric friction. To shield the vehicle from this heat, thermal management systems, such as insulators or cooling systems are employed. The most intense heating occurs at stagnation points, such as the nosecap and wing leading edges, necessitating meticulous thermal management system design at these points. This study focuses on thermal management of re-entry vehicle’s nosecap. A python code called USHMA was developed to estimate heat flux levels using engineering correlations. Thermal analysis was carried out to identify temperatures surpassing safe thresholds, prompting parametric studies to comprehend heat transfer mechanisms. A thermal management strategy was devised, incorporating multiple radiation shields, spacers, internal coatings, and material selection based on analysis. The design was validated ensuring the safe operation of all components.