Abstract: Flow separation occurs over the compression corners generated by deflected control surfaces on hypersonic vehicles. This phenomenon is coupled with shock/shock interactions, shock/boundary layer interactions and real gas effects. In order to design control surfaces of hypersonic vehicles and insure flight stability, it is important to predict this complex flowfield features on hypersonic vehicles. In this paper, a numerical study was conducted to assess the effects of real gas on the local flow separation of lifting body rudder. The effects of Mach number, wall temperature and flight altitude on separated flowfield features were highlighted. To this end, several numerical results derived for perfect gas, equilibrium gas and chemical nonequilibrium reacting gas were provided and compared. The calculation results show that dissociation reaction occurs within boundary layer and reduces the temperature when real gas effects is considered. Viscosity of gas within boundary layer becomes smaller, the loss of kinetic energy decreases and this strengthens the ability to overcome the adverse pressure gradient, so the flow becomes harder to separate. Real gas effects significantly decreases separation zone, and change the separation/attach shock wave position and strength. As a result of shock/shock and shock/boundary layer interactions, pressure and heat flux distribution near the separation zone are changed by real gas effects at the same time. From the comparison of equilibrium gas and perfect gas result at different altitude, a conclusion can be drawn that the flow separation becomes more sensitive to real gas effects at higher altitude. Wall temperature is an important factor to determine separated flowfield features. As wall temperature increases, separation zone size becomes larger as well as heat flux on the rudder decreases. As much number increases, the separation zone size decreases and the differences between real gas and perfect gas become larger, real gas effects is more significant.