Supersonic inlet is often considered to be the most crucial element in the air-breathing engine of a flying vehicle designed to operate at supersonic speeds. The function of the inlet is to provide subsonic flow at the desired velocity to the downstream components of the engine with minimum loss in the total pressure recovery, flow uniformity, and flow stability, all of which are important to the overall engine efficiency. Accordingly proper simulation of the inlet is one of the most important steps in the designing of the propulsion system of any flying vehicle particularly flying at supersonic speeds. A numerical code with possibility of proper simulation of normal and oblique shocks, boundary layer, interaction of shock and boundary layer and … can help designer with respect to the time, cost and accuracy of the design process. So, in this research we developed a numerical code to solve the Navier-Stokes equations with explicit discretization around an axisymmetric supersonic external compression inlet. To validate the numerical code results, wind tunnel tests were conducted in a trisonic wind tunnel in Iran. The discretized governing equations were applied in a structured grid that was generated with the elliptic grid generator. Because of the flow is steady, the time step has been calculated with the local time stepping method to accelerate the convergence. Inviscid flux has been calculated with Roe, Van Leer, AUSM+ and AUSMPW methods. Different methods have been used to find the best one. In order to increase the spatial accuracy of discretization the MUSCL method has been used. For proper simulation of the boundary layer and its interaction with the shocks, the turbulent viscosity must be added to the molecular viscosity. In this research the Baldwin-Lomax model has been used to compute the turbulent viscosity. Finally, applying the above numerical methods, a good agreement has been observed between the numerical and experimental results.

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