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This paper aims to numerically investigate the influence of shock wave and freestream turbulence interaction on the parabolic and spherically blunted nose cones at supersonic speed.

Using density-based solver, the three-dimensional steady-state simulation is carried out. The working fluid is calorically perfect that obeys ideal gas law and the no-slip boundary conditionis given to the surface of the nose cone. Pressure far-field boundary condition is imposed at the boundary of the computational domain by giving freestream Mach number, freestream static pressure and temperature.

The growth rate of the boundary layer is faster on the spherically blunted nose cone, hence, the overall drag force is higher than the parabolic nose cone. Temperature at the edge of the boundary layer is increased due to the early ampli-fication of instabilities by the upstream disturbance. In this sense, the effects of freestream turbulence depend on its level, freestream conditions, strength and type of shock wave and zone of influence.

Simulations are carried out for the flow Mach number 2.0 at zero angles of attack for the freestream conditions of the flow at an altitude of 10,000 m.

The phenomenon of shock wave–turbulence interaction occurs in flow regimes from transonic to hypersonic speeds and finds a wide range of applications, especially in the design of aircraft and missiles configurations.

The phenomenon of compression wave and freestream turbulence interaction around the commonly used nose cones in the case of aircraft, missiles, etc., is investigated. The performance characteristics such as aerodynamic drag, boundary layer dynamics and the nature of flow around the different nose cones at zero angle of attack are illustrated.

The purpose of this paper is to present a non‐equilibrium viscous shock layer (VSL) solution procedure that considerably improves computational efficiency, especially for long slender bodies.

The VSL equations are solved in a shock oriented coordinate system. The method of solution is spatial marching, implicit, finite‐difference technique, which includes coupling of the normal momentum and continuity equations. In the nose region, the shock shape is specified from an algebraic expression and corrected through global passes through that region. The shock shape is computed as part of the solution beyond the nose region and requires only a single global pass. For this study, a seven‐species (O₂, N₂, O, N, NO, NO⁺, e⁻) air model is used.

The present approach eliminates the need for initial shock shape, which was required by previous method of solution. This method generates its own shock shape as a part of solution and the input shock shape obtained from a different solution is not required. Therefore, in comparison with the other VSL methods, the present approach dramatically reduces the CPU time of calculations. Moreover, by using the shock oriented coordinate systems the junction point problem in sphere‐cone configurations is solved.

This method is an excellent tool for parametric study and preliminary design of hypersonic vehicles.

The present method provides a computational capability which reduces the CPU time, and expands the range of application for the prediction of hypersonic heating rates.