Efficient combustion and heat release in scramjet flows depend on effective mixing of the fuel in supersonic streams. Usually, transverse sonic injection in-stages are employed as one of the suitable means for efficient supersonic combustor design. Numerical simulations are carried out to study the mixing characteristics of staged sonic air injections in supersonic stream (M ¼ 2.07) behind a backward-facing step in scramjet combustor by solving threedimensional Navier -Stokes equations along with K -1 turbulence model with a commercial CFD software CFX-TASCFlow. Computed results of the jet penetration and spreading show very good agreement with the experimental values and the results of other computations. A good overall match has been obtained between the experimental values and the computation for various flow profiles at various axial locations in the combustor. However, the values differ in the near-field region at the injection plane. The assumed uniformity of the flow-field properties at the injection orifice and/or the inadequacy of the turbulence model considered in this study is conjectured to be the cause of the difference.

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Numerical simulations are carried out for the internal flow field of a dual pulse solid rocket motor port to understand the flow behaviour. Three dimensional Reynolds Averaged Navier Stokes equations are solved alongwith shear stress transport turbulence model using commercial code. The combustion gas is assumed as a mixture of alumina and gases and single phase flow calculations are done with the thermo chemical properties provided for the mixture. The simulation captures all the essential features of the flow field. The flow accelerates through the pulse separation device (PSD) port and high temperature and high velocity gas is seen to impinge the motor wall near the PSD port. The overall total pressure drop through motor port and through PSD is found to be moderate.

Conjugate heat transfer studies are presented for high speed aerospace vehicle using commercial CFD software. Navier Stokes equations in the fluid domain and transient heat conduction equations in the solid domain are solved simultaneously to obtain the skin temperature history and other flow parameters. The computational methodology is applied to predict the surface temperature of high speed aerospace vehicle after validating the methodology against experimental results. Validation cases include laminar flow past axisymmetric double cone at Mach 4.57 and turbulent flow past circular cylinder at Mach 6.7. Computed flow field including cold wall heat flux, surface temperature distribution, surface temperature history match nicely with experimental as well as other numerical results. Temperature dependent material properties are found to have significant effect on the surface temperature prediction. Computed surface temperature of a high speed aerospace vehicle show good overall match with flight measured values.