Numerical Investigations on the Impact of Turbulent Prandtl Number and Schmidt Number on Supersonic Combustion

The flow field inside the combustor of a scramjet is highly complicated and the related turbulent Prandtl and Schmidt numbers have a significant impact on the effective numerical prediction of such dynamics.As in many cases researchers set these parameters on the basis of purely empirical laws,assessing their impact(via parametric numerical simulations)is a subject of great importance.In the present work,in particular,two test cases with different characteristics are selected for further evaluation of the role played by these non-dimensional numbers:Burrows-Kurkov case and DLR case.The numerical results indicate that these parameters influence ignition location.Moreover,the temperature distribution is more sensitive to them than to H2O mass fraction and velocity distributions.

Numerical Investigation on Flow and Heat Transfer Performance of Supercritical Carbon Dioxide Based on Variable Turbulent Prandtlnumber Model

Flow and heat transfer characteristic of supercritical carbon dioxide(SCO_(2))are numerically investigated in the horizontal and vertical tubes.TWL turbulent Prandtl number model could well describe the behavior of SCO_(2) affected by the buoyancy.Under the cooling condition,the heat transfer performance of SCO_(2) along the upward direction is best and that along the downward direction is worst when bulk fluid temperatures are below the pseudocritical temperature.Reducing the ratio of heat flux to mass flux could decrease the difference of convective heat transfer coefficient in three flow directions.Under the heating condition,heat transfer deterioration only occurs in vertical upward and horizontal flow directions.Heat transfer deterioration of SCO_(2) could be delayed by increasing the mass flux and the deterioration degree is weakened in the second half of tube along the vertical upward flow direction.Compared with the straight tube,the corrugated tube shows better comprehensive thermal performance.

Numerical simulation of compression corner flows at Mach number 9

A hypersonic vehicle encounters a wide range of conditions during its complete flight regime.These flight conditions may vary from low to high Mach numbers with varying angles of attack.The near-wall viscous dissipation associated with flows at combined high Mach and Reynolds numbers leads to significant wall heat transfer rates and shear stresses.The shock wave/boundary-layer interaction results in a flow separation region,which commonly augments total pressure losses in the flow and lowers the efficiency of aerodynamic control surfaces such as fins installed on a vehicle.The standard turbulence models,when used to resolve such flows,result in incorrect separation bubble size for large separated flows.Therefore,it results in an inaccurate aerodynamic load,such as the wall pressures,skin friction distribution,and heat transfer rate.In previous studies,the application of the shock-unsteadiness correction to the standard two-equation k-ωturbulence model improved the separation bubble size leading to an accurate pressure prediction and shock definition with the assumption of constant Prandtl number.In the present work,the new shock-unsteadiness modification to the k-ωturbulence model is applied to the hypersonic compression corner flows.This new model with variable Prandtl number is based on the model parameter,which depends upon the local density ratio.The computed wall pressures,heat flux and flow field are compared to the experimental data.A parametric study is carried out by varying compression deflection angles,free stream Reynolds number and wall temperatures to compute the flow field and wall data accurately,particularly in the shock boundary layer interaction region.The new shockunsteadiness modified k-ωmodel with variable Prandtl number shows an accurate prediction of initial pressure rise location,pressure distribution in the plateau region and heat flux in comparison to the standard k-ωmodel.