The results of experimental investigation of a turbulent boundary layer on compression and expansion surfaces are presented. They include the study of the shock wave and/or expansion fan action upon the boundary layer...The results of experimental investigation of a turbulent boundary layer on compression and expansion surfaces are presented. They include the study of the shock wave and/or expansion fan action upon the boundary layer, boundary layer separation and its relaxation. Complex events of paired interactions and the flow on compression convex-concave surfaces were studied. The possibility and conditions of the boundary layer relaminarization behind the expansion fan and its effect on the relaxation length are presented. Different model configurations for wide range conditions were investigated. Comparison of results for different interactions was carried out.展开更多
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 dissipati...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.展开更多
Hydrogen peroxide(H_2O_2) has its significance during the combustion of heavy hydrocarbons in the internal combustion(IC) engines. Owing to its importance the measurements of H_2O_2 dissociation rate have been reporte...Hydrogen peroxide(H_2O_2) has its significance during the combustion of heavy hydrocarbons in the internal combustion(IC) engines. Owing to its importance the measurements of H_2O_2 dissociation rate have been reported mostly using the shock tube apparatus. These types of experimental measurements are although quite reliable but require high cost. On the other hand, numerical simulations provide low cost and reliable solutions especially using computation fluid dynamics(CFD) software. In the current study an experimental shock tube flow is modeled using open access platform OpenFOAM to investigate the thermal decomposition of H_2O_2. Using two different convective schemes, limited Linear and upwind, the propagation of shock wave and resultant dissociation reaction are simulated. The results of the simulations are compared with the experimental data. It is observed that the rate constant measured using the simulation data deviates from the experimental results in the low temperature range and approaches the experimental values as the temperature is raised.展开更多
文摘The results of experimental investigation of a turbulent boundary layer on compression and expansion surfaces are presented. They include the study of the shock wave and/or expansion fan action upon the boundary layer, boundary layer separation and its relaxation. Complex events of paired interactions and the flow on compression convex-concave surfaces were studied. The possibility and conditions of the boundary layer relaminarization behind the expansion fan and its effect on the relaxation length are presented. Different model configurations for wide range conditions were investigated. Comparison of results for different interactions was carried out.
基金financially supported by the Deanship of Scientific Research(DSR),King Abdulaziz University,Jeddah,under grant No.DF-043-135-1441。
文摘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.
文摘Hydrogen peroxide(H_2O_2) has its significance during the combustion of heavy hydrocarbons in the internal combustion(IC) engines. Owing to its importance the measurements of H_2O_2 dissociation rate have been reported mostly using the shock tube apparatus. These types of experimental measurements are although quite reliable but require high cost. On the other hand, numerical simulations provide low cost and reliable solutions especially using computation fluid dynamics(CFD) software. In the current study an experimental shock tube flow is modeled using open access platform OpenFOAM to investigate the thermal decomposition of H_2O_2. Using two different convective schemes, limited Linear and upwind, the propagation of shock wave and resultant dissociation reaction are simulated. The results of the simulations are compared with the experimental data. It is observed that the rate constant measured using the simulation data deviates from the experimental results in the low temperature range and approaches the experimental values as the temperature is raised.
