Text

theses

In this Ph.D. dissertation, two separate phenomena have been numerically studied: flow control using a laser discharge in a supersonic flow and laminar shock wave boundary layer interaction in a hypersonic flow. In the first section of the study, the interaction of a laser-generated plasma with a hemisphere cylinder at Mach 3.45 is simulated using the Euler and Navier Stokes equations, separately and assuming a perfect gas with no chemical reactions in the laser discharge. The instantaneous laser discharge creates a plasma region which in this study is assumed to be spherical. From this spherical plasma region, a blast wave and an expansion wave form which propagate radially outward and inward, respectively. The heated region convects with the flow and interacts with the blunt body shock in the upstream of the hemisphere and changes the flow structure and parameters in that region. The impact of the blast wave with the hemisphere surface momentarily raises the pressure on the hemisphere. When the heated region reaches the blunt body shock lensing of the shock wave occurs and a toroidal vortex forms due to the Richtmyer-Meshkov instability; as a result, the pressure on the hemisphere drops momentarily. Later on, the flow parameters converge to their steady state condition as the heated region convects to the downstream of the hemisphere. The results are compared with experimental data of a separate study to validate the numerical model used in these simulations. To investigate the hypersonic shock wave laminar boundary layer interaction, two separate geometric configurations are used: axisymmetric flow over a hollow-cylinder flare, and three-dimensional flow over a cylindrically blunted fin mounted on a flat plate. In the first case, the capability of the chosen numerical model in predicting the pressure and heat transfer in a hypersonic shock wave boundary layer interaction over an axisymmetric hollow cylinder flare at a Mach 10 flow is investigated. In the second case, the assessment of the capability of a laminar perfect gas model to predict the heat transfer in a three-dimensional hypersonic flow with shock wave boundary layer interaction was studied. In this study, the freestream Mach number and Reynolds number - based on the diameter of the cylindrical fin - are 14 and 8,000, respectively. Numerical heat transfer on the blunt fin is compared with the experimental data for validation. Moreover, investigation of the effects of the sweep angle of the blunt fin on the shock wave boundary layer interaction is the other objective of this research. Three discrete sweep angles of zero, 22.5 and 45 degree have been chosen and comparison of their results have been made. It has been shown that the adverse pressure gradient imposed from the shock wave to the boundary layer can separate the boundary layer. The separation shock wave formed over the separated region can interact with the other shock waves and create a lambda shock wave structure with a transmitted shock wave. As the separated boundary layer reattaches to the surface, it increases the localized heat transfer and produces a reattachment shock wave, which increases the pressure on the surface of the vehicle. The localized high aerothermodynamic loads as well as the low frequency oscillations regarding the shock wave boundary layer interaction impose design limitations on the hypersonic aircrafts and show the importance of fully understanding the physics behind these phenomena as well as gaining the ability to predict the flow with such interactions.

ETD_8578

Ph.D.

Includes bibliographical references

by Mahsaalsadat Mortazavi Ravari

doi:10.7282/T3251ND2

ETD doctoral

The author owns the copyright to this work.