Shock boundary layer interactions (SBLI) is an undesirable event occurring in high-speed air-breathing propulsion system that stimulates boundary layer separation due to adverse pressure gradients and consequently lead to ow distortion and pressure loss in the intake section. Therefore it is essential to apply ow control mechanisms to prevent this phenomenon. This study involves a novel ow control device called micro-ramp, which is a part of the micro-vortex generator family that has shown great potential in solving the adverse phenomenon. The term micro refers to the height of the device, which is smaller than the boundary layer thickness, δ. It is important to highlight the two main novelties of this investigation. Firstly, it is the first micro-ramp study conducted in the hypersonic ow regime (Mach 5) since most of the previous micro-ramp studies were only performed in subsonic, transonic and supersonic flows. Another novelty is the various experimental techniques that were used in single study for example schlieren photography, oil-dot and oil- ow visualisation and conventional pressure transducers. In addition, advanced ow diagnostic tools such as infrared thermography, pressure sensitive paints (PSP) and particle image velocimetry (PIV) were also employed. T

A body exposed to hypersonic flow is subjected to extremely high wall heating rates, owing to the conversion of the kinetic energy of the oncoming flow into heat through the formation of shock waves and viscous dissipation in the boundary layer and this is one of the main concerns in the design of any hypersonic vehicle. The conventional way of tackling this problem is to use a blunt fore-body, but it also results in an increase in wave drag and puts the penalty of excessive load on the propulsion system. An alternative approach is to alter the flow field using external means without changing the shape of the body; and several such methods are reported in the literature. The superiority of such methods lie in the fact that the effective shape of the body can be altered to meet the requirements of low wave drag, without having to pay the penalty of an increased wall heat transfer rate. Among these techniques, the use of local energy addition in the freestream to alter the flow field is particularly promising due to the flexibility it offers. By the suitable placement of the energy source relative to the body, this method can be effectively used to reduce the wave drag, to generate control forces and to optimise the performance of inlets. Although substantial number of numerical investigations on this topic is reported in the literature, there is no experimental evidence available, especially under hypersonic flow conditions, to support the feasibility of this concept. The purpose of this thesis is to experimentally investigate the effect of energy deposition on the flow-field of a 120� apex angle blunt cone in a hypersonic shock tunnel. Energy deposition is done using an electric arc discharge generated between two electrodes placed in the free stream and various parameters influencing the effectiveness of this technique are studied. The effect of energy deposition on aerodynamic parameters such as the drag force acting on the model and the wall heat flux has been investigated. In addition, the unsteady flow field is visualised using a standard Z-type schlieren flow visualisation setup. The experimental studies have shown a maximum reduction in drag of 50% and a reduction in stagnation point heating rate of 84% with the deposition of 0.3 kW of energy. The investigations also show that the location of energy deposition has a vital role in determining the flow structure; with no noticeable effects being produced in the flow field when the discharge source is located close to the body (0.416 times body diameter). In addition, the type of the test gas used is also found to have a major influence on the effectiveness of energy deposition, suggesting that thermal effects of energy deposition govern the flow field alteration mechanism. The freestream mass flux is also identified as an important parameter. These findings were also confirmed by surface pressure measurements. The experimental evidence also indicates that relaxation of the internal degrees of freedom play a major role in the determination of the flow structure. For the present experimental conditions, it has been observed that the flow field alteration is a result of the interaction of the heated region behind the energy spot with the blunt body shock wave. In addition to the experimental studies, numerical simulations of the flow field with energy deposition are also carried out and the experimentally measured aerodynamic drag with energy deposition is found to match reasonably well with the computed values.