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.

The interaction of a shock wave and boundary layer often occurs in high speed flows. For sufficiently strong shock strengths the boundary layer separates, generating shock patterns in the contiguous inviscid flow (termed strong interactions); which may also affect the performances of the systems where they occur, demanding control of the interaction to enhance the performances. The case of impinging shock wave boundary layer interaction is of fundamental importance and can throw light on the physics of the interaction in general. Although various aspects of the interaction are studied at supersonic speeds, the complexities involved in the interaction at hypersonic speeds are not well understood. Of importance is the high total enthalpy associated with hypersonic flows the simulation of which requires shock tunnels. The present experimental study focuses on the interaction between strong impinging shock and boundary layer in hypersonic flows of moderate to high total enthalpies. Experiments are performed in hypersonic shock tunnels HST-2 and FPST (free piston driven shock tunnel), at nominal Mach numbers 6 and 8, with total enthalpy ranging from 1.3 MJ/kg to 6 MJ/kg, and freestream Reynolds number ranging from 0.3 million/m to 4 million/m. The strong impinging shock is generated by a wedge of angle 30.960 to the freestream. The shock is made to impinge on a flat plate (made of Hylem which is adiabatic, except for one case with plate made of aluminium which allows heat transfer). The position of (inviscid) shock impingement may be varied (from 55 mm from the leading edge to 100 mm from the leading edge) by moving the plate back and forth on the fixture which holds the wedge and the plate. Expectedly the strong shock generates a large separation bubble of length comparable to the distance of the location of shock impingement from the leading edge of the plate. Such large separation bubbles are typical of supersonic/hypersonic intakes at off-design operation. The evolution of the flow field- including the evolution of impinging shock and subsequent evolution of the large separation bubble- within the short test duration of the shock tunnels is one of the main concerns addressed in the study. Time resolved schlieren flow visualizations using high speed camera, surface pressure measurements using PCB, kulite and MEMS sensors, surface convective heat transfer measurements using platinum thin film sensors are the flow diagnostics used. From the time resolved visualizations and surface pressure measurements with the fast response sensors, the flow field, even with a separation bubble as large as 75 mm (at Mach 5.96, with shock impingement at 95 mm from the leading edge) was found to be established within the short shock tunnel test time. The effects of various parameters- freestream Mach number, distance of the location of shock impingement, freestream total enthalpy and wall heat transfer- on the interaction are investigated. With increase in Mach number from 5.96 to 8.67, for nearly the same shock impingement locations (95 mm and 100 mm from the leading edge respectively), the separation length decreased from 75 mm to 60 mm despite the fact that the shocks are doubly stronger at the higher Mach number. Inflectional trend in separation length was observed with enthalpy at nominal Mach number 8- separation length increased from 60 mm at 1.6 MJ/kg to 70 mm at 2.4 MJ/kg, and decreased drastically to ~40 mm at 6 MJ/kg (when dissociations are expected). The separation length Lsep for all the experiments, except the experiments at 6 MJ/kg, were found to be large, i.e. comparable with the distance xi of location of shock impingement from the leading edge of the flat plate. The scaled separation length (with Hylem wall) was found to obey the inviscid similarity law proposed from the present study for large separation bubbles with strong impinging shocks, where M∞ is the freestream Mach number, p∞ is the freestream pressure and pr is the measured reattachment pressure; this holds for freestream total enthalpy ranging from 1.3 MJ/kg to 2.4 MJ/kg and Reynolds number (based on location of shock impingement) ranging from 1x105 to 4x105. While the increase in separation length from 1.6 MJ/kg to 2.4 MJ/kg could thus be attributed to the small difference in Mach number between the cases (due to inverse variation with cube of Mach number), the decrease in separation length and the non-confirmation to the proposed similarity law for the 6 MJ/kg case is attributed to the real gas effects. At Mach 6 the flow was observed to separate close to the leading edge, even when the (inviscid) shock impingement was at 95 mm from the leading edge. This prompted the proposal of an approximate inviscid model of the interaction for the Mach 6 case with separation at leading edge, and reattachment at the location of (inviscid) shock impingement; Accordingly, the closer the location of impingement, the more the angle that the separated shear layer makes with the plate and hence more the pressure inside the separation bubble. A small reduction in separation length was also observed with aluminium wall when compared with Hylem wall, emphasizing the importance of wall heat conductivity (especially when concerning separated flows) even within the short test durations of shock tunnels. The free interaction theory over adiabatic wall was found to predict the pressure at the location of separation, but under-predict the plateau pressure (at nominal Mach number 8). Numerical simulations (steady, planar) were also carried out using commercial CFD solver FLUENT to complement the experiments. Simulations using one equation turbulence model (Spalart-Allmaras model) were closer to the experimental results than the laminar simulations, suggesting that the flow field may be transitional or turbulent after separation. Significant reduction of the separation bubble length was demonstrated with the control of the interaction using boundary layer bleed within the short test time of the shock tunnel; with tangential blowing at the separation location20% reduction in separation length was observed, while with suction at separation location the reduction was 13.33 %.