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 %.

Shock-shock interactions are among the most basic gas-dynamic problem, and are almost unavoidable in any high speed light, where shock waves generating from different sources crosses each other paths. These interactions when present very close to the solid surface lead to very high pressure and thermal loads on the surface. The related practical problem is that experienced at the cowl lip of a scramjet engine, where the interfering shock waves leads to high heat transfer rates which may also lead to the damage of the material. The classification by Edney (1968) on the shock-shock interaction patterns based on the visualization has since then served the basis for such studies. Though the problem of high heating on the surface in the vicinity of the shock-shock interactions has been studied at length at supersonic Mach numbers, the study on the topic at the hypersonic Mach numbers is little sparse. Even in the studies at hypersonic Mach numbers, the high speeds are not simulated, which is the measure of the kinetic energy of the ow. Very few experimental studies have addressed this problem by simulating the energy content of the ow. Also, some of the numerical studies on the shock-shock interactions suggest the presence of unsteadiness in the shock-shock interaction patterns as observed by Edney (1968), though this observation is not made very clearly in the experimental studies undertaken so far. In the present study, experiments are carried out in a conventional shock tunnel at Mach number of 5.62 (total enthalpy of 1.07 MJ/kg; freestream velocity of 1361 m/s), with the objective of mapping the surface pressure distribution and surface convective heat transfer rate distribution on the hemispherical body in the presence of the shock-shock interactions. A shock generator which is basically a wedge of angle = 25 , is placed at some dis-dance in front of the hemispherical body such that the planar oblique shock wave from the shock generator hits the bow shock wave in front of the hemi-spherical body. The relative distance between the wedge tip and the nose of the hemispherical body is allowed to change in di erent experiments to capture the whole realm of shock-shock interaction by making the planar oblique shock wave interact with the bow shock wave at different locations along its trajectory. The study results in a bulk of data for the surface pressure and heat transfer rates which were obtained by placing 5 kulites pressure transducers, 1 PCB pressure transducer and 21 platinum thin lm gauges along the surface of the hemispherical body in a plane normal to the freestream velocity direction. Along with the measurement of the surface pressure and the surface heat transfer rates, the schlieren visualization is carried out to capture the shock waves, expansion fans, slip lines, present in a certain shock-shock interaction pattern and the measured values were correlated with the captured schlieren images to evaluate the ow build up and steady and useful test time thereby helping in understanding the ow physics in the presence of the shock-shock interactions. From the present study it has been observed that in the presence of Edney Type-I and Edney Type-II interaction, the heat transfer rates on the hemi-spherical body are symmetrical about the centerline of the body, with the peak heating at the centerline which drops towards the shoulder. For Edney Type-III, Edney Type-IV, Edney Type-V and Edney Type-VI interaction pattern, the distribution in not symmetrical and shifts in peak heat transfer rates being on the side of the hemispherical from which planar oblique shock wave is incident. Also, it is observed that for the interactions which appear within the sonic circle, Edney Type-III and Edney Type-IV, the heat transfer rates observe an unsteadiness, such that the gauges located close to the interaction region experiencing varying heat transfer rates during the useful test time of the shock tunnel. Few experiments were conducted at Mach 8.36 (total enthalpy of 1.29 MJ/kg; freestream velocity of 1555.25 m/s) and Mach 10.14 (total enthalpy of 2.67 MJ/kg; freestream velocity of 2258.51 m/s) for the con gurations representing Edney Type-III interaction pattern to further evaluate the unsteady nature observed at Mach 5.62 ows. The unsteadiness was evident in both the cases. It is realized that the short test times in the shock tunnels pose a constraint in the study of unsteady flow fields, and the use of tailored mode operation of shock tunnel can alleviate this constraint. Also, limited number of experiments in the present study, which are carried out in a Free Piston Shock Tunnel, helps to understand the need to conduct such study in high enthalpy test conditions.

The re-entry space vehicles encounter high temperatures when they enter the earth atmosphere and the high temperature air in the shock layer around the body undergoes partial dissociation. Also, the gas molecules injected into the shock layer from the ablative thermal protection system (TPS) undergo pyrolysis which helps in reducing the net heat flux to the vehicle surface. The chemical species due to the pyrolysis add complexity to the stagnation flow chemistry (52 chemical reactions) models which include species like NOx, CO and hydrocarbons (HCs). Although the ablative TPS is responsible for the safety of re-entry space vehicle, the induced chemical species result in variety of adverse effects on environment such as global warming, acid rain, green house effect etc. The well known three-way-catalyst (TWC) involves simultaneous removal of all the three gases (i.e, NOx, CO, Hydrocarbons) present in the shock layer. Interaction of such three-way-catalyst on the heat shield materials or on the wall of the re-entry space vehicle is to reduce the heat flux and to remove the gases in the shock layer, which is an important issue. For the re-entry vehicle the maximum aerodynamic heating occurs at an altitude ranging about 68 to 45 km during which the vehicle is surrounded by high temperature dissociated air. Then the simplest real gas model of air is the five species model which is based on N2, O2, O, NO and N. This five species model assumes no ionization and no pyrolysis gases are emitted from the heat shield materials. The experimental research work presented in this thesis is directed towards the understanding of catalytic and non-catalytic surface reactions on high temperature materials in presence of strong shock heated test gas. We have also explored the possibility of using shock tube as a high enthalpy device for synthesis of new materials. In the first Chapter, we have presented an overview of re-entry space vehicles, thermal protection system (TPS) and importance of real gas effects in the shock layer. Literature survey on TPS, ablative materials and aerothermochemistry at the stagnation point of reentry capsule, in addition to catalytic and non-catalytic surface reactions between the wall and dissociated air in the shock layer are presented. In Chapters 2 and 3, we present the experimental techniques used to study surface reactions on high temperature materials. A brief description of HST2 shock tunnel is presented and this shock tunnel is capable of generating flow stagnation enthalpies ranging from 0.7 to 5 MJ/kg and has an effective test time of ~ 800 µs. High speed data acquisition system (National Instruments and Yokogawa) used to acquire data from shock tube experiments. The experimental methods like X-ray Photoelectron Spectroscopy (XPS), Scanning Electron Microscopy (SEM), X-ray diffraction (XRD), Raman and FTIR spectroscopy have been used to characterize the shock-exposed materials. Preliminary research work on surface nitridation of pure metals with shock heated nitrogen gas is discussed in Chapter 2. Surface nitridation of pure Al thin film with shock heated N2 is presented in Chapter 3. An XPS study shows that Al 2p peak at 74.2 eV is due to the formation AlN on the surface of Al thin film due to heterogeneous non-catalytic surface reaction. SEM results show changes in surface morphology of AlN film due to shock wave interaction. Thickness of AlN film on the surface increased with the increase in temperature of the shock heated nitrogen gas. However, HST2 did not produce sufficient temperature and pressure to carry out real conditions of re-entry. Therefore design and development of a new high enthalpy shock tunnel was taken up. In Chapter 4, we present the details of design and fabrication of free piston driven shock tunnel (FPST) to generate high enthalpy test gas along with the development of platinum (Pt) and thermocouple sensors for heat transfer measurement. A free piston driven shock tunnel consists of a high pressure gas reservoir, compression tube, shock tube, nozzle, test section and dump tank connected to a vacuum pumping system. Compression tube has a provision to fill helium gas and four ports, used to mount optical sensors to monitor the piston speed and pressure transducer to record pressure at the end of the compression tube when the piston is launched. Piston can attain a maximum speed of 150 m/s and compress the gas inside the compression tube. The compressed gas bursts the metal diaphragm and generates strong shock wave in the shock tube. This tunnel produces total pressure of about 300 bar and temperature of about 6000 K and is capable of producing a stagnation enthalpy up to 45 MJ/kg. The calibration of nozzle was carried out by measuring the pitot tube pressure in the dump tank. Experimentally recorded P5 pressure at end of the shock tube is compared with Numerical codes. Calibrated pressure P5 values are used to calculate the temperature T5 of the reflected shock waves. This high pressure and high temperature shock heated test gas interacts with the surface of the high temperature test materials. For the measurement of heat transfer rate, platinum thin film sensors are developed using DC magnetron sputtering unit. Hard protective layer of aluminum nitride (AlN) on Pt thin film was deposited by reactive DC magnetron sputtering to measure heat transfer rate in high enthalpy tunnel. After the calibration studies, FPST is used to study the heat transfer rate and to investigate catalytic/non-catalytic surface reaction on high temperature materials. In Chapter 5, an experimental investigation of non-catalytic surface reactions on pure carbon material is presented. The pure carbon C60 films and conducting carbon films are deposited on Macor substrate in the laboratory to perform shock tube experiments. These carbon films were exposed to strong shock heated N2 gas in the shock tube portion of the FPST tunnel. The typical shock Mach number obtained is about 7 with the corresponding pressure and temperature jumps of about 110 bar and 5400 K after reflection at end of the shock tube. Shock exposed carbon films were examined by different experimental techniques. XPS spectra of C(1s) peak at 285.8 eV is attributed to sp2 (C=N) and 287.3 eV peak is attributed to sp3 (C-N) bond in CNx due to carbon nitride. Similarly, N(1s) core level peak at 398.6 eV and 400.1 eV observed are attributed to sp3-C-N and sp2-C=N of carbon nitride, respectively. SEM study shows the formation of carbon nitride crystals. Carbon C60 had melted and undergone non-catalytic surface reaction with N2 while forming carbon nitride. Similar observations were made with conducting carbon films but the crystals were spherical in shape. Micro Raman and FTIR study gave further evidence on the formation of carbon nitride film. This experimental investigation confirms the formation of carbon nitride in presence of shock-heated nitrogen gas by non-catalytic surface reaction. In Chapters 6 and 7, we present a novel method to understand fully catalytic surface reactions after exposure to shock heated N2, O2 and Ar test gas with high temperature materials. We have employed nano ZrO2 and nano Ce0.5Zr0.5O2 ceramic high temperature materials to investigate surface catalytic reactions in presence of shock heated test gases. These nano crystalline oxides are synthesized by a single step solution combustion method. Catalytic reaction was confirmed for both powder and film samples of ZrO2. As per the theoretical model, it is known that the catalytic recombination reaction produces maximum heating on the surface of re-entry space vehicles. This was demonstrated in this experiment when a metastable cubic ZrO2 changed to stable monoclinic ZrO2 phase after exposure to shock waves. The change of crystal structure was seen using XRD studies and needle type monoclinic crystal growth with aspect ratio (L/D) more than 15 was confirmed by SEM studies. XPS of Zr(3d) core level spectra show no change in binding energy before and after exposure to shock waves, confirming that ZrO2 does not change its chemical nature, which is the signature of catalytic surface reaction. When a shock heated argon gas interacted with Ce0.5Zr0.5O2 compound, there was a change in colour from pale yellow to black due to reduction of the compound, which is the effect of heat transfer from the shock wave to the compound in presence of argon gas. The reduction reaction shows the release of oxygen from the compound due to high temperature interaction. The XPS of Ce(3d) and Zr(3d) spectra confirm the reduction of both Ce and Zr to lower valent states. The oxygen storage and release capacity of the Ce0.5Zr0.5O2 compound was confirmed by analyzing the reduction of Ce4+ and Zr4+ with high temperature gas interaction. When Ce0.5Zr0.5O2 (which is same as Ce2Zr2O8) in cubic fluorite structure was subjected to strong shock, it changed to pyrochlore (Ce2Zr2O7) structure by releasing oxygen and on further heating it changed to Ce2Zr2O6.3 which is also crystallized in pyrochlore structure by further releasing oxygen. If this heating is carried out in presence of argon test gas, fluorite structure can easily change to pyrochlore Ce2Zr2O6.3 structure, which is a good electrical conductor. Due to its oxygen storage capability (OSC) and redox (Ce4+/Ce3+) properties, Ce0.5Zr0.5O2 had been used as oxygen storage material in three-way-catalyst. Importance of these reactions is that the O2 gas released from the compound will react with gas released from the heat shield materials, like NOx, CO and hydrocarbon (HCs) species which results in reduction of temperature in the shock layer of the re-entry space vehicle. The compound Ce0.5Zr0.5O2 changes its crystal structure from fluorite to pyrochlore phase in presence of shock heated test gas. The results presented in these two Chapters are first of their kind, which demonstrates the surface catalytic reactions. In Chapter 8, we present preliminary results of the oxygen recombination on the surface of heat shield material procured from Indian Space Research Organization (ISRO) used as TPS in re-entry space capsule (Space capsule Recovery Experiment SRE-1) and on thin film SiO2 deposited on silicon substrate. The formation of SiO between the junctions of SiO2/Si was confirmed using XPS study when shock exposed oxygen reacted on these materials. The surface morphology of the ablated SiO2 film was studied using SEM. The damage induced due to impact of shock wave in presence of oxygen gas was analyzed using Focused Ion Beam (FIB) microscope. The results reveal the damage on the surface of SiO2 film and also in the cross-section of the film. We are further investigating use of FIB, particularly related to residual stress developed on thin films due to high pressure and high temperature shock wave interaction. In Chapter 9, conclusions on the performance of FPST, synthesis of high temperature materials, catalytic and non-catalytic surface reactions on the high temperature material due to shock-heated test gases are presented. Possible scope for future studies is also addressed in this Chapter.

One of the characteristics of the high speed ows over blunt bodies is the detached shock formed in front of the body. The distance of the shock from the stagnation point measured along the stagnation streamline is termed as the shock stand o distance or the shock detachment distance. It is one of the most basic parameters in such ows. The need to know the shock stand o distance arises due to the high temperatures faced in these cases. The biggest challenge faced in high enthalpy ows is the high amounts of heat transfer to the body. The position of the shock is relevant in knowing the temperatures that the body being subjected to such ows will have to face and thus building an efficient system to reduce the heat transfer. Despite being a basic parameter, there is no theoretical means to determine the shock stand o distance which is accepted universally. Deduction of this quantity depends more or less on experimental or computational means until a successful theoretical model for its predictions is developed. The experimental data available in open literature for spherical bodies in high speed ows mostly lies beyond the 2 km/s regime. Experiments were conducted to determine the shock stand o distance in the velocity range of 1-2 km/s. Three different hemispherical bodies of radii 25, 40 and 50 mm were taken as test models. Since the shock stand o distance is known to depend on the density ratio across the shock and hence gamma (ratio of specific heats), two different test gases, air and carbon dioxide were used for the experiments here. Five different test cases were studied with air as the test gas; Mach 5.56 with Reynolds number of 5.71 million/m and enthalpy of 1.08 MJ/kg, Mach 5.39 with Reynolds number of 3.04 million/m and enthalpy of 1.42 MJ/kg Mach 8.42 with Reynolds number of 1.72 million/m and enthalpy of 1.21 MJ/kg, Mach 11.8 with Reynolds number of 1.09 million/m and enthalpy of 2.03 MJ/kg and Mach 11.25 with Reynolds number of 0.90 million/m and enthalpy of 2.88 MJ/kg. For the experiments conducted with carbon dioxide as test gas, typical freestream conditions were: Mach 6.66 with Reynolds number of 1.46 million/m and enthalpy of 1.23 MJ/kg. The shock stand o distance was determined from the images that were obtained through schlieren photography, the ow visualization technique employed here. The results obtained were found to follow the same trend as the existing experimental data in the higher velocity range. The experimental data obtained was compared with two different theoretical models given by Lobb and Olivier and was found to match. Simulations were carried out in HiFUN, an in-house CFD package for Euler and laminar own conditions for Mach 8 own over 50 mm body with air as the test gas. The computational data was found to match well with the experimental and theoretical data

Aerodynamic cavities are common features on hypersonic vehicles which are caused in both large and small scale features like surface defects, pitting, gap in joints etc. In the hypersonic regime, the presence of such cavities alters the flow phenomenon considerably and heating rates adjacent to the discontinuities can be greatly enhanced due to the diversion of flow. Since the 1960s, a great deal of theoretical and experimental research has been carried out on cavity flow physics and heating. However, most of the studies have been done to characterize the effect downstream and within the cavity. In the present study, a series of were carried out in the shock tunnel to investigate the heating characteristics, upstream and on the lateral side of the cavity. Heat flux measurement has been done using indigenously developed high resistance platinum thin film gauges. High resistance gauges, as contrary to the conventionally used low resistance gauges were showing good response to the extremely low heat flux values on a flat plate with sharp leading edge. The experimental measurements of heat done on a flat plate with sharp leading edge using these gauges show good match with theoretical relation by Crabtree et al. Flow visualization using high speed camera with the cavity model and shock structures visualized were similar to reported in supersonic cavity flow. This also goes to state that in spite of the fluctuating shear layer-the main feature of hypersonic flow over a cavity ,reasonable studies can be done within the short test time of shock tunnel. Numerical Simulations by solving the Navier-Stokes equation, using the commercially available CFD package FLUENT 13.0.0 has been done to complement the experimental studies.