The advancement of each component of aerospace vehicles is necessary as the continual demand for more aggressive missions are created. Improvements in propulsion and guidance system electronics are invaluable; however without material development to protect the vehicle from its environment those advances will not have a practical application. Thermal protection systems (TPS) are required in both external applications; for example on reentry vehicles, as well as in internal applications; to protect the casing of rockets and missiles. This dissertation focuses on a specific type of internal solid rocket motor TPS, ablatives. Ablatives have been used for decades on aerospace vehicles. To protect the motor from the hostile environment, these materials pyrolyze and char. Both of these mechanisms produce a boundary between the combustion gases and the motor as well as release the heat that the decomposed material has absorbed. These sacrificial materials are intended to protect the casing that it is attached to. With the development of polymer nanocomposites (PNCs) in the last couple of decades, it is of interest to see how these two fields can merge. Three different nanomaterials (carbon nanofibers, multiwall carbon nanotubes, and nanoclays) are examined to observe how each behaves in environments that simulate the motor firing conditions. These nanomaterials are individually added to a thermoplastic polyurethane elastomer (TPU) at different loadings, creating three distinct families of polymer nanocomposites. To describe a materials ablative performance, a number of material properties must be individually studied; such as thermal, density, porosity, char strength, and rheology. Different experiments are conducted to isolate specific ablative processes in order to identify how each nanomaterial affects the ablative performance. This dissertation first describes each material and the ablative processes which are characterized by each experiment. Then basic material properties of each family of materials are described. Degradation and flammability experiments then describe the degassing processes. Studies of the material char are then performed after full blown rocket experiments are done. These tests have shown that of the three nanomaterials, nanoclay enhances the TPU ablative performance the most while the CNF provides the least enhancement. / text

Submitted by PEDRO GUILHERME SILVA PESCI null (peg_pesci@hotmail.com) on 2017-11-30T13:40:46Z No. of bitstreams: 1 Versão Final.pdf: 56735184 bytes, checksum: 4e1d7d9c3a8db1009fa3e028174b3593 (MD5) / Approved for entry into archive by Pamella Benevides Gonçalves null (pamella@feg.unesp.br) on 2017-11-30T17:38:21Z (GMT) No. of bitstreams: 1 pesci_pgs_me_guara.pdf: 56735184 bytes, checksum: 4e1d7d9c3a8db1009fa3e028174b3593 (MD5) / Made available in DSpace on 2017-11-30T17:38:21Z (GMT). No. of bitstreams: 1 pesci_pgs_me_guara.pdf: 56735184 bytes, checksum: 4e1d7d9c3a8db1009fa3e028174b3593 (MD5) Previous issue date: 2017-11-24 / Materiais utilizados em componentes de veículos espaciais, como em tubeiras ou superfícies expostas à reentrada atmosférica, são sujeitos a ambientes termicamente agressivos. Este trabalho apresenta estudos envolvendo o desempenho de materiais compósitos utilizados em sistemas de proteção térmica, a partir da exposição a jatos de plasma, onde os fluxos de calor são comparáveis aos da reentrada atmosférica de componentes de veículos espaciais. Amostras de compósitos ablativos de carbono/fenólica foram ensaiadas no túnel de plasma do Laboratório de Plasmas e Processos do ITA (Instituto Tecnológico de Aeronáutica), por meio de uma tocha de plasma alimentada por uma fonte de energia elétrica de corrente contínua de 50kW. Os parâmetros de operação do túnel de plasma foram otimizados para reproduzirem as condições próximas do ponto crítico de reentrada das cargas úteis dos veículos espaciais desenvolvidos pelo IAE (Instituto de Aeronáutica e Espaço). As amostras em estudo foram desenvolvidas e fabricadas no Brasil, a partir de materiais de especial interesse do IAE. Para comparação, foi também ensaiado outro material com propriedades já bem estabelecidas como o teflon, sob as mesmas condições ablativas. Foram determinadas as perdas de massa e as taxas de perda de massa específicas das amostras, as temperaturas radiométricas superficiais e termométricas internas, em função do tempo de exposição ao fluxo térmico. Foi realizada também a avaliação da evolução das interfaces por comparação entre simulação e a amostra após o ensaio. Os resultados obtidos permitiram estimar as propriedades do comportamento ablativo dos materiais testados e validar o modelo teórico usado na simulação computacional para sua utilização em geometrias próximas às dos sistemas de proteção térmica utilizadas no setor aeroespacial / Materials used in space vehicles components, such as nozzles or surfaces exposed to atmospheric reentry, are subjected to thermally aggressive environments. This work presents studies involving the performance of composite materials used in thermal protection systems, through the exposure to plasma jets, where the heat fluxes are comparable to atmospheric reentry of space vehicle components. Samples of ablative carbon/phenolic composites were tested in the plasma tunnel of ITA’s (Aeronautics Institute of Technology) Plasma and Process Laboratory, by a plasma torch with a 50kW DC power source. The plasma tunnel operating parameters were optimized to reproduce the conditions close to the critical re-entry point of the space vehicles payloads developed by the IAE (Aeronautics and Space Institute). The samples in study were developed and manufactured in Brazil, from materials of special interest to IAE. For comparison, another material with well established properties such as teflon was also tested under the same ablative conditions. The mass loss and the specific mass loss rates of the samples, the surface radiometric and internal thermometric temperatures, as a function of the exposure time to the thermal flow, were determined. The evolution of the interfaces was also performed by comparison between simulation and the sample after the test. The results allowed to estimate the properties of the ablative behavior of the materials tested and to validate the theoretical model used in the computational simulation for its use in geometries close to the thermal protection systems used in the aerospace sector

In the design of a thermal protection system for atmospheric entry, aerothermal heating presents a major impediment to efficient heat shield design. Recombination of atomic species in the boundary layer results in highly exothermic surface-catalyzed recombination reactions and an increase in the heat flux experienced at the surface. The degree to which these reactions increase the surface heat flux is partly a function of the heat shield material. Characterization of the catalytic behavior of these materials takes place in experimental facilities, however there is a dearth of detailed computational models for the fluid dynamic and chemical behavior of such facilities. A numerical model coupling finite rate chemical kinetics and high temperature thermodynamic and transport properties with a computational fluid dynamics flow solver has been developed to model the chemically reacting flow in the inductively coupled plasma torch facility at the University of Vermont. Simulations were performed modeling the plasma jet for hybrid oxygen-argon and nitrogen plasmas in order to validate the models developed in this work by comparison to experimentally-obtained data for temperature and relative species concentrations in the boundary layer above test articles. Surface boundary conditions for wall temperature and catalytic efficiency were utilized to represent the different test article materials used in the experimental facility. Good agreement between measured and computed data is observed. In addition, a code-to-code validation exercise was performed benchmarking the performance of the models developed in this dissertation by comparison to previously published results. Results obtained show good agreement for boundary layer temperature and species concentrations despite significant differences in the codes. Lastly, a series of simulations were performed investigating the effects of recombination reaction rates and pressure on the composition of a nitrogen plasma jet in chemical nonequilibrium in order to better understand the composition at the boundary layer edge above a test article. Results from this study suggest that, for typical test conditions, the boundary layer edge will be in a state of chemical nonequilibrium, leading to a nonequilibrium condition across the entire boundary layer for test article materials with high catalytic efficiencies.

The Kentucky Re-entry Universal Payload System (KRUPS) is an adaptable testbed for atmosphere entry science experiments, with an initial application to thermal protection systems (TPS). Because of the uniqueness of atmospheric entry conditions that ground testing is unable to replicate, scientists principally rely on numerical models for predicting entry conditions. The KRUPS spacecraft, developed at the University of Kentucky, provides an inexpensive means of obtaining validation data to verify and improve these models. To increase the technology readiness level (TRL) of the spacecraft, two sub-orbital missions were developed. The first mission, KUDOS, launched August 13th, 2017 on a Terrier-Improved Malamute rocket to an altitude of ~150 km. The second mission, KOREVET, launched on March 25th, 2018 on the same type of rocket to an altitude of ~170 km. The chief purpose of both missions was to validate the spacecraft design, ejection mechanism, on-board power, data transmission, and data collection. After both missions, the overall TRL improved from 4 to 5 by validating most subsystems in a relevant environment. Both of these missions were invaluable preparation for the project's ultimate goal of releasing multiple experimental testbeds from the ISS.

Thermal Protection System (TPS) is required to shield an atmospheric entry vehicle against the high surface heating environment experienced during hypersonic flight. There are significant uncertainties in the tools and models currently used for the prediction of entry aeroheating and TPS material thermal response. These uncertainties can be reduced using experimental data. Analysis of TPS ground and flight data has been traditionally performed in a direct fashion. Direct analyses center upon comparison of the computational model predictions to data. Qualitative conclusions about model validity may be drawn based on this comparison and a limited number of model parameters may be iteratively adjusted to obtain a better match between predictions and data. The goal of this thesis is to develop a more rigorous methodology for the estimation of surface heating and TPS material response using inverse estimation theory. Built on theoretical developments made in related fields, this methodology enables the estimation of uncertainties in both the aeroheating environment and material properties from experimental temperature data. Unlike direct methods, the methodology developed here is capable of estimating a large number of independent parameters simultaneously and reconstructing the time-dependent surface heating profile in an automated fashion. This methodology is applied to flight data obtained from thermocouples embedded in the Mars Pathfinder and Mars Science Laboratory entry vehicle heatshields.

<p dir="ltr">A current engineering challenge is to understand and validate material systems capable of maintaining structural viability under the elevated temperature and environmental conditions of hypersonic flight. One aspect of this challenge is the joining of multiple materials with thermal expansion mismatch, which can lead to residual stress, resulting in debits in component lifetime under in-service loading. The focus of this work is a series of studies focused on a ceramic-metal composite (WC/Cu), a zirconia coating applied to a carboncarbon (C/C) composite, and a silicide (R512E) coating applied to a Nb-based alloy (C103). Each of these material systems are candidates for elevated temperature applications in which dissimilar constituents result in residual stress in the material. Each study leveraged experimental residual strain measurements, with the primary focus on the use of synchrotron X-ray diffraction, in conjunction with representative models, and microscopy to illuminate the active mechanisms in the development and evolution of residual stress in the bulk material. The combination of experimental and modeling predictions provides a framework to inform the viability and lifing of material systems exhibiting dissimilar expansion properties.</p>

Aviation industry is continuously growing especially for very long distance flights due to the globalisation of local economies around the world and the explosive economic growth in Asia. Reducing the time of intercontinental flights from 16-20 hours to 4 hours or less would therefore make the, already booming, ultra-long distance aviation sector even more attractive. To accomplish this drastic travel time reduction for civil transport, hypersonic cruise aircraft are considered as a potential cost-effective solution. Such vehicles should also be fuelled by liquid hydrogen, which is identified as the only viable propellant to achieve antipodal hypersonic flight with low environmental impact. Despite considerable research on hypersonic aircraft and hydrogen fuel, several major challenges should still be addressed before such airliner becomes reality. The current thesis is therefore motivated by the potential benefit of hydrogen-fuelled hypersonic cruise vehicles associated with their limited state-of-the-art.Hypersonic cruise aircraft require innovative structural configurations and thermal management solutions due to the extremely harsh flight environment, while the uncommon physical properties of liquid hydrogen, combined with high and long-term heat fluxes, introduce complex design and technological storage issues. Achieving hypersonic cruise vehicles is also complicated by the multidisciplinary nature of their design. In the scope of the present research, appropriate methodologies are developed to assess, design and optimize the thermo-structural model and the cryogenic fuel tanks of long-range hydrogen-fuelled hypersonic civil aircraft. Two notional vehicles, cruising at Mach 5 and Mach 8, are then investigated with the implemented methodologies. The design analysis of light yet highly insulated liquid hydrogen tanks for hypersonic cruise vehicles indicates an optimal gravimetric efficiency of 70-75% depending on insulation system, tank wall material, tank diameter, and flight profile. A combination of foam and load-bearing aerogel blanket leads to the lightest cryogenic tank for both the Mach 5 and the Mach 8 aircraft. If the aerogel blanket cannot be strengthened sufficiently so that it can bear the full load, then a combination of foam and fibrous insulation materials gives the best solution for both vehicles. The aero-thermal and structural design analysis of the Mach 5 cruiser shows that the lightest hot-structure is a titanium alloy construction made of honeycomb sandwich panels. This concept leads to a wing-body weight of 143.9 t, of which 36% accounts for the wing, 32% for the fuselage, and 32% for the cryogenic tanks. As expected, hypersonic thermal loads lead to important weight penalties (of more than 35%). The design of the insulated cold structure, however, demonstrates that the long-term high-speed flight of the airliner requires a substantial thermal protection system, such that the best configuration (obtained by load-bearing aerogel blanket) leads to a titanium cold design of only 4% lighter than the hot structure. Using aluminium 7075 rather than titanium offers a further weight saving of about 2%, resulting in a 135.4 t wing-body weight (with a contribution of 23%, 25%, 18% and 34% from the TPS, the wing, the fuselage, and the cryogenic tanks respectively). Given the design hypotheses, the difference in weight is not significant enough to make a decisive choice between hot and cold concepts. This requires the current methodologies to be further elaborated by relaxing the simplifications. Investigation of the thermal protection must be extended from one single point to different regions of the vehicle, and the TPS thickness and weight should be considered in the structural sizing of the cold design. More generally, the design process should be matured by including additional (static, dynamic and transient) loads, special structural concepts, multi-material configurations and other parameters such as cost and safety aspects. / Doctorat en Sciences de l'ingénieur et technologie / This thesis was conducted in co-tutelle between University of Sydney and Université Libre de Bruxelles.Professor Dries Verstraete was my supervisor at the University of Sydney (so as a member of SydneyUni), but is automatically registered here as a member of ULB because he worked at ULB almost ten years ago.Ben Thornber is also a member of the University of Sydney but the application does not save it for an unknown reason. / info:eu-repo/semantics/nonPublished

A porous silica –silica composite was processed with varying fiber diameters using the slurry moulding technique. The advantage of the process was that the density of the composite could be processed to the required levels. The reinforcements used were fibers obtained by leaching E-glass cloth, imported silica fibers with diameter <1.8µ and hollow silica fibers processed using sol-gel method. All the properties depend on the density of the composite. The compressive strength was measured in the perpendicular and parallel directions. Strength was high when the load axis was along the fiber direction. The composite with fine fibers (< 1.8 µ pure silica fibers) showed higher strength compared to the leached silica fibers. The thermal conductivity measurement on these composites showed an increase with temperature owing to the domination of radiation at high temperatures. As the vacuum level was approached, the thermal conductivity showed a decrease due to the absence of the convective part of the thermal transfer process. For use as a thermal protection system, it is important to measure the thermal response of these tiles in a simulated re-entry environment. Tests were done to measure this response for a given heat flux conditions at 38W/cm2 to 75W/cm2 and the backwall temperature was measured for various types of silica -silica composites. The role of impurities like sodium and B2O3 was also studied with respect to the conversion from amorphous to crystalline forms of SiO2. The severe increase in the coefficient of thermal expansion when SiO2 converted from amorphous to α– crystoballite was also measured. CSZP CSZP which belongs to the NZP family was processed using the co-precipitation technique. The influence of substituting the ‘P’ site with ‘Si’ atom was studied for its influence on thermal expansion – both at the bulk level by dilatometry and at the intrinsic level using high temperature XRD. For many anisotropic materials micro-cracking is a serious issue while cooling from the sintering temperature. It has been previously proved that this extent of micro-cracking depends on the particle size. Smaller the particle size is therefore preferred. One of the significant results obtained in this study was the successful use of microwaves to process crack free CSZP with fine grain size. CSZP with 95% density having a grain size as small as 1µ have been processed using microwave sintering. Dielectric property evaluation namely dielectric constant, dielectric loss and temperature coefficient of resonant frequency which are vital parameters required if this material is to be used as a candidate TPS have also been measured. The thermal conductivity of the sample was measured using Laser flash apparatus and was found to be 0.9 W/mk which provides an indication that this material can be used as a successful material for TPS. Finally a composite consisting of silica fiber with CSZP as matrix was processed and tested for heat flux. The low back wall temperature indicates that this material is a potential replacement for silica tile.

Les travaux présentés dans cette thèse CIFRE financée par AIRBUS Defence & Space s’intègrent dans une problématique de développement de nouveaux Systèmes de Protection Thermique (TPS) pour l’entrée atmosphérique. Ils se focalisent sur l’étude du transfert radiatif dans la zone d’ablation du TPS et son couplage avec le transfert conductif au travers de la matrice fibreuse de faible densité. Pour réaliser cette étude, il a été nécessaire d’évaluer les propriétés thermiques de ces matériaux, notamment les propriétés radiatives qui, contrairement aux conductivités thermiques, demeurent mal connues. La première étape de cette étude a donc visé à caractériser les propriétés optiques et radiatives de certains matériaux fournis par AIRBUS Defence & Space et par le CREE Saint-Gobain. Pour réaliser ces caractérisations, nous avons développé une méthode originale d’identification des propriétés radiatives basée sur des mesures de l’émission propre. Les spectres d’émission à haute température, réalisés sur des échantillons en fibre de silice ou en feutre de carbone nécessaires à l’identification, sont obtenus sur un banc de spectrométrie FTIR développé lors de ces travaux. Les échantillons sont chauffés à haute température à l’aide d’un laser CO2 et un montage optique permet de choisir entre la mesure du flux émis par l’échantillon ou un corps noir servant à l’étalonnage du banc. L’identification des propriétés repose sur la modélisation des facteurs de distribution du rayonnement calculés à l’aide d’une méthode de lancé de rayons Monte Carlo utilisant la théorie de Mie pour un cylindre infini pour le calcul des propriétés radiatives. Les températures identifiées sont comparées aux températures mesurées par pyrométrie au point de Christiansen dans le cas de la silice et montrent un bon accord avec ces dernières. Enfin la dernière partie de ce document est consacrée au couplage conduction-rayonnement dans ce type de milieu. Les échantillons ayant une très forte extinction, le modèle utilisé repose sur la définition d’une conductivité équivalente de Rosseland pour traiter les transferts radiatifs volumiques et ainsi simuler les champs de température au sein des échantillons dans les conditions de chauffage utilisées lors de l’identification. Dans le cas de la silice, cependant, les températures prédites par le modèle utilisant la conductivité équivalente de Rosseland, sont nettement supérieures à celles obtenues par identification ou par pyrométrie au point de Christiansen. Le fait que la conductivité équivalente de Rosseland ne fasse pas la distinction entre une forte extinction due à la diffusion ou à l’absorption est probablement la cause de cette différence. / The work presented in this thesis has been financed by AIRBUS Defence and Space. It is part of the development strategy of new Thermal Protection Systems (TPS) for atmospheric reentry purposes. The aim is to study the radiative transfer in the ablation zone of the TPS as well as the coupling of the radiative and conductive heat transfer in the low density fibrous matrix. To this end, radiative properties of the materials have to be evaluated since they are not well known. The first step of this study is therefore to characterize the optical and radiative properties of sample provided by AIRBUS Defence and Space and the CREE Stain-Gobain laboratory. Thus, an original identification method based on radiative emission measurement was developed to obtain the radiative properties. The needed emission spectra are measured on silica or carbon samples at high temperature with an experimental setup based on Fourrier Transformed InfraRed spectrometry. The samples are heated using a CO2 laser. An optical setup allows us to measure emission spectra on the sample or a black body used to calibrate the experiment. The identification process is based on the modeling of the radiative distribution factor computed by a Monte Carlo ray-tracing method. It uses Mie theory for infinite cylinder to compute the radiative properties. Temperature are also identified and, for silica, compared to the one measured by a Christiansen pyrometry technique. The last part of this study focuses on the coupled radiative/conductive heat transfer modeling in low density fibrous media. Samples being greatly absorbing, we used the Rosseland equivalent conductivity to model the radiative transfer in volume and obtain the thermal response of the samples in the conditions of the experimental setup used for the identification. For silica, predicted temperatures are superior to the identified ones or those measured with the Christiansen pyrometry technique. This is probably because the Rosseland equivalent conductivity makes no difference between extinction due to absorption and extinction due to scattering.

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.