Knowledge of flow separation is very essential for proper understanding of both external and internal aerothermodynamics of bodies. Because of unique flow features such as thick boundary layers, merged shock layers, strong entropy layers, flow separation in the flow field of bodies at hypersonic speeds, is both complex as well as interesting. The problem of flow separation is further complicated at very high stagnation enthalpies because of the real gas effects. Notwithstanding the plethora of information available in open literature even for simple geometric configurations the experimentally determined locations of flow separation and re-attachment points do not match well with the results from the computational studies even at hypersonic laminar flow conditions. In this backdrop the main aim of the present study is to generate a reliable experimental database of classical separated flow features around generic configurations at hypersonic laminar flow conditions. In the present study, flow visualization using high speed camera, surface convective heat transfer rate measurements using platinum thin film sensors, and direct skin friction measurements using PZT crystals have been carried out for characterizing the separated flow field around backward facing step, double cone and double wedge models. The numerical simulations by solving the Navier-Stokes equations have also been carried out to complement the experimental studies. The generic models selected in the present study are simple configurations, where most of the classical hypersonic separated flow features of two-dimensional, axi-symmetric and three dimensional flow fields can be observed. All the experiments are carried out in IISc hypersonic shock tunnel (HST2) at Mach 5.75 and 7.6. For present study, helium and air have been used as the driver and test gases respectively. The high speed schlieren flow visualization is carried out on backward facing step (2 and 3 mm step height), double cone (semi-apex angles of 150/350 and 250/680) and double wedge (semi-apex angles of 150/350) models by using high speed camera (Phantom 7.1). From the visualized shockwave structure in the flow field the flow reattachment point after separation has been clearly identified for backward facing step, double cone and double wedge models at hypersonic Mach numbers while the separation point could not be clearly identified because of the low free stream density in shock tunnels. However the flow visualization studies helped clearly identifying the region of flow separation on the model. Based on the results from the flow visualization studies both the physical location and distribution of platinum thin film gauges was finalized for the heat transfer rate measurements. Surface heat transfer rates along the length of two backward facing step (2 and 3 mm step height) models have been measured using platinum thin film gauges deposited on Macor substrate. The Eckert reference temperature method is used along the flat plate for predicting the heat flux distribution. Theoretical analysis of heat flux distribution down stream of the backward facing step model has been carried out using Gai’s dimensional analysis. The study reveals for the first time that at moderate stagnation enthalpy levels (~2 MJ/kg) the hypersonic separated flow around a backward facing step reattaches rather smoothly without any sudden spikes in the measured values of surface heat transfer rates. Based on the measured surface heating rates on the backward facing step, the reattachment distance was estimated to be approximately 10 and 8 step heights downstream of 2 and 3 mm step respectively at nominal Mach number of 7.6. Convective surface heat transfer experiments have also been carried out on axi-symmetric double cone models (semi-apex angles of 15/35 and 25/68), which is analogous to the Edney’s shock interactions of Type VI and Type IV respectively. The flow is unsteady on the double cone model of 25/68 and measured heat flux is not constant. The heat transfer experiments were also carried out on the three-dimensional double wedge model (semi-apex angles of 15/35). The separation and reattachment points have been clearly identified from the experimental heat transfer measurements. It has been observed that the measured heat transfer rates on the double wedge model is less than the double cone model (semi-apex angles of 150/350) for the identical experimental conditions at the same gauge locations. This difference could be due to the three-dimensional entropy relieving effects of double wedge model. PZT-5H piezoelectric based skin friction gauge is developed and used for direct skin friction measurements in hypersonic shock tunnel (HST2). The bare piezoelectric PZT-5H elements (5 mm × 5 mm with thickness of 0.75 mm) polarized in the shear mode have been used as a skin friction gauge by operating the sensor in the parallel shear mode direction. The natural frequency of the skin friction sensor is ~80 kHz, which is suitable for impulse facilities. The direct skin friction measurements are carried out on flat plate, backward facing step (2 mm step height) and double wedge models. The measured value of skin friction coefficient (integrated over an area of 25 sq. mm; sensor surface area) at a distance of 23 mm from the leading edge of the sharp leading edge backward facing step model is found to be ~ 0.0043 while it decreases to ~ 0.003 at a distance of 43 mm from the leading edge at a stagnation enthalpy of ~ 2MJ/kg. The measured skin friction matches with the Eckert reference temperature within ± 10%. The skin friction coefficient is also measured on the double wedge at a distance of 73 mm from the tip of the first wedge along the surface and is found to be 4.56 × 10-3. Viscous flow numerical simulations are carried out on two-dimensional backward facing step, axi-symmetric double cone and three-dimensional double wedge models using ANSYS-CFX 5.7 package. Navier-Stokes Simulations are carried out at Mach 5.75 and 7.6 using second order accurate (both in time and space) high resolution scheme. The flow is assumed to be laminar and steady throughout the model length except on the double cone (semi-apex angles of 250/680) model configuration, which represents the unsteady flow geometry. Analogous Edney Type VI and Type IV shock interactions are observed on double cone, double wedge (semi-apex angles of 150/350) and double cone (semi-apex angles of 250/680) models respectively from the CFD results. Experimentally measured convective heat transfer rates on the above models are compared with the numerical simulation results. The numerical simulation results matches well with the experimental heat transfer data in the attached flow regions. Considerable differences are observed between the measured surface heat transfer rates and numerical simulations both in the separated flow region and on the second cone/wedge surfaces. The separation and reattachment points can be clearly identified from both experimental measurements and numerical simulations. The results from the numerical simulations are also compared with results from the high speed flow visualization experiments. The experimental database of surface convective heating rates, direct skin friction coefficient and shockwave structure in laminar hypersonic flow conditions will be very useful for validating CFD codes

<div>Many important measurements of low-amplitude instabilities related to hypersonic laminar-turbulent boundary-layer transition have been successfully performed with 1-MHz PCB132 pressure sensors. However, there is large uncertainty in measurements made with PCB132 sensors due to their poorly understood response at high frequency. The current work continues efforts to better characterize the PCB132 sensor with a low-pressure shock tube, using the pressure change across the incident shock as an approximate step input. </div><div> </div><div> New vacuum-control valves provide precise control of pre-run pressures in the shock tube, generally to within 1\% of the desired pressure. Measurements of the static-pressure step across the shock made with Kulite sensors showed high consistency for similar pre-run pressures. Skewing of the incident shock was measured by PCB132 sensors, and was found to be negligible across a range of pressure ratios and static-pressure steps. Incident-shock speed decreases along the shock tube, as expected. Vibrational effects on the PCB132 sensor response are significantly lower in the final section of the driven tube.</div><div> </div><div> Approximate frequency responses were computed from pitot-mode responses. The frequency-response amplitude varied by a factor of 5 between 200--1000 kHz due to significant resonance peaks. Measurements with blinded PCB132 sensors indicate that the resonances in the frequency response are not due to vibration. </div><div> </div><div> Using the approximate frequency response measured with the shock tube to correct the spectra of wind-tunnel data produced inconclusive results. Correcting pitot-mode PCB132 wind-tunnel data removed a possible resonance peak near 700 kHz, but did not agree with the spectrum of a reference sensor in the range of 11--100 kHz. </div>

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ABSTRACT A theoretical and experimental study of the vibrational and chemical relaxation rates of diatomic gases, in flows behind shock waves and along nozzles,is made here. ¶ The validity of the conventional relaxation rate models, which are generally used to analyse experiments, is tested by developing a detailed microscopic description of the diatomic relaxation processes. Assuming the diatomic molecules to be represented by the anharmonic Morse Oscillator, the vibrational Master equation, which describes the time variation of each vibrational energy level population, is constructed by allowing one-quantum vibration to translation (V-T) energy exchanges and vibration to vibration (V-V) energy exchanges between the molecules. Dissociation and recombination are allowed to occur from, and to, the uppermost vibrational level. Solving the Master equation, it is found that a number of effects are explained by the inclusion of V-V transitions. In particular it is found that V-V energy exchanges cause the induction time for H2 dissociation to be increased; suggest that the linear rate law, for H2 and Ar mixtures, fails for a H2 mole fraction above 20%; give an acceleration of vibrational excitation as equilibrium is approached for H2 and N2; cause the vibrational temperature to be lower than the value found without V-V transitions for vibrational de-excitation in nozzle flows of H2 and N2, and conversely for recombination of H2 in nozzle flows. The most important result is the demonstration that conventional nozzle flow calculations, with shock-tube-determined dis-sociation and vibrational excitation rates, appear to be valid for the recombining and vibrationally de-excitating flows considered. ¶ The dissociation rates of undiluted nitrogen are measured in the free-piston shock tube DDT, using time-resolved optical interferometry, over a temperature range of 6000-14000K and confirm the strong temperature dependence of the pre-exponential factor observed by Hanson and Baganoff (1972). ¶ The vibrational de-excitation and excitation rates are determined in the small free-piston shock tunnel T2 over temperature ranges of 2000-4000K and 7000-10300K, respectively, by measuring the shock angles and curvatures, from optical interferograms, of flow over an inclined flat plate in the nonequilibrium nozzle flow. The de-excitation rate is found to be within a factor of ten of the excitation rate, while the excitation rate of N2 by collision with N is found to be less than about 50 times the excitation rate of N2 by N2. The dissociation rates of nitrogen, in the flow behind a shock attached to a wedge, are investigated in the large free-piston shock tunnel, using the shock curvature technique. The discrepancy, reported by Kewley and Hornung (1974b), between theory and experiment at the highest enthalpy is found to be resolved by including the measured helium contamination (Crane 1975) in the free-stream. Reasonable agreement is obtained between experimental shock curvatures and calculations using accepted dissociation rates.

For flights beyond Mach 6 ramjets are inefficient engines due to huge total pressure loss in the normal shock systems, combustion conditions that lose a large fraction of the available chemical energy due to dissociation and high structural loads. However if the flow remains supersonic inside the combustion chamber, the above problems could be alleviated and here the concept of SCRAMJET(supersonic combustion ramjet) comes into existence. The scramjets could reduce launching cost of satellites by carrying only fuel and ingesting oxygen from atmospheric air. Further applications could involve defense and transcontinental hypersonic transport. In the current study an effort is made to achieve supersonic combustion in a ground based short duration test facility(combustion driven shock-tunnel), which in addition to flight Mach number can simulate flight Reynolds number as well. In this study a simple method of injection i.e. wall injection of the fuel into the combustion chamber is used. The work starts with threedimensional numerical simulation of a non-reacting gas(air) injection into a hypersonic cross-flow of air to determine the conditions in which air penetrates reasonably well into the cross-flow. Care is taken so that the process does not induce huge pressure loss due to the bow shock which appears in front of the jet column. The code is developed in-house and parallelized using OpenMp model. This is followed by experiments on air injection into a hypersonic cross-flow of air in a conventional shock-tunnel HST2 existing in IISc. The most tricky part is synchronization of injection with start of test-flow in such a short duration(test time 1 millisecond) facility. Next part focuses on numerical simulations to determine the free-stream conditions, mainly the temperature and pressure of air, so that combustion takes place when hydrogen is injected into a supersonic cross-flow of air. The simulations are two-dimensional and includes species conservation equations and source terms due to chemical reactions in addition to the Navier-Stokes equations. This code is also built in-house and parallelized because of more number of operations with the inclusion of species conservation equations and chemical non-equilibrium. However, the predicted conditions were not achievable by HST2 due to low stagnation conditions of HST2. Therefore, a new shock-tunnel which could produce the required conditions is built. The new tunnel is a combustion driven shock-tunnel in which the driver gas is at higher temperature than conventional shock-tunnel. The driver gas is basically a mixture of hydrogen, oxygen and helium at a mole ratio of 2:1:10 initially. The mixture is ignited by spark plugs and the hydrogen and oxygen reacts releasing heat. The heat released raises the temperature of the mixture which is now predominantly helium and small fractions of water vapour and some radicals. The composition of the driver gas and initial pressure are determined through numerical simulations. Experiments follow in the new tunnel on hydrogen injection into a region of supersonic cross-flow between two parallel plates with a wedge attached to the bottom plate. The wedge reduces the hypersonic free-stream to Mach 2. A high-speed camera monitors the flow domain around injection point and sharp rise in luminosity is observed. To ascertain whether the luminosity is due to combustion or not, two more driven gases namely nitrogen and oxygen-rich air are used and the luminosity is compared. In the first case, the free-stream contains no oxygen and luminosity is not observed whereas in the second case higher luminosity than air driver case is visible. Additionally heat-transfer rates are measured at the downstream end of the model and at a height midway between the plates. Similar trend is observed in the relative heat-transfer rates. Wall static pressure at a location downstream of injection port is also measured and compared with numerical simulations. Results of numerical simulations which are carried out at the same conditions as of experiments confirm combustion at supersonic speed. Experiments and numerical simulations show presence of supersonic combustion in the setup. However, further study is necessary to optimize the parameters so that thrust force could be generated efficiently.

The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body. Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.

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A body in a hypersonic flow field will experience very high heating especially during re-entry. Conventionally this problem is tackled to some extent by the use of large angle blunt cones. At the cost of increased drag, the heat transfer rate is lower over most parts of the blunt body, except in a region around the stagnation point. Thus even with blunt cones, management of heat transfer rates and drag on bodies at hypersonic speeds continues to be an interesting research area. Various thermal protection systems have been proposed in the past, like heat sink cooling, ablation cooling and aerospikes. The ablative cooling system becomes extremely costly when reusability is the major concern. Also the shape change due to ablation can lead to issues with the vehicle control. The aerospikes themselves may become hot and ablate at hypersonic speeds. Hence an alternate form of cooling system is necessary for hypersonic flows, which is more feasible, cost effective and efficient than the conventional cooling systems. Injection of a mass of cold fluid into the boundary layer through the surface is one of the potential cooling techniques in the hypersonic flight corridors. These kinds of thermal protection systems are called mass transfer cooling systems. The injection of the mass may be through discrete slots or through a porous media. When the coolant is injected through a porous media over the entire surface, the coolant comes out as a continuous mass. Such a cooling system is also referred as “transpiration cooling system”. When the fluid is injected through discrete slots, the system is called as “film cooling system”. In either case, the coolant absorbs the incoming heat through its rise in enthalpy and thus modifies the boundary layer characteristics in such a way that the heat flow rate to the surface is less. Injection of a forward facing jet (opposite to the freestream direction) from the stagnation point of a blunt body can be used for mitigating both the aerodynamic drag and heat transfer rates at hypersonic Mach numbers. If the jet has enough momentum it can push the bow shock forward, resulting in reduced drag. This will also reduce heat transfer rate over most part of the body except around the jet re-attachment region. A reattachment shock impinging on the blunt body invariably increases the local heat flux. At lower momentum fluxes the forward facing jet cannot push the bow shock ahead of the blunt body and spreads easily over the boundary layer, resulting in reduced heat transfer rates. While the film cooling performance improves with mass flow rate of the jet, higher momentum flow rates can lead to a stronger reattachment leading to higher heat transfer rate at the reattachment zone. If we are able to reduce the momentum flux of the coolant for the same mass flow rate, the gas coming out can easily spread over the boundary layer and it is possible to improve the film cooling performance. In all the reported literature, the mass flow rate and the momentum flux are not varied independently. This means, if the mass flow rate is increased, there is a corresponding increase in the momentum flux. This is because the injection (from a particular orifice and for a particular coolant gas) is controlled only by the total pressure of injection and free stream conditions. The present investigation is mainly aimed at demonstrating the effect of reduction in momentum of the coolant (injected opposing a hypersonic freestream from the stagnation point of a blunt cone), keeping the mass flow rate the same, on the film cooling performance. This is achieved by splitting a single jet into a number of smaller jets of same injection area (for same injection total pressure and same free stream conditions). To the best of our knowledge there is no report on the use of forward facing micro-jet array for film cooling at hypersonic Mach numbers. In this backdrop the main objectives of the present study are: • To experimentally demonstrate the effect of splitting a single jet into an array of closely spaced smaller micro-jets of same effective area of injection (injected opposite to a hypersonic freestream from the stagnation zone), on the reduction in surface heat transfer rates on a large angle blunt cone. · Identifying various parameters that affect the flow phenomenon and doing a systematic investigation of the effect of the different parameters on the surface heat transfer rates and drag. Experimental investigations are carried out in the IISc hypersonic shock tunnel on the film cooling effectiveness. Coolant gas (nitrogen and helium) is injected opposing hypersonic freestream as a single jet (diameter 2 mm and 0.9 mm), and as an array of iv micro jets (diameter 300 micron each) of same effective area (corresponding to the respective single jet). The coolant gas is injected from the stagnation zone of a blunt cone model (58o apex angle and nose radius of 35 mm). Experiments are performed at a flow freestream Mach number of 5.9 at 0o angle of attack, with a stagnation enthalpy of 1.84 MJ/Kg, with and without injections. The ratios of the jet stagnation pressure to the pitot pressure (stagnation pressure ratio) used in the present study are 1.2 and 1.45. Surface convective heat transfer measurements using platinum thin film sensors, time resolved schlieren flow visualization and aerodynamic drag measurements using accelerometer force balance are used as flow diagnostics in the present study. The theoretical stagnation point heat transfer rate without injection for the given freestream conditions for the test model is 79 W/cm2 and the corresponding aerodynamic drag from Newtonian theory is 143 N. The measured drag value without injection (125 N) shows a reasonable match with theory. As the injection is from stagnation zone it is not possible to measure the surface heat transfer rates at the stagnation point. The sensors thus are placed from the nearest possible location from the stagnation point (from 16 mm from stagnation point on the surface). The sensors near the stagnation point measures a heat transfer rate of 65 W/cm2 on an average without any injection. Some of the important conclusions from the study are: • Up to 40% reduction in surface heat transfer rate has been measured near the stagnation point with the array of micro jets, nitrogen being the coolant, while the corresponding reduction was up to 30% for helium injection. Considering the single jet injection, near the stagnation point there is either no reduction in heat transfer rate or a slight increase up to 10%. · Far away from stagnation point the reduction in heat transfer with array of micro-jets is only slightly higher than corresponding single jet for the same pressure ratio. Thus the cooling performance of the array of closely spaced micro jets is better than the corresponding single jet almost over the entire surface. • The time resolved flow visualization studies show no major change in the shock standoff distance with the low momentum gas injection, indicating no major changes in other aerodynamic aspects such as drag. · The drag measurements also indicate that there is virtually no change in the overall aerodynamic drag with gas injection from the micro-orifice array. · The spreading of the jets injected from the closely spaced micro-orifice array over the surface is also seen in the visualization, indicating the absence of a region of strong reattachment. · The reduction in momentum flux of the injected mass due to the interaction between individual jets in the case of closely spaced micro-jet array appears to be the main reason for better performance when compared to a single jet. The thesis is organized in six chapters. The importance of film cooling at hypersonic speeds and the objectives of the investigation are concisely presented in Chapter 1. From the knowledge of the flow field with counter-flow injection obtained from the literature, the important variables governing the flow phenomena are organized as non-dimensional parameters using dimensional analysis in Chapter 2. The description of the shock tunnel facility, diagnostics and the test model used in the present study is given in Chapter 3. Chapter 4 describes the results of drag measurements and flow visualization studies. The heat transfer measurements and the observed trends in heat transfer rates with and without coolant injection are then discussed in detail in Chapter 5. Based on the obtained results the possible physical picture of the flow field is discussed in Chapter 6, followed by the important conclusions of the investigation.