The growing requirements to achieve hypersonic flights, as in the case of reentry vehicles, pose a serious challenge to the designers. This demands an understanding of the features of hypersonic flow and its effect on hypersonic vehicles. Hypersonic shock tunnels are one of the most widely used facilities for the purpose of obtaining valuable design data by conducting experiments on scaled down models. They are operated by conventional shock tubes by rupturing metal diaphragms placed between the driver and driven sections of the shock tube. Shock tunnels are being extensively used in spite of some of the drawbacks they possess. Due to the varying nature of metal diaphragm rupture, reproducibility of the experiment results is difficult to obtain. Damage to model and inner surface of the shock tube can happen when the diaphragm petal breaks away from the diaphragm. Lastly the time consuming diaphragm replacement process is not desired in applications which require quick loading of shock waves on the specimen. All these disadvantages call for the replacement of the diaphragm mode of operation with a diaphragmless mode of operation for the generation of shock waves. The main objective of the present study is to design and demonstrate the working of a diaphragmless hypersonic shock tunnel. The motivation for the present study comes from the fact that the diaphragmless operation of a shock tunnel has not been reported so far in the open literature. All the research works carried out deal with diaphragmless drivers operating only a shock tube. In the present work, the conventional metal diaphragm is substituted by fast acting pneumatic valves which serve the purpose of quickly opening the driven section of the shock tube to allow the driver gas to rush in, resulting in the formation of a shock wave. To design a diaphragmless driver, a detailed study of the shock formation process is accomplished which helps in understanding the effect of valve opening time on the shock formation distance. Also the theoretical basis for the design of a pneumatic cylinder is understood. Following the theoretical studies, three types of diaphragmless drivers are designed and tested. The first setup incorporates a rubber membrane, which acts as a valve. The rubber membrane when bulged closes the mouth of the driven section and on retraction the driven section is opened to the driver gas. The second and the third setups utilise two different types of double acting pneumatic cylinders. Experimental results of the three diaphragmless drivers operating a shock tube are analysed and compared with the ideal shock tube theory. Better repeatability in terms of shock Mach number is shown with all three diaphragmless shock tubes when compared with a conventionally operated shock tube. Finally, the best among the three systems is identified to operate the hypersonic shock tunnel 2 (HST2) facility of the Shock Waves laboratory, IISc. Demonstration of the working of the diaphragmless shock tunnel is shown by performing heat transfer measurements on a 3 mm backward facing step flat plate model. The experimental results are compared with those obtained in a conventional shock tunnel. CFD studies on diaphragmless shock tube model are done to have an idea on the flow in the shock tube there by identifying the shock formation distance. ANSYS-CFX package is used for this purpose. Further, results from the numerical simulation of hypersonic flow over the backward facing step model are compared with the experimental results thus validating the code.

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

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.

L’exploration des quatre planètes géantes gazeuses, Jupiter, Saturne, Neptune et Uranus, est importante pour comprendre l’évolution de notre système solaire et plus généralement de l’univers. Les sondes entrant dans l’atmosphère des géantes gazeuses ont des vitesses de 20 à 50 km/s, largement supérieures aux vitesses d’entrée atmosphérique sur les autres planètes du système solaire. Il s’agit d’un problème complexe car les conditions d’entrées sont brutales et les vitesses associées dépassent largement les capacités des installations d’essai au sol actuelles. Cette thèse examine la possibilité de simuler expérimentalement les conditions d’entrées proposées pour Uranus et Saturne à 22.3 et 26.9 km/s avec un tube d’expansion à piston libre. D’abord, la possibilité de simuler les conditions directement en recréant la vitesse d’entrée réelle a été étudiée. Il a été trouvé qu’il était possible de simuler l’entrée d’Uranus mais seulement avec de grandes incertitudes. Pour cette raison, il a été proposé d’utiliser une substitution du gaz d’essai établie, dans lequel soit le pourcentage d’hélium dans l’atmosphère H2/He est augmenté, soit l’hélium est remplacé par du néon, un gaz noble plus lourd. Cela permet de simuler uniquement les conditions postchoc des entrées. Théoriquement, il a été constaté que ces substitutions permettaient de simuler l’entrée Uranus ou Saturne, ce qui a été confirmé expérimentalement à l’aide d’hélium. Notant l’intérêt actuel d’envoyer des sondes d’entrée atmosphérique vers ces deux planètes, cette étude a démontré que les capacités expérimentales requises sont disponibles pour la réalisation d’expériences simulées avec les modèles d’essais. / Exploration of the four gas giant planets, Jupiter, Saturn, Uranus, and Neptune, is important for understanding the evolution of both our solar system and the greater universe. Due to their size, flight into the gas giants involves atmospheric entry velocities between 20 and 50 km/s. This is a complex issue because the entry conditions are harsh but the related velocities are mostly beyond the capabilities of current ground testing facilities. As such, this thesis examines the possibility of experimentally simulating proposed Uranus and Saturn entries at 22.3 and 26.9 km/s in a free piston driven expansion tube, the most powerful type of impulse wind tunnel. Initially, the possibility of simulating the conditions directly by re-creating the true flight velocity was investigated. It was found to be possible to simulate the 22.3 km/s Uranus entry, but not without large uncertainties in the test condition. For this reason, it was proposed to use an established test gas substitition where the percentage of helium in the H2/He atmosphere is increased, or the helium is substituted for the heavier noble gas neon. This allows just the post-shock conditions of the entries to be simulated. Theoretically it was found that these substitutions allowed both Uranus or Saturn entry to be simulated, which was confirmed experimentally using helium. Noting the current interest in sending atmospheric entry probes to both of these planets, this study has demonstrated that the required experimental capabilities are available for performing simulated experiments using test models.

No description available.

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.

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

The 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.

Shock Boundary Layer Interactions (SBLIs) and shock-shock interactions are some of the most fundamental problems in high speed aerodynamics. These interactions are of particular importance in scramjet intakes at hypersonic speeds. In hypersonic own with strong SBLI accompanied by own separation, large separation bubbles can form due to high impinging shock strengths. While experiments involving large separation lengths for the impinging shock boundary layer interactions near sharp leading edge are well documented in the literature, only few investigations on the effect of leading edge bluntness on the interactions are studied. In the present study, experiments were carried out to study the role of leading edge bluntness on the impinging shock boundary layer interactions. An oblique shock generated by a wedge (wedge angle 31 degrees) is made to impinge on a at plate (length 200 mm) over which a boundary layer develops. Different leading edge inserts were used on a at plate to get either a sharp or a blunt (radii from 2 to 8 mm) leading edge. The position of the at plate was moved horizontally with respect to the wedge to vary the shock impingement location relative to the leading edge. Experiments were carried out at two freestream conditions - Mach 5.88 (total enthalpy of 1.26 MJ/kg and freestream Reynolds number of 3.85 million per meter) and Mach 8.54 (total enthalpy of 1.85 MJ/kg and freestream Reynolds number of 1.41 million per meter). The various features of the interaction along with different parameters were obtained from schlieren visualizations and surface pressure measurements. The schlieren visualization was used to obtain the separation length, while the reattachment pressure was obtained from the surface pressure distribution. From the present experimental study, a reduction in separation length was observed with an increase in leading edge bluntness. It was also seen that the sharp leading edge had the maximum separation length. Correlations for the separation length and the reattachment pressure have been proposed for these experimental conditions. Numerical simulations were also carried out using commercial software and they had a qualitative agreement with the experiments.