The main aim of this thesis was to understand the mechanism behind the classical transition scenario inside the boundary layer over an airfoil and eventually attempting to control this transition utilizing passive devices for transition delay. The initial objective of analyzing the transition phenomenon based on TS wave disturbance growth was conducted at 90 Hz using LDV and CTA measurement techniques at two different angles of attack. This was combined with the studies performed on two other frequencies of 100 and 110 Hz, in order to witness its impact on the neutral stability curve behavior. The challenges faced in the next phase of the thesis while trying to control the transition location, was to understand and encompass the effect of adverse pressure gradient before setting up the passive control devices, which in this case was miniature vortex generators. Consequently, several attempts were made to optimize the parameters of the miniature vortex generators depending upon the streak strength and stability. Finally, for 90 Hz a configuration of miniature vortex generators have been found to successfully stabilize the TS wave disturbances below a certain forcing amplitude, which also led to transition delay.

No description available.

Fluid dynamicists are always in need of innovative instruments for flow velocity measurements. An ideal instrument would be non-intrusive, have a very fine spatial resolution as well as a very fine temporal resolution, be able to measure three-components of velocity, and be compact. Through recent advancements, laser Doppler velocimetry can now meet all of those requirements making it an important part of aerodynamicist's research toolbox. The first paper presented in this manuscript style thesis explains the development of an advanced three-velocity component, spatially-resolving laser-Doppler velocimetry (LDV) system for highly resolved velocity measurements in situations with limited optical access. The new instrument, a next generation version of the previously developed 'comprehensive' LDV technology, enables measurements of three components of velocity and particle position in the axial direction all through a single transceiving lens. Described here is the design process and the final design for the 'compact, comprehensive' LDV (Comp²LDV). The probe was designed to achieve ± 10 micron root-mean-square uncertainties in axial particle position, which combined with the long measurement volume, allow researchers to obtain a three-velocity-component velocity statistics profiles over a span of approximately 1.5mm without the need for traversing. Results from measurements in a flat plate turbulent boundary layer very near the wall have compared favorably to data from previous studies. The second paper focuses on the motion and evolution of coherent structures in supersonic jet flows and how that relates to the intense noise the flows generate. As a preliminary study to experimentally address these relationships, novel non-intrusive measurements using two-component laser Doppler velocimetry (LDV) have been conducted at exceptionally high data rates to lend insight into the statistical behavior of noise-generating flow structures. A new heated supersonic jet facility has been constructed to provide supersonic flow at total temperatures ratios (T₀/Tₐ) up to 3. In the present work, the instrumentation is validated via comparison of LDV measurements along the centerline of a screeching cold jet with microphone and high-speed shadowgraph results. Reynolds stress spectra are presented for an over-expanded case (nozzle pressure ratio of 3.2) of a design Mach number 1.65 nozzle operated cold (T₀/Tₐ = 1). A preliminary study was then conducted in the near-nozzle shear layer, up to x/d = 4.0, at design nozzle pressure ratio (4.58) and total temperature ratio of 2.0. Results are presented for Reynolds stress time-delay correlations and power spectra at Re_d = 1.1M for this case. The stream-wise Reynolds normal stress spectra are compared with published spectral behavior reported by other researchers, indicating a similar spectral shape in the downstream stations as previously measured with LDV and hot wire anemometry for cold jets, but which differ in shape from density-based techniques. / Master of Science

The linear instability of laterally diverging/converging flows is an important aspect towards understanding the laminar-transition process in many viscous flows. In this work the linear instability of constant pressure laterally diverging/converging flow has been investigated. The laminar velocity field for laterally diverging/converging flows, under the source/sink approximation, has been reduced to two-dimensional flows. This reduction is alternative to the Mangier transformation used earlier. For a constant pressure laterally strained flow, the laminar velocity is found to be governed by the Blasius equation for flow over a flat plate. The non-parallel linear instability of constant pressure laterally strained flows has been examined. The instability equation is found to be same as that for the Blasius flow. This implies that the stability is same as that for the Blasius flow. A lateral divergence/convergence is shown to alter the Reynolds number from that in a two-dimensional flow. The instability of a laterally converging/diverging flow thus can be obtained from the available results for the Blasius flow by scaling the Reynolds numbers. This leads to the result that while a diverging flow is more unstable than the Blasius flow, a converging flow is more stable. Some additional relevant results are also presented.

The present thesis is an experimental study of some aspects of unsteady two dimensional boundary layers subject to adverse pressure gradient. An adverse pressure gradient usually leads to boundary layer separation or an instability which may result in transition to turbulence. Unsteady boundary layer separation is not yet fully understood and there is no specific criterion proposed in literature for its occurrence. The details of separation depend on the Reynolds number, the geometry of the body (streamlined or bluff) and the type of imposed unsteady motion (impulsive, oscillatory etc.). Similarly there are many unknowns with respect to instability and transition in unsteady boundary layers, especially those having a streamwise variation. For unsteady flows it is useful to break up the pressure gradient term in the unsteady boundary layer equation into two components:(Formula) is the velocity at the edge of the boundary layer. The first term of the right hand side of this equation may be called the temporal component (Πt) which signifies acceleration or deceleration in time of the free stream and the second term is the spatial component (Πx) which represents the spatial or convective acceleration of the free stream. Many of the studies on instability in unsteady flows found in literature are carried out in straight tubes or channels, where the Πx term is absent. However, in many cases, especially in biological systems both terms are present. An example is the unsteady flow over the moving body of a fish. To study the effects of Πt and Πx on unsteady separation and instability we have built an unsteady water tunnel where the two components can be systematically varied. The flow is created by a controlled motion of a piston. By a suitable combination of the geometry of the model and the piston motion, different types of separation bubbles may be generated. In our studies the piston motion follows a trapezoidal variation: constant acceleration from rest, followed by constant velocity and then deceleration to zero velocity. We have chosen two geometries. One is a bluff body and thus has a high value of Πx and other is a small angle diffuser with a divergence angle 6.2° and thus having a small value of Πx. Upstream and downstream of the diffuser are long lengths of constant cross section. We have performed experiments with the above mentioned geometries placed in the tunnel test section. Flow is visualized using the laser induced fluorescence technique by injecting a thin layer of fluorescein dye on the test wall. Numerical simulations have been done using the software FLUENT. Boundary layer parameters like boundary layer, displacement and momentum thicknesses are calculated from the simulations and used to analyze the experimental results. For the flow in the diffuser, quasi-steady stability analysis of the instantaneous velocity profiles gives a general idea of stability behavior of the flow. Two types of experiments have been done with the bluff body. One is the unsteady boundary layer separation and the formation of the initial vortex for a flow that is uniformly accelerated from rest. We have found some scalings for the formation time (tv) of the separation vortex. The second type of experiment was to study the vortex shedding from the separating shear layer after the boundary layer has fully separated. At high enough Reynolds number shear layer vortices are seen to shed from the separation bubble. The Strouhal number based on the momentum thickness and the velocity at the edge of the boundary layer just upstream of the separation point is found to vary between 0.004 and 0.008. This value is close to the Strouhal number value of 0.0068 found in laminar separation bubbles on a flat plate. The second part of the study concerns with the evolution of the flow in the small angle diffuser with a mild variation of the spatial component of the pressure gradient. From the experimental visualizations we have found that the ratio of Πx and Πt at the start of the deceleration phase of the piston motion is an important parameter that determines the type of instability. This value of Πx/Πt is controlled by controlling the piston deceleration: a large deceleration gives a low Πx/Πt value and a low deceleration gives a large Πx/Πt value. Three types of instabilities have been observed in our experiments. In Type I, the first vortex forms at the maximum pressure gradient point (MPGP) and which grows disproportionately with time. However, instability vortices are seen later at other locations around the MPGP. In type II an array of vortices over a certain length are observed; the vortices grow with time. In Type III, which we observe for low decelerations, we observe initial vortices only in the diffuser section in the deceleration phase of the piston motion. Type III instability is similar to the one observed in dynamic stall experiments. In all cases the instability is very localized - it occurs only over some length of the boundary layer. Transition to turbulence, which is also localized, is observed at higher Reynolds numbers. The non-dimensional time for vortex formation is not very different from that found in straight channel experiments. Quasi-steady linear stability analyses for the boundary layer at the MPGP both for the top and the bottom walls show that the flow is absolutely unstable for some cases. In summary, the thesis looks at in a unified way the separation and instability of unsteady boundary layers with reverse flow. It is hoped that the results will be useful in predicting and understanding onset of separation and instability in practically occurring unsteady flows.

No description available.

The interaction of a liquid drop with a solid surface is being actively studied to understand practically encountered scenarios such as the impact of fuel spray droplets onto the walls of engine combustion chamber, the formation of thermal barrier coating on the surfaces of turbine blades, the process of ink-jet printing, etc. The surface topography of solid surface is one of the major parameters influencing the dynamics of drop-surface interaction process. Understanding the precise role of surface topography features such as micro asperities and grooves on the spreading and receding processes of impacting liquid drops is crucial for the improvement in abovementioned applications. Recent developments in the fabrication of micro- and nano-structures on solid surfaces provide fabulous opportunities to investigate the role of single/multiple micro asperities and grooves on the dynamics of impacting drops. The thesis deals with an experimental work on the impact of water drops on stainless steel surfaces comprising unidirectional parallel grooves. A group of six target grooved surfaces covering a wide range of surface wettability were considered. The target surfaces were prepared using the techniques of photolithography, electro discharge machining, and laser machining. Scanning electron microscope and optical surface profilometer were used to characterize the groove texture geometrical parameters of the target grooved surfaces. The experiments of drop impact were carried out in an experimental apparatus consisting of a liquid drop generator, a substrate table, and a digital video imaging system. Free-falling distilled water drops released from a certain height were allowed to impact normally on the target surfaces. The image sequences of drop impact dynamics were constructed from the images captured using the digital video imaging system. Majority of the drop impact experiments were captured using a high speed video camera operating with frame speed ranging from 3000 to 10000 fps. For the target grooved surfaces, the impact dynamics was analyzed for the impacting drop liquid oriented both in the direction perpendicular to the grooves ( ) and in the direction parallel to the grooves (||) via independent test runs. The captured digital frames were used to deduce the temporal variation of impacting drop parameters such as drop contact diameter, drop contact angle, and drop height at the center of impacting drop with the aid of image processing software. The impacting drops were characterized in terms of Weber number, We expressed in terms of drop impact velocity and drop diameter measured just before the start of impact process. The study covered We ranging from 1.8 to 170. In general, the groove texture on the solid surface influences the drop impact process for all We examined in the study. The effect is more pronounced for the receding of impacting drops. For high We drops, the groove texture enhances the perturbations seen at the periphery of spreading lamella. The study showed quantitatively that the drop impact process on a target grooved surface comprising unidirectional parallel grooves develops a non-axisymmetric drop flow on the grooved surface exhibiting different spreading and receding processes of impacting drop liquid in the directions perpendicular ( ) and parallel (||) to the grooves. The maximum spreading diameter reached immediately after the completion of early inertia-dominated spreading in is less than that obtained in || due to the loss of drop kinetic energy caused by the pinned motion of drop liquid in . The non- axisymmetric drop flow on the target grooved surface develops a difference between the frequencies of contact angle oscillation of impacting drop liquid in  and ||. The frequency difference in contact angle oscillation causes the beating phenomenon in the temporal variation of the contact angle anisotropy, Δθ and drop height at the center of impacting drop, Z. For a given target grooved surface, the experimental measurements suggested that the beat frequency is almost independent of We. The temporal variation of Δθ and Z do not show the traces of beating phenomenon for the impact of high We drops. Owing to the non-axisymmetric drop flow, the final equilibrium drop shape is eccentric for the impact of low We drops and approaches a circular shape for the impact of high We drops. The role of groove texture geometrical parameters is seen in the drop impact process via the surface wettability especially for the impact of low We drops. Larger surface roughness factor makes the target grooved surface to exhibit hydrophobic characteristics.

Mieux comprendre les spécificités de la dynamique des couches limites à grand nombre de Reynolds malgré les contraintes métrologiques et son coût de simulation numérique est crucial. A titre d'exemple, cette dynamique peut déterminer plus de la moitié de la traînée d'un avion en croisière. Décrire la turbulence pariétale peut guider la résolution numérique d'une partie des fluctuations à un coût maîtrisé par des stratégies WMLES (simulation des grandes échelles avec modèle de paroi). Les présentes analyses physiques de couches limites turbulentes incompressibles à gradient de pression nul et à grand nombre de Reynolds s'appuient sur des simulations numériques avancées. Après validation d'une base de données, le frottement moyen pariétal est décomposé selon l'identité FIK (Fukagata et al. (2002)), dont l'application malgré le développement spatial est discutée. Une analyse spectrale montre que les grandes échelles (\lambda_x > \delta) contribuent à environ la moitié du frottement vers Re_\theta = 10^4. Les limitations de l'identité FIK motivent la dérivation d'une décomposition physique de la génération du frottement dont le comportement asymptotique est alors relié à la production d'énergie cinétique turbulente dans la zone logarithmique. Pour mieux reconstruire les spectres spatiaux, une nouvelle méthode d'estimation de la vitesse de convection turbulente en fonction de la longueur d'onde des fluctuations, adaptée au développement spatial et à des signaux temporels de durée finie, est dérivée, interprétée et évaluée à Re_\theta = 13000. Certaines des conclusions éclairent des modifications d'une stratégie WMLES, le mode III de la méthode ZDES. / Better understanding the specificities of the dynamics of high-Reynolds number boundary layers despite metrological constraints and its numerical simulation cost is crucial. For instance, this dynamics can determine more than half of the drag of a cruising aircraft. Describing wall turbulence can guide the numerical resolution of some of the fluctuations at a limited cost by WMLES strategies (wall-modelled large eddy simulation). The present physical analyses of zero-pressure gradient incompressible turbulent boundary layers at high Reynolds number rely on advanced numerical simulations. After validating a database, mean skin friction is decomposed by means of the FIK identity (Fukagata et al. (2002)), whose application despite the spatial growth is discussed. A spectral analysis shows that the large scales (\lambda_x > \delta) contribute approximately half of the friction near Re_\theta = 10^4. The limitations of the FIK identity motivate the derivation of a physical decomposition of the generation of friction whose asymptotic behaviour is then related to turbulent kinetic energy production in the logarithmic layer. In order to better reconstruct spatial spectra, a new method to estimate the turbulent convection velocity as a function of the wavelength of the fluctuations, adapted to spatial growth and to temporal signals of finite duration, is derived, interpreted, and assessed at Re_\theta = 13000. Some of the conclusions enlighten modifications to a WMLES strategy, mode III of the ZDES method.

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

A key component of research in the aerospace industry constitutes hypersonic flights (M>5) which includes the design of commercial high-speed aircrafts and development of rockets. Computational analysis becomes more important due to the difficulty in performing experiments and reliability of its results at these harsh operating conditions. There is an increasing demand from the industry for the accurate prediction of wall-shear and heat transfer with a low computational cost. Direct Numerical Simulations (DNS) create the standard for accuracy, but its practical usage is difficult and limited because of its high cost of computation. The usage of Reynold's Averaged Navier Stokes (RANS) simulations provide an affordable gateway for industry to capitalize its lower computational time for practical applications. However, the presence of existing RANS turbulence closure models and associated wall functions result in poor prediction of wall fluxes and inaccurate solutions in comparison with high fidelity DNS data. In recent years, machine learning emerged as a new approach for physical modeling. This thesis explores the potential of employing Machine Learning (ML) to improve the predictions of wall fluxes for hypersonic turbulent boundary layers. Fine-grid RANS simulations are used as training data to construct a suitable machine learning model to improve the solutions and predictions of wall quantities for coarser meshes. This strategy eliminates the usage of wall models and extends the range of applicability of grid sizes without a significant drop in accuracy of solutions. Random forest methodology coupled with a bagged aggregation algorithm helps in modeling a correction factor for the velocity gradient at the first grid points. The training data set for the ML model extracted from fine-grid RANS, includes neighbor cell information to address the memory effect of turbulence, and an optimal set of parameters to model the gradient correction factor. The successful demonstration of accurate predictions of wall-shear for coarse grids using this methodology, provides the confidence to build machine learning models to use DNS or high-fidelity modeling results as training data for reduced-order turbulence model development. This paves the way to integrate machine learning with RANS to produce accurate solutions with significantly lesser computational costs for hypersonic boundary layer problems.