No description available.

A recently-developed experimental technique based on the sublimation of naphthalene, which enables imaging of the dispersion of a passive scalar using planar laser-induced fluorescence (PLIF), is applied to a Mach 5 turbulent boundary layer and a NASA Orion capsule flowfield. To enable the quantification of naphthalene PLIF images, quantitative fluorescence and quenching measurements were made in a temperature- and pressure-regulated test cell. The test cell measurements were of the naphthalene fluorescence lifetime and integrated fluorescence signal over the temperature range of 100 K to 525 K and pressure range of 1 kPa to 40 kPa in air. These data enabled the calculation of naphthalene fluorescence yield and absorption cross section over the range of temperatures and pressures tested, which were then fit to simple functional forms for use in the calibration of the PLIF images. Quantitative naphthalene PLIF images in the Mach 5 boundary layer revealed large-scale naphthalene vapor structures that were regularly ejected out to wall distances of approximately y/δ = 0.6 for a field of view that spanned 3δ to 5δ downstream of the trailing edge of the naphthalene insert. The magnitude of the calculated naphthalene mole fraction in these structures at y/δ = 0.2 ranged from approximately 1-6% of the saturation mole fraction at the wind tunnel recovery temperature and static pressure. An uncertainty analysis showed that the uncertainty in the inferred naphthalene mole fraction measurements was ± 20%. Mean mole fraction profiles collected at different streamwise locations were normalized by the mole fraction measured at the wall and a characteristic height of the scalar boundary layer, causing the profiles to collapse into one “universal” mole fraction profile. Two-dimensional fields of naphthalene mole fraction were also obtained simultaneously with velocity by using particle image velocimetry (PIV) and PLIF. The images show large-scale naphthalene vapor structures that coincide with regions of relatively low streamwise velocity. The covariance of naphthalene mole fraction with velocity indicates that an ejection mechanism is transporting low-momentum, high-scalar-concentration fluid away from the wall, resulting in the protrusions of naphthalene vapor evident in the instantaneous PLIF images. Lastly, naphthalene PLIF was used to visualize the dispersion of gas-phase ablation products on a scaled Orion capsule model at four different angles of attack at Mach 5. High concentrations of scalar were imaged in the capsule recirculation region. Additionally, intermittent turbulent structures were visualized on the heat shield surface, particularly for the 12° and 52° AoA cases.

Cette thèse a été effectuée au sein de l’Institut de Physique de Rennes, et qui porte sur le transfert d’énergie rotationnelle lors des collisions de CO-Ar et CO-H2 à très basses températures pour des applications astrophysiques. Comprendre la constitution du milieu interstellaire (MIS), son évolution et ses propriétés physiques telles que la température et la densité, nécessite la connaissance de l’efficacité des collisions atomiques et moléculaires. Cette thèse expérimentale a été motivée par cet objectif. Le MIS et plus particulièrement les nuages moléculaires froids sont caractérisés par des températures très basses atteignant ~ -260 °C. Afin de reproduire ces conditions, on a utilisé la technique CRESU (Cinétique des Réactions en Ecoulement Supersonique Uniforme). Deux systèmes de collisions ont été étudiés : CO-Ar et CO-H2 pour leur impact dans les modèles astrophysiques (dits aussi modèles de transfert radiatifs). Une technique spectroscopique IR-VUV (Infrarouge-Ultraviolet dans le vide) en double résonance à base de lasers pulsés a été utilisée pour la détection et le diagnostic de l’efficacité des collisions et la détermination les constantes de collisions. Les résultats expérimentaux obtenus (pour la première fois) ont été comparés à des prédictions basées sur des calculs théoriques très avancés de mécanique quantique. Un très bon accord a été obtenu, ce qui a permis de tester et valider ces calculs théoriques d’un côté, et aussi de pouvoir fournir des constantes de collisions robustes qui vont être utilisées par les astrophysiciens pour modéliser et déterminer les propriétés physiques du MIS, ainsi qu’interpréter les spectres astrophysiques obtenus par des télescopes ou des satellites. / In the quest to understand the universe, astrophysicists observe and make models for astrophysical objects in the sky. The interstellar medium, ISM, in particular is of central importance since it represents the matter that exists in the space between stars in a galaxy, and in which stars and planets form. Understanding it, its constituents and its evolution and characteristics requires the quantification of several chemical and physical processes, including collision processes. In this work, we used the CRESU technique to reproduce very cold environments of astrophysical media, in particular dense molecular clouds in the ISM. We studied experimentally rotational energy transfer, RET, resulting from inelastic collisions at very low temperatures using a pump-probe laser-based spectroscopic technique for the purpose of measuring constants quantifying collisions. Two types of constants were determined: the first are total removal constants of RET resulting from a specific rotational quantum state to all possible final rotational quantum states, and the second are more detailed information consisting in rate constants from a specific rotational quantum state to another specific rotational quantum state, so-called state-to-state rate constants. Two experiments have been performed involving Carbone monoxide molecule, CO, as a target molecule of collisions. The first involves argon, Ar, as a projectile atom, and the second molecular hydrogen, H2, as a projectile molecule. Both collisional systems play an important role in a wide range of areas including gas-phase phenomena and astrophysical applications. In the first experiment, we investigated collisions between CO and Ar, from 293 K down to 30 K. IR-VUV double resonance technique has been exploited to measure, directly for the first time, absolute values of total removal and state-to-state constants of collisions. The experimental results have been compared to theoretical predictions based on a diatom-atom model of collision, where very good agreement was observed. In the second experiment, we investigated collisions between CO and H2 (the most abundant molecules in the ISM) from 293 K down to 5.5 K focusing on the very low temperatures of dense molecular clouds in the ISM. For the first time, total removal and state-to-state constants have been measured and compared to theoretical predictions of a highly accurate diatom-diatom model of collisions, where excellent agreement was observed. The results obtained in this thesis served to validate theoretical models of molecular collisions, helping the continuous efforts for pushing the frontiers of theoretical models. In the astrophysical side, the obtained collisional constants will be taken into account in modeling of many astrophysical media.

<div>The atomization of jets and droplets undergoing breakup in high-speed flows has been experimentally measured and theoretically modeled. Systems for producing individual droplet breakup and full jet breakup were designed, and a wide range of diagnostics were developed and adapted to measure the results with reduced uncertainty.</div><div><br></div><div>A detailed methodology for investigating high-speed sprays in the Purdue Experimental Turbine Aerothermal Lab is presented. Optical diagnostic techniques were carefully selected and optimized for the test section geometries and flow features, such that images could be collected at high frequencies of 20 kHz with high resolutions. Developed image processing routines are outlined to demonstrate how backlit imaging with specialized lenses allowed for more accurate spray depth measurements in supersonic conditions, which were then used in regression modeling routines to derive empirical correlations that factored in test section geometry, flow conditions, and injector design. A Mie scattering imaging technique was used for quantitative analysis of the supersonic spray plume profile and measurement of the spray width. 20 kHz shadowgraphy provided sufficient gradients for analysis of the unsteadiness of the spray and surrounding supersonic flow at the point of injection. Droplet sizes and velocities were measured in subsonic conditions using digital in-line holography, in which recent advancements to the reconstruction algorithm were implemented to reduce out-of-plane measurement uncertainty, and phase Doppler particle analysis.</div><div><br></div><div>The breakup of a single drop undergoing multi-mode breakup was analytically characterized, with the proposal of a new breakup criterion in the Taylor analogy breakup model. Hill vortices within the drop were proposed as a new flow mechanism promoting multi-mode breakup. Product drop sizes from the ring breakup were predicted and compared with experimental results.</div>

No description available.

Grid fins (lattice fins) are used as a lifting and control surface for highly manoeuvrable missiles. Grid fins also find their applications for air-launched submunitions. The main advantages are its low hinge moment requirement and good high angle of attack performance characteristics. Two dimensional analysis has been carried out using linear and shock-expansion theo- ries. The results indicate that above certain depth-to-height ratio, (called critical depth-to-height ratio,) the local normal force becomes negative due to shock reflection from the opposite side. Hence, depth (chord) for grid fin cell should not exceed a critical value. A prediction method has been developed for the estimation of aerodynamic character- istics of grid fin-body combinations at supersonic Mach numbers based on shock-expansion theory. Body upwash theory has been used for the effect of body; method of images has been used for carry-over forces onto the body. Empirical relation has been used for the modelling of separated body vortices and their effect on the leeward side fins. The method has been validated with experimental results for three configurations. The comparison is good for individual fin characteristics as well as overall characteristics for all the cases at higher supersonic Mach numbers. For lower supersonic Mach numbers at higher angles of attack, the prediction deviates from experiment. The reason for the deviation is due to shock detachment and shock reflection from opposite side, which is not modelled in the present method. Vortex lattice method has been used for prediction of linear aerodynamic character- istics of grid-fins at subsonic Mach numbers. Empirical relation based on trends from available experimental data has been used for the non-linear effect. The method has been validated with experimental results for several configurations without and with control surface deflections. The predicted aerodynamic characteristics compare well with experimental results for all the cases and the difference is within 15%. Based on the subsonic and supersonic analytical methods, a prediction code for the aerodynamic analysis of configuration with grid fins has been developed. Flow field computations inside isolated cells have been carried out using CFD code, PARAS-3D. Effects of depth-to-height ratio, web thickness, web leading edge angle and cell width-to-height ratio have been studied. Increase in thickness reduces the critical depth and increases the normal force. This increment in normal force is due to shock wave formation at the expansion side and its interaction with the opposite side. Effect of cell cross sectional shape has been studied using inviscid computation over isolated cells. Square, right triangular, equilateral triangular and hexagonal cross sections have been considered for this study. The normal force for square cell at zero roll is higher compared to 45 deg roll (diamond shape). Triangular cells show large variation in normal force with roll orientation due to large variation in projected area with roll angle. To compare the characteristics of different cross sectional cells, the normal force is normalised with respect to total internal web area. The comparison shows that the hexagonal cell gives maximum normal force and right triangular cell gives the minimum. Packaging efficiency of different cross sections is analysed by normalising the normal force with frontal area. The results show that triangular cells are preferred for packaging efficiency. Viscous flow computations over complete configuration have been carried out using FLUENT. GAMBIT has been used for geometry definition and grid generation. Hexahedral finite volumes are used to generate the grids including the nose region. Flow computations have been carried out at supersonic Mach numbers. To reduce the compu- tational time, Flow computations upto 0.5 calibre ahead of grid fin have been carried out with body-alone configuration. Flow over the fin-body section has been computed sep- arately taking the inlet pressure condition from the body-alone computed results. This procedure has reduced the grid size to around 1/5th and the computations converged faster due to imposition of converged solution at the pressure inlet. The computed results on the body show that the Flow separation occurs on the lee- ward side of the body and formation of separated vortices. The comparison of pressure distribution on the body with experiment is good. Flow computations over the fin-body section have been carried out at different Mach numbers and angles of attack. The computed normal force coefficient on the horizontal fin compares well with experimental data. Computations with fin deflection of -15 deg have also been carried out and the computed results are within 10% of the experimental data. Flow computations over another grid fin configuration have been carried out at dif- ferent roll angles. The comparison of individual fin force and overall normal force and pitching moment coefficients with experiment is good. The comparison demonstrates the capability of prediction methods as well as CFD in analysing aerodynamic performance of grid fin configurations.

Evaluating aerodynamic noise from aircraft engines is a design stage process, so that it conform to regulations at airports. Aerodynamic noise is also a principal source of structural vibration and internal noise in short/vertical take off and landing and rocket launches. Acoustic loads may be critical for the proper functioning of electronic and mechanical components. It is imperative to have tools with capability to predict noise generation from turbulent flows. Understanding the mechanism of noise generation is essential in identifying methods for noise reduction. Lighthill (1952) and Lighthill (1954) provided the first explanation for the mechanism of aerodynamic noise generation and a procedure to estimate the radiated sound field. Many such procedures, known as acoustic analogies are used for estimating the radiated sound field in terms of the turbulent fluid flow properties. In these methods, the governing equations of the fluid flow are rearranged into two parts, the acoustic sources and the propagation terms. The noise source terms and propagation terms are different in different approaches. A good description of the turbulent flow field and the noise sources is required to understand the mechanism of noise generation. Computational aeroacoustics (CAA) tools are used to calculate the radiated far field noise. The inputs to the CAA tools are results from CFD simulations which provide details of the turbulent flow field and noise sources. Reynolds-Averaged Navier Stokes (RANS) solutions can be used as inputs to CAA tools which require only time-averaged mean quantities. The output of such tools will also be mean quantities. While complete unsteady turbulent flow details can be obtained from Direct Numerical Simulation (DNS), the computation is limited to low or moderate Reynolds number flows. Large eddy simulations (LES) provide accurate description for the dynamics of a range of large scales. Most of the kinetic energy in a turbulent flow is accounted by the large-scale structures. It is also the large-scale structures which accounts for the maximum contribution towards the radiated sound field. The results from LES can be used as an input to a suitable CAA tool to calculate the sound field. Numerical prediction of turbulent flow field, the acoustic sources and the radiated sound field is at the focus of this study. LES based on explicit filtering method is used for the simulations. The method uses a low-pass compact filter to account for the sub-grid scale effects. A one-parameter fourth-order compact filter scheme from Lele (1992) is used for this purpose. LES has been carried out for four different flow situations: (i) round jet (ii) plane jet (iii) impinging round jet and (iv) impinging plane jet. LES has been used to calculate the unsteady flow evolution of these cases and the Lighthill’s acoustic sources. A compact difference scheme proposed by Hixon & Turkel (1998) which involves only bi-diagonal matrices are used for evaluating spatial derivatives. The scheme provides similar spectral resolution as standard tridiagonal compact schemes for the first spatial derivatives. The scheme is computationally less intensive as it involves only bi-diagonal matrices. Also, the scheme employs only a two-point stencil. To calculate the radiated sound field, the Helmholtz equation is solved using the Green’s function approach, in the form of the Kirchhoff-Helmholtz integral. The integral is performed over a surface which is present entirely in the linear region and covers the volume where acoustic sources are present. The time series data of pressure and the normal component of the pressure gradient on the surface are obtained from the CFD results. The Fourier transforms of the time series of pressure and pressure gradient are then calculated and are used as input for the Kirchhoff-Helmholtz integral. The flow evolution for free jets is characterised by the growth of the instability waves in the shear layer which then rolls up into large vortices. These large vortical structures then break down into smaller ones in a cascade which are convected downstream with the flow. The rms values of the Lighthill’s acoustic sources showed that the sources are located mainly at regions immediately downstream of jet break down. This corresponds to the large scale structures at break down. The radiated sound field from free jets contains two components of noise from the large scales and from the small scales. The large structures are the dominant source for the radiated sound field. The contribution from the large structures is directional, mainly at small angles to the downstream direction. To account for the difference in jet core length, the far field SPL are calculated at points suitably shifted based on the jet core length. The peak value for the radiated sound field occurs between 30°and 35°as reported in literature. Convection of acoustic sources causes the radiated sound field to be altered due to Doppler effect. Lighthills sources along the shear layer were examined in the form of (x, t) plots and phase velocity pattern in (ω, k) plots to analyse for their convective speeds. These revealed that there is no unique convective speeds for the acoustic sources. The median convective velocity Uc of the acoustic sources in the shear layer is proportional to the jet velocity Uj at the center of the nozzle as Uc ≈ 0.6Uj. Simulations of the round jet at Mach number 0.9 were used for validating the LES approach. Five different cases of the round jet were used to understand the effect of Reynolds number and inflow perturbation on the flow, acoustic sources and the radiated sound field. Simulations were carried out for an Euler and LES at Reynolds number 3600 and 88000 at two different inflow perturbations. The LES results for the mean flow field, turbulence profiles and SPL directivity were compared with DNS of Freund (2001) and experimental data available in literature. The LES results showed that an increase in inflow forcing and higher Reynolds number caused the jet core length to reduce. The turbulent energy spectra showed that the energy content in smaller scale is higher for higher Reynolds number. LES of plane jets were carried out for two different cases, one with a co-flow and one without co-flow. LES of plane jets were carried out to understand the effect of co-flow on the sound field. The plane jets were of Mach number 0.5 and Reynolds number of 3000 based on center-line velocity excess at the nozzle. This is similar to the DNS by Stanley et al. (2002). It was identified that the co-flow leads to a reduction in turbulence levels. This was also corroborated by the turbulent energy spectrum plots. The far field radiation for the case without co-flow is higher over all angles. The contribution from the low frequencies is directional, mainly towards the downstream direction. The range of dominant convective velocities of the acoustic sources were different along shear layers and center-line. The plane jet results were also used to bring out a qualitative comparison of flow and the radiation characteristics with round jets. For the round jet, the center-line velocity decays linearly with the stream-wise distance. In the plane jet case, it is the square of the center-line velocity excess which decays linearly with the stream-wise distance. The turbulence levels at any section scales with the center-line stream-wise velocity. The decay of turbulence level is slower for the plane jet and hence the acoustic sources are present for longer distance along the downstream direction. Subsonic impinging jets are composed of four regions, the jet core, the fully developed jet, the impingement zone and the wall jet. The presence of the second region (fully developed free jet) depends on the distance of the wall from the nozzle and the length of the jet core. In impinging jets, reflection from the wall and the wall jet are additional sources of noise compared to the free jets. The results are analysed for the contribution of the different regions of the flow towards the radiated sound field. LES simulations of impinging round jets and impinging plane jet were carried out for this purpose. In addition, the results have been compared with equivalent free jets. The directivity plots showed that the SPL levels are significantly higher for the impinging jets at all angles. For free jets, a typical time scale for the acoustic sources is the ratio of the nozzle size to the jet velocity. This is ro/Uj for round jets and h/Uj for plane jets. For impinging jets, the non-dimensionlised rms of Lighthill’s source indicates that the time scale for acoustic sources is the ratio of the height of the nozzle from the wall to the jet velocity be L/Uj. LES of impinging round jets was carried out for two cases with different inflow perturbations. The jets were at Reynolds number of 88000 and Mach number of 0.9, same as the free jet cases. The impingement wall was at a distance L = 24ro from the nozzle exit. For impinging round jets, the SPL levels are found to be higher than the equivalent free jets. From the SPL levels and radiated noise spectra it was shown that the contribution from the large scale structures and its reflection from the wall is directional and at small angles to the wall normal. The difference in the range of angles where the radiation from the large scale structures were observed shows the significance of refraction of sound waves inside the flow. The rms values of the Lighthill’s sources indicate two dominant regions for the sources, just downstream of jet breakdown and in the impingement zone. The LES of impinging plane jet was done for a jet of Mach number 0.5 and Reynolds number of 6000. The impingement wall was at a distance L = 10h from the nozzle exit. The radiated sound field appears to emanate from this impingement zone. The directivity and the spectrum plots of the far field SPL indicate that there is no preferred direction of radiation from the impingement zone. The Lighthill’s sources are concentrated mainly in the impingement zone. The rms values of the sources indicate that the peak values occur in the impingement zone. The results from the different flow situations demonstrates the capability of LES with explicit filtering method in predicting the turbulent flow and radiated noise field. The method is robust and has been successfully used for moderate Reynolds number and an Euler simulation. An important feature is that LES can be used to identify acoustic sources and its convective speeds. It has been shown that the Lighthill source calculations, the calculated sound field and the observed radiation patterns agree well. An explanation for these based on the different turbulent flow structures has also been provided.

Lattice Boltzmann Method has been quite successful for incompressible flows. Its extension for compressible (especially supersonic and hypersonic) flows has attracted lot of attention in recent time. There have been some successful attempts but nearly all of them have either resulted in complex or expensive equilibrium function distributions or in extra energy levels. Thus, an efficient Lattice Boltzmann Method for compressible fluid flows is still a research idea worth pursuing for. In this thesis, a new Lattice Boltzmann Method has been developed for compressible flows, by using the concept of a relaxation system, which is traditionally used as semilinear alternative for non-linear hypebolic systems in CFD. In the relaxation system originally introduced by Jin and Xin (1995), the non-linear flux in a hyperbolic conservation law is replaced by a new variable, together with a relaxation equation for this new variable augmented by a relaxation term in which it relaxes to the original nonlinear flux, in the limit of a vanishing relaxation parameter. The advantage is that instead of one non-linear hyperbolic equation, two linear hyperbolic equations need to be solved, together with a non-linear relaxation term. Based on the interpretation of Natalini (1998) of a relaxation system as a discrete velocity Boltzmann equation, with a new isotropic relaxation system as the basic building block, a Lattice Boltzmann Method is introduced for solving the equations of inviscid compressible flows. Since the associated equilibrium distribution functions of the relaxation system are not based on a low Mach number expansion, this method is not restricted to the incompressible limit. Free slip boundary condition is introduced with this new relaxation system based Lattice Boltzmann method framework. The same scheme is then extended for curved boundaries using the ghost cell method. This new Lattice Boltzmann Relaxation Scheme is successfully tested on various bench-mark test cases for solving the equations of compressible flows such as shock tube problem in 1-D and in 2-D the test cases involving supersonic flow over a forward-facing step, supersonic oblique shock reflection from a flat plate, supersonic and hypersonic flows past half-cylinder.