<p>Rising fuel costs and growing environmental concerns have forced gas turbine engine manufacturers to place high value on reducing fuel burn. This trend has pushed compressor technology into new design spaces that are not represented by historical experience. Specifically, centrifugal compressor diffusers are trending toward higher pressure recovery and smaller diameters. The internal fluid dynamics in these new flow regimes are not well understood and additional study is necessary. This work outlines detailed experimental and numerical observations of the flow field through a vaned diffuser for aeroengine applications.</p> <p>The experimental data consist of extensive Laser Doppler velocimetry measurements of the unsteady velocity field from the impeller trailing edge through the majority of the diffuser passage. These data were obtained non-intrusively and yielded all three components of the velocity vector field at approximately 2,000 geometric points. The correlation between fluctuations in the three velocity components were also observed at several key locations to determine the components of the local Reynolds stress tensor.</p> <p>These data indicated a jet/wake profile at the impeller exit represented by a consistent velocity deficit region from hub to shroud adjacent to the suction surface of the passage. This region was more prevalent adjacent to the splitter blade. The unsteady fluctuations due to the propagation of the jet and wake through the diffuser passage persist to 40% downstream of the throat. A complex secondary flow field was also observed with large axial velocities and a passage-spanning vortex developing through the diffuser passage. The velocity data and total-pressure data indicated a region of flow separation developing along the pressure surface of the vane near the hub due to the unsteady propagation of the jet and wake flow through the diffuser. Although this region was stable in time, its development arose due to unsteady aspects of the flow. Finally, the strong interconnection between the jet and wake flow, unsteady fluctuations, secondary velocities, incidence, and flow separation was demonstrated. </p> <p>Computationally, a “best-practice” methodology for the modelling of a centrifugal compressor was developed by a systematic analysis of various turbulence models and many modelling features. The SST and BSL-EARSM turbulence models with the inclusion of fillets, surface roughness, and non-adiabatic walls was determined to yield the best representation of the detailed flow development through the diffuser in steady (mixing-plane) simulations. The accurate modelling of fillets was determined to significantly impact the prediction of flow separation with the SST turbulence closure model. Additionally, the frozen rotor approach was shown to not accurately approximate the influence of unsteady effects on the flow development.</p> <p>Unsteady simulations were also compared to the detailed experimental data through the diffuser. The BSL-EARSM turbulence model best matched the experimentally observed flow field due to the SST model’s prediction of flow separation in the shroud-pressure side corner of the passage. In general, lower levels of axial velocity were predicted numerically that resulted in less spanwise mixing between the endwall and freestream flows. Additionally, the turbulent kinetic energy levels in the computational results showed little streamwise variation through the vaneless and semi-vaneless space. The large variation observed experimentally indicated that the production and dissipation of turbulent kinetic energy through this region was not accurately predicted in the two turbulence models implemented for the unsteady simulations.</p>

This study investigates the effects of near wall disturbances in the form of roughness on a compressible turbulent boundary layer. The studies were carried out using numerical methods which directly solve the Navier-Stokes equations. This provides for unique opportunities to investigate three dimensional structures of the flow as well as avoid the loss of physical fidelity with turbulence modeling. Three cases were ran, a smooth wall case, and two rough wall cases with different heights of the roughness elements between the cases. The results are first visualized with different approaches. Then statistical methods were used to characterize the flow.

<div>High speed flow problems of practical interest require a solution of nonequilibrium aerothermochemistry to accurately predict important flow phenomena including surface heat transfer and stresses. As a majority of these flow problems are in the continuum regime, Computational Fluid Dynamics (CFD) is a useful tool for flow modeling. This work presents the development of a nonequilibrium add-on solver to ANSYS Fluent utilizing user-defined-functions to model salient aspects of nonequilibrium flow in air. The developed solver was verified for several benchmark nonequilibrium flow problems and compared with the available experimental data and other nonequilibrium flow simulations. <br></div><div><br></div><div>The rate of dissociation behind a strong shock in thermochemical nonequilibrium depends on the vibrational excitation of molecules. The Macheret-Fridman (MF) classical impulsive model provides analytical expressions for nonequilibrium dissociation rates. The original form of the model was limited to the dissociation of homonuclear molecules. In this work, a general form of the MF model has been derived and present macroscopic rates applicable for modeling dissociation in CFD. Additionally, some improvements to the prediction of mean energy removed in dissociation in the MF-CFD model has been proposed based on the comparisons with available QCT data. In general, the results from the MF-CFD model upon investigating numerous nonequilibrium flows are promising and the model shows a possibility of becoming the standard tool for investigating nonequilibrium flows in CFD.</div><div><br></div><div>The aerodynamic deorbit experiment (ADE) CubeSat has dragsail to accompany accelerated deorbiting of a CubeSat post-mission. A good estimation of the aerothermal load on a reentry CubeSat is paramount to ensure a predictable reentry. This study investigates the aerothermal load on an ADE CubeSat using the direct simulation Monte Carlo (DSMC) methods and Navier-Stokes-Fourier continuum based methods with slip boundary conditions. The aerothermal load on an ADE CubeSat at 90 km altitude from the DSMC and continuum methods were consistent with each other. The continuum breakdown at a higher altitude of 95 km resulted in a strong disagreement between the continuum and DSMC solutions. Overall, the continuum methods could offer a considerable computational cost saving to the DSMC methods in predicting aerothermal load on an ADE CubeSat at low altitudes.<br> </div>

<div>The development work of a race car revolves around numerous goals such as drag reduction,</div><div>maximizing downforce and side force, and maintaining balance. Commonly, these goals</div><div>are to be met at the same time thus increasing the level of difficulty to achieve them. The</div><div>methods for data acquisitions available to a race team during the season is mostly limited to</div><div>wind tunnel testing and computational fluid dynamics, both of which are being heavily regulated</div><div>by sanctioning bodies. While these methods enable data collection on a regular basis</div><div>with repeat-ability they are still only a simulation, and as such they come with some margin</div><div>of error due to a number of factors. A significant factor for correlation error is the effect of</div><div>tires on the flow field around the vehicle. This error is a product of a number of deficiencies</div><div>in the simulations such as inability to capture loaded radius, contact patch deformation in</div><div>Y direction, sidewall deformation and overall shifts in tire dimensions. These deficiencies</div><div>are evident in most WT testing yet can be captured in CFD. It is unknown just how much</div><div>they do affect the aerodynamics performance of the car. That aside, it is very difficult to</div><div>correlate those findings as most correlation work is done at WT which has been said to be</div><div>insufficient with regards to tire effect modeling. Some work had been published on the effect</div><div>of tire deformation on race car aerodynamics, showing a large contribution to performance</div><div>as the wake from the front tires moves downstream to interact with body components. Yet</div><div>the work done so far focuses mostly on open wheel race cars where the tire and wheel assembly</div><div>is completely exposed in all directions, suggesting a large effect on aerodynamics.</div><div>This study bridges the gap between understanding the effects of tire deformation on race car</div><div>aerodynamics on open wheel race cars and closed wheel race cars. The vehicle in question</div><div>is a hybrid of the two, exhibiting flow features that are common to closed wheel race cars</div><div>due to each tire being fully enclosed from front and top. At the same time the vehicle is</div><div>presenting the downstream wake effect similar to the one in open wheel race cars as the</div><div>rear of the wheelhouse is open. This is done by introducing a deformable tire model using</div><div>FEA commercial code. A methodology for quick and accurate model generation is presented</div><div>to properly represent true tire dimensions, contact patch size and shape, and deformed dimension,</div><div>all while maintaining design flexibility as the model allows for different inflation</div><div>pressures to be simulated. A file system is offered to produce CFD watertight STL files that</div><div>can easily be imported to a CFD analysis, while the analysis itself presents the forces and</div><div>flow structures effected by incorporating tire deformation to the model. An inflation pressure</div><div>sweep is added to the study in order to evaluate the influence of tire stiffness on deformation</div><div>and how this results in aerodynamic gain or loss. A comparison between wind tunnel</div><div>correlation domain to a curved domain is done to describe the sensitivity each domain has</div><div>with regards to tire deformation, as each of them provides a different approach to simulating</div><div>a cornering condition. The Study suggests introducing tire deformation has a substantial</div><div>effect on the flow field increasing both drag and downforce.In addition, flow patterns are</div><div>revealed that can be capitalized by designing for specific cornering condition tire geometry.</div><div>A deformed tire model offers more stable results under curved and yawed flow. Moreover,</div><div>the curved domain presents a completely different side force value for both deformed and</div><div>rigid tires with some downforce distribution sensitivity due to inflation pressure.</div>

<div>Initial experiments on separated shock/boundary-layer interactions were carried out within the Boeing/AFOSR Mach-6 Quiet Tunnel. Measurements were made of hypersonic laminar-turbulent transition within the separation above a compression corner. This wind tunnel features freestream fluctuations that are similar to those in</div><div>flight. The present work focuses on the role of traveling instabilities within the shear layer above the separation bubble.</div><div>A 7 degree half-angle cone with a slice and a finite-span compression ramp was designed and tested. Due to a lack of space for post-reattachment sensors, early designs of this</div><div>generic geometry did not allow for measurement of a post-reattachment boundary layer. Oil flow and heat transfer measurements showed that by lengthening the ramp, the post-reattachment boundary layer could be measured. A parametric study was completed to determine that a 20 degree ramp angle caused reattachment at 45% of the</div><div>total ramp length and provided the best flow field for boundary-layer transition measurements.</div><div>Surface pressure fluctuation measurements showed post-reattachment wave packets and turbulent spots. The presence of wave packets suggests that a shear-layer</div><div>instability might be present. Pressure fluctuation magnitudes showed a consistent transition Reynolds numbers of 900000, based on freestream conditions and distance</div><div>from the nosetip. Pressure fluctuations grew exponentially from less than 1% to roughly 10% of tangent-wedge surface pressure during transition.</div><div>A high-voltage pulsed plasma perturber was used to introduce controlled disturbances into the boundary layer. The concept was demonstrated on a straight 7 degree half-angle circular cone. The perturbations successfully excited the second-mode instability at naturally unstable frequencies. The maximum second-mode amplitudes prior to transition were measured to be about 10% of the mean surface static pressure. </div><div>The plasma perturber was then used to disturb the boundary layer just upstream of the separation bubble on the cone with the slice and ramp. A traveling instability was measured post-reattachment but the transition location did not change for any tested condition. It appears that the excited shear-layer instability was not the dominant mechanism of transition.</div>

<div>Experiments were performed on a sharp-nosed 7° half-angle cone at a 6° angle of attack in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) to study the stationary crossflow instability and its receptivity to small surface roughness. Heat transfer measurements were obtained using temperature sensitive paint (TSP) and Schmidt Boelter (SB) heat transfer gauges. Great care was taken to obtain repeatable, quantitative measurements from TSP.</div><div></div><div>Consecutive runs were performed at a 0° angle of attack, and the heat transfer measured by the SB was found to drop as the initial model temperature increased, while other initial conditions such as stagnation pressure were held constant. This agreed with calculations done using a similarity solution. It was found that repeatable measurements at a 6° angle of attack could be made if the initial model temperature was controlled and the patch location that was used to calibrate the TSP was picked in a reasonable and consistent manner.</div><div></div><div>The Rod Insertion Method (RIM) roughness, which was used to excite the stationary crossflow instability, was found to be responsible for the appearance of the streaks that were analyzed. The signal-to-noise ratio in the TSP was too low to properly measure the streaks directly downstream of the roughness insert. The heat transfer along the streak experienced linear growth, peaked, and then slightly decayed. It is possible this peak was saturation. The general trend was that the growth of the streaks moved farther upstream as the roughness element height increased, which agreed with past computations and low speed experiments. The growth of the streak also moved farther upstream as the freestream Reynolds number increased. The amplitude of the streaks was calculated by non-dimensionalizing the heat transfer using the laminar theoretical mean-flow solution for a 7° half-angle cone at a 6° angle of attack. The relationship between the amplitude and the non-dimensional roughness height was approximately linear in the growth region of the streaks.</div>

The objective of this research is to investigate the effects of tip clearance on the stage and component performance in a high-speed centrifugal compressor. The experimental data were compared against results from a numerical model to assess the ability of the numerical simulation to predict the effects of tip clearance. Experimental data were collected at Purdue University on the Single Stage Centrifugal Compressor (SSCC), a high-speed, high-pressure ratio test compressor sponsored by Honeywell Aerospace. Numerical simulations were completed using the ANSYS CFX software suite and part of the research computing clusters located at Purdue University.<div><br></div><div>Two tip clearances were tested, the nominal tip clearance and a tip clearance that is 66% larger than the nominal clearance, at speeds from 60% to 100% corrected speed. To compare data points with different tip clearances, various parameters were evaluated, and one was chosen. The value of TPR/inlet corrected mass flow rate best represented similar loading conditions, and thus similar incidences, for each tip clearance and was chosen as the best method for comparing similar data points taken with different clearances. Stage and component performance were focused on the sensitivity of each performance parameter to the changing of the tip clearance. The stage total pressure ratio and stage efficiency showed moderate sensitivity while the stage work factor showed much lower sensitivity. The impeller is more sensitive to changing tip clearances than the stage is, showing greater changes when comparing data from each tip clearance. The diffuser was on the same order of sensitivity as the impeller, with marginally higher sensitivities for some parameters. It was found that by the typical performance metrics, the diffuser performs worse at the nominal clearance than at the larger clearance. Upon further investigation though, the impeller is providing a higher static pressure and therefore, more diffusion, at the nominal clearance so the diffuser must perform less diffusion during nominal clearance operation.<br></div><div><br></div><div>To assess the validity of a prediction of the performance and sensitivity of the stage and components to the tip clearance, a numerical model was developed and validated. The numerical model was able to reasonably predict the stage performance with better comparisons of performance in the impeller and worse in the diffuser. The instrumentation in the experiment was replicated in the software to calculate performance the same way it is calculated experimentally so that the results would be comparable. While the performance of the stage and components was lacking in some areas, the trends predicted were similar to those calculated from the experimental data. As with the performance, the trends in the impeller matched very well between the experiment and the numerical simulation. The trends in stage and diffuser performance were predicted more accurately than the stage and diffuser performance maps and were able to capture the magnitude of the change in performance caused by changing the tip clearance. <br></div>

<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>

<p>A study has been undertaken to investigate the flow structure in the vicinity of discontinuities in the surface of a high-speed air vehicle. The effect of gaps and steps on aerodynamic heating is of particular interest. The present    thesis presents Reynolds-averaged Navier Stokes  (RANS) calculations of this class of flow. This thesis consists of two studies: a parametric study of cavity flow at Mach 2 and a study to compare with wind tunnel experiments at Mach 6. The calculations for the parametric study used the Menter two-equation SST turbulence model at fully turbulent conditions. These are two-dimensional cavity flows that were carried out to identify the influence of cavity geometry on flow structure and heating distribution inside the cavity, and to categorize cavity flow regimes. The second study employed RANS calculations for conditions  corresponding to Mach 10 wind tunnel experiments carried out by Nestler et al. (AIAA Paper 1968-673) for Mach 6 boundary layer edge conditions. The SST model used in the parametric study was paired with the Menter oneequation transition model and the two-equation realizable κ-ϵ model in CFD++ was used for the computations. The results showed that, even with adjustment of model parameters, the Menter transition model cannot match the location of laminar to turbulent transition, but it demonstrated good agreement with the experimental data in fully turbulent conditions. The two-equation realizable κ-ϵ model, available in CFD++, was able to accurately model transition and showed favorable agreement for fully turbulent conditions as well.</p>

<p>Laminar or turbulent flows over porous surfaces have received extensive attention in the past few decades, due to their potential to achieve passive flow controls. These surfaces either in natural exhibit roughness or are engineered in purpose, and usually entail special features such as increasing/reducing surface drags. An increasing interest has arisen in the interaction between these surfaces and high-speed compressible flows, which could inform the next-level flow control studies at supersonic and hypersonic speeds for the designs of high-speed vehicles. In this dissertation, the interaction between high-speed compressible turbulent flows and acoustically permeable surface is investigated. The surface property is modeled via the Time-Domain Impedance Boundary Condition (TDIBC), which avoids the inclusion of the geometric details in the numerical simulations.</p> <p>We first perform Large-Eddy Simulations of compressible turbulent channel flows over one impedance wall for three bulk Mach numbers:Mb = 1.5, 3.5 and 6.0. The bulk Reynolds number Reb is tuned to achieve similar viscous Reynolds number Re∗τ ≈ 220 across all Mb to ensure a nearly common state of near-wall turbulence structures over impermeable walls. The TDIBC based on the auxiliary differential equations (ADE) method is applied to bottom wall of the channel. A three-parameter complex impedance model with a resonating frequency tuned to the large-eddy turn-over frequency of the flow is adopted. With a sufficiently high permeability, a streamwise traveling instability wave that is confined in nature and that increases the surface drag, is observed in the near-wall region and changes the local turbulent events. As a result, the first and second order mean flow statistics are found to deviate from that of a flow over impermeable walls. We then perform a linear stability analysis using a turbulent background base flow and confirm that the instability wave is triggered by a sufficiently high permeability and manifests a confined nature. The critical resistance Rcr (interpreted as the inverse of the permeability), above which the instability is suppressed, is found to be sub-linearly proportional to the bulk Mach number Mb, indicating less permeability required to trigger the instability in high Mach number flows.</p> <p>Due to the extremely high computational cost in high Mach number wall-bounded flow calculations, the next-phase optimization/flow control design using the porous surface becomes unaffordable. An ’economical’ flow setup that can server the purpose of rapid flow generation would greatly benefit the planned research. For such reason, we carry out a study about the effect of the domain size on the near-wall turbulence structures in compressible turbulent channel flows, to identify such type of flow setup. Apart from the concept of minimal flow units (MFU, as in the literature) entailing a minimal domain size required for near-wall turbulence to be sustained, efforts have also been made to identify a range of the domain size that can sustain both the inner and outer layer turbulence, and lead to only small deviations in mean flow statistics from the baseline data, which herein defined as minimal turbulent channel (MTC). The motivation of proposing the concept of MTC is to provide a computationally efficient setup for the rapid generation of near-wall turbulence with minimal compromise on the fidelity of the simulated field for investigations requiring numerous simulations, such as machine learning, flow control/optimization designs. It is found that the mean flow statistics from a computational domain spanning 700 − 1100 and 230 − 280 local viscous units in streamwise and spanwise directions, respectively, agree reasonably well with the reference calculations of all three Mach numbers under investigation, and are thus identified as the range in which the MTC stays. The large scale near-wall turbulence structures observed in full scale DNS simulations, and their spatially coherent connections, are roughly preserved in MTC, indicated by the existence of the grouped streamwise aligned hairpin vortices of various sizes and the resulted patterns of uniform momentum zones and thermal zones in the instantaneous flow field. In an MTC, the energy transfer paths among the kinetic energy of the mean field, turbulent kinetic energy and mean internal energy are slightly modified, with the most significant change observed in the viscous dissipation. The mean wall-shear stress and mean wall heat flux see less than 5% error as compared to the full scale simulations. Such reduced-order flow setup requires less than 3% of the computational resource as compared to the full scale simulations.</p>