The initial slow viscous growth of the Tollmein-Schlichting wave in a canonical boundary layer transition is absent in bypass and wake-induced transitions. Although there have been a great deal of studies pertaining to bypass transition in boundary layers, the underlying breakdown mechanism is not clearly understood and it continues to be a subject of interest. Similarly, a wake-induced transition caused by Karman wake in the freestream remains poorly understood. The breakdown in this case is caused by anisotropic disturbances containing large scale unsteadiness in the freestream. Differing view points among workers on the transition process have also added to the complexities. In this thesis, bypass and wake-induced boundary layer transitions studied experimentally towards understanding various flow breakdown features are reported. The measurements were made on a flat plate boundary layer in a low-speed wind tunnel. The particle image velocimetry (PIV) technique was extensively used. Various grids were used to generate nearly isotropic freestream turbulence. A circular cylinder was placed at different heights from the plate leading edge to generate Karman wake in the freestream. Two cylinders of different diameters were used to vary the Reynolds number(based on the cylinder diameter). The PIV measurements being simultaneous over a large spatial domain enabled to assess various spatial transitional flow structures. In the case of bypass transition, the streamwise velocity fluctuation, u, is found to exhibit some organized negative and positive fluctuations that dominate the flow during transition, and confirm the simulation results reported in the literature. These positive and negative u fluctuations are found to be associated with the streak unsteadiness. By conditional sampling of these positive and negative u fluctuations, we find that urms (root-mean-squaredof u)can be expressed as a linear combination of urms,f and urms,b,i.e. urms = a(urms,f + urms,b); ais constant, and the subscripts fand bdenote the positive and nega-tive ufluctuations, respectively. Both urms,f and urms,b arefoundto follow the non-modal growth distribution. The wall-normal results clearly show that an inclined shear layer is often associated with an organized structure of negative ufluctuations and an inflectional in-stantaneous velocity profile. These inclined shear layers appear to be similar to those in ribbon-induced transition. The turbulent spot precursor appears to be the vortex shedding from an oscillating in-clined shear layer. Interestingly, the normalized vortex shedding fre-quency is found to be Reynolds number invariant, as in the case of ribbon-induced transition. The present study also confirms the sim-ulated turbulent spot features, including a thin log-law at the break-down stage. The spanwise plane PIV results reveal the signature of streak secondary instability in the flow in terms of symmetric and anti-symmetric streaks oscillations. The initial growth of streak amplitude is followed by a slow decay. The maximum streak amplitude is well above30% of the freestream velocity. These two aspects provide support to the streak instability analysis reported in the literature. While the present wake-induced transition study provides some sup-port to the available numerical simulation and experimental results, some new results have also emerged. The measured sharp rise in the disturbance energy during transition is found to be closer to the simulated result, compared to the difference reported in the literature. The spanwise vortices in the early stage, as also seen in other experimental studies, deform leading to the formation of lambda structures, the signature of which is found by the linear stochastic analysis. With increased Reynolds number and decreased cylinder height from the plate, the physical size of the lambda structure is found to decrease. These lambda structures are often found to appear in a staggered manner in the spanwise plane, as in the case of sub-harmonic boundary layer transition. Although a sub-harmonic peak in the frequency spectra is reported in the literature, as also in the present study, the clear staggered pattern went unnoticed. Streamwise streaks are subsequently generated due to the mean shear stretching of these lambda vortices. The spanwise spacing of these streamwise streaks is found to be comparable with the recent simulation results. Also, these streaks are found to undergo somewhat sinuous-like oscillations, compared to the only varicose type oscillations reported in the literature. The streak amplitude is found to saturate at about 35% of the freestream speed. Here again an inclined shear layer in the wall-normal plane is associated with organized negative u fluctuations and an inflectional instantaneous velocity profile. The movement of the peak urms towards the wall is found to be due to the positive u fluctuation, which follows a hairpin-like structure. The inclined shear layers herein are associated with the lambda or a hairpin-like structure. As in a by-pass transition, an inclined shear layer, vortex shedding from it, the imprint of which is also found in the linear stochastic analysis are present. The normalized high frequency shed vortices is found to be Reynolds number invariant in the present wake-induced transition, as in ribbon-induced and bypass transitions. Compared to the re-cent suggestion that the parent-offspring mechanism is the governing self-sustaining mechanism in the boundary layer, the present study suggests that streak-instability mechanism is also present. The proper orthogonal decomposition(POD) analysis of the experimental data was carried out with an emphasis on the bypass transition case studied. The first few energetic POD modes are found to capture the dominant flow structures, i.e. the organized positive and negative u fluctuations. In the case of bypass transition, the first two energetic POD modes are self-similar, i.e. independent of the freestream turbulent intensity and the Reynolds number. An attempt is also made to construct a low-dimensional model with the POD eigenfunction modes to predict the qualitative dynamics of bypass transition. This has revealed the existence of a traveling disturbance in the bypass transition. On the whole, the present study shows some similar breakdown features in bypass and wake-induced transitions, although more studies in this regard are essential.

Laminar flow control holds major promise for increasing aircraft efficiency and increasing laminar flow over aerodynamic surfaces could decrease drag by up to 30 percent. The Flight Research Lab at Texas A&M University has studied laminar flow over a wing with 30 degrees of leading edge sweep with Discrete Roughness Elements (DREs) installed and has indicated that DREs can be used to increase laminar flow at Reynolds numbers up to 7.5 million at Mach 0.3. A new project, termed SARGE, has been commissioned in conjunction with NASA for studying DREs on a swept wing glove at conditions relevant to jet transports. The SARGE project must have an instrumentation system capable of accurately measuring flow conditions and transition location on the suction side of the glove. Infrared (IR) thermography has been selected as the primary transition detection tool. A heat transfer analysis has shown that solar radiation will warm the surface of the glove above the adiabatic wall temperature and therefore the laminar region will appear to be warmer. The FLIR SC8000 IR camera has been selected for this application due to its ability to produce high-resolution images in the appropriate IR band. High quality air data is also required for the experiment. A five-hole probe will be used to measure flow angle and velocity near the glove. This instrument will provide meanflow conditions due to its limited frequency response. High quality pressure transducers coupled with careful probe calibration will allow for differential measurements to be made with an uncertainty of +/- 0.03 degrees. Static pressure ports and high frequency response Kulite transducers will also be employed. Hotfilm sensors will be used to verify the state of the boundary layer on the glove through spectral analysis. A unique hotfilm array has been proposed that will enable the measurement of traveling wave vectors through a spectral technique. An experiment on the Flight Research Lab's Cessna O-2 to investigate the veracity of this technique has also been suggested. Thermocouples will also be installed on the glove's surface to monitor temperatures and verify transition location. The layout of the hotfilms and thermocouples is also detailed.

A key remaining challenge in the design of hypersonic vehicles is the incomplete understanding of the process of boundary-layer transition. Turbulent heating rates are substantially higher than those for a laminar boundary layer, and large uncertainties in transition prediction therefore demand conservative, inefficient designs for thermal protection systems. It is only through close collaboration between theory, experiment, and computation that the state of the art can be advanced, but experiments relevant to flight require ground-test facilities with very low disturbance levels. To enable this work, a unique Mach 6 low-disturbance wind tunnel, previously of NASA Langley Research Center, is established within a new pressure-vacuum blow-down infrastructure at Texas A&M. A 40-second run time at constant conditions enables detailed measurements for comparison with computation. The freestream environment is extensively characterized, with a large region of low-disturbance flow found to be reliably present for unit Reynolds numbers Re < 11×10^6 m-1. Experiments are performed on a 5º half-angle flared cone model at Re = 10×10^6 m-1 and zero angle of attack. For the study of the second-mode instability, well-resolved boundary-layer profiles of mean and fluctuating mass flux are acquired at several axial locations using hot-wire probes with a bandwidth of 330 kHz. The second mode instability is observed to undergo significant growth between 250 and 310 kHz. Mode shapes of the disturbance agree well with those predicted from linear parabolized stability equation (LPSE) computations. A 17% (40 kHz) disagreement is observed in the frequency for most-amplified growth between experiment and LPSE. Possible sources of the disagreement are discussed, and the effect of small misalignments of the model is quantified experimentally. A focused schlieren deflectometer with high bandwidth (1 MHz) and high signal-to-noise ratio is employed to complement the hot-wire work. The second-mode fundamental at 250 kHz is observed, as well as additional harmonic content not discernible in the hot-wire measurements at two and three times the fundamental. A bispectral analysis shows that after sufficient amplification of the second mode, several nonlinear mechanisms become significant, including ones involving the third harmonic, which have not hitherto been reported in the literature.

A high level of free-stream turbulence and surface roughness are known to cause breakdown of an otherwise stable laminar flow. In transition induced by free-stream turbulence, streaks are formed due to the lift-up effect and low-speed streaks with high shear breakdown to turbulence. Streaks are also present in transition caused by a roughness element and they may breakdown via sinuous or varicose instability. In general, streamwise streaks, their lift-up and streak instability are integral to the bypass transition process. If the lift-up of a high-shear layer or its breakdown is manipulated by some external means, then the downstream flow is expected to change. An experimental study was carried out to understand the effect of flow modification caused by a mesh placed normal to the flow and at different wall-normal locations in the late stage of bypass transitions induced separately by an isolated cylindrical roughness element and a high level of free-stream turbulence. The measurements were made on a flat plate boundary layer in a low-speed wind tunnel using the particle image velocimetry technique. The mesh causes an approximately 30% reduction in the free-stream velocity, and mild acceleration in the boundary layer, irrespective of its wall-normal location. Interestingly, when located near the wall, the mesh suppresses several transitional events leading to transition delay over a large downstream distance. The transition delay is found to be mainly caused by suppression of the lift-up of the high-shear layer and its distortion, along with modification of the spanwise streaky structure to an orderly one. However, with the mesh well away from the wall, the lifted-up shear layer remains largely unaffected, and the downstream boundary layer velocity profile develops an overshoot which is found to follow a plane mixing layer type profile up to the free stream. Reynolds stresses, and the size and strength of vortices increase in this mixing layer region. The high-intensity disturbance in this region can possibly enhance the transition of accelerated flow far downstream, although a reduction in streamwise turbulence intensity occurs over a short distance downstream of the mesh. However, the shape of large-scale streamwise structure in the wall-normal plane is found to be more or less the same as that without the mesh.

<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>

<p> Hypersonic boundary-layer transition greatly affects aerodynamic heating, skin friction, aircraft stability and other characteristics on flight vehicles. Understanding the factors leading to laminar-turbulent transition is pivotal in hypersonic aircraft design. Various instabilities and modes may facilitate transition at hypersonic speeds including first and second-mode waves, Görtler vortices, and cross-flow which may be stationary or traveling. The research presented here will focus on investigating traveling cross-flow instabilities on a 7° half-angle cone at 6° angle of attack. The experiments were conducted in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) at Purdue University. The low freestream noise of the quiet tunnel facility made it ideal for studying boundary layer transition due to its more, ”flight like” environment when compared to traditional tunnel environments. Previous experiments by Ryan Henderson, Chris Ward, and Joshua Edelman focused on studying the cross-flow instability on right circular cones at angle of attack (AoA) in the BAM6QT. From these experiments it was decided that a means for taking off-surface pressure measurements on a cone was needed. This work sets out to create a micro-pitot traverse system capable of doing such. The system is able to measure pressure fluctuations within the boundary layer of cone models at precise axial, azimuthal and wall-normal locations. The design for the traverse was based off a traverse used at Notre Dame which was designed by David Cavalieri in his PhD dissertation for Illinois Institute of Technology. Micro-pitot probes created using hypodermic tubing and Kulite sensors were created to attach to the end of the traverse and take pressure measurements. The micro-pitot probes were placed such that they formed two distinct spatial pairs capable of measuring both the phase speed and propagation angle of traveling cross-flow instabilities using the difference in time of arrival of the traveling instability between the sensor pairs. The micro-pitot probes developed were made from telescoped hypodermic tubes housing Kulite XCE-061-15A sensors. The telescoped tubing assembly caused attenuation at higher frequencies affecting the micro-pitot probes ability to measure pressure fluctuations at higher frequencies. It was necessary to increase the dynamic performance of the micro-pitot probes in order to capture the cross-flow instability. To accomplish this a custom built frequency 17 compensator was designed to correct for this attenuation. The process for designing the compensator utilized a Mach 4 supersonic jet system (SSJ) to estimate a transfer function model for the tubing assembly. This was done by comparing the spectral content of an untubed Kulite sensor and a micro-pitot sensor in the SSJ. The transfer function model was then used to develop the compensator improving measurements made with the micro-pitot up to 50 kHz. The micro-pitot traverse system was then used in a series of tests in the BAM6QT to validate its ability to function as designed. The traverse needed to provide a rigid platform for the micro-pitot probes during tunnel operation. The deflection of the pitot head was recorded using a shadowgraph system. This allowed real time measurements for the deflection of the pitot head during tunnel operation to be taken. These measurements were compared to theoretical calculations to ensure deflections were within acceptable limits. Also, of key importance was the survivability of the traverse system after repeated runs in the BAM6QT. This focused on the ability of the traverse to continue providing movement in all three-directions and its ability to resist wear in the tunnel environment. The only cause for concern noted over the course of three tunnel entries centered around the motor used for wall-normal movement. This motor suffered repeated damage impairing the traverses ability to function as intended. Observations regarding this issue and solutions implemented to mitigate the impact of this damage are discussed. Finally, the micro-pitot was combined with the traverse system and used in conjunction with surface mounted sensors on a axisymmetric cone to measure traveling cross-flow instabilities. Damage to Kulites needed for the micro-pitot prohibited three sensors from being used in the tunnel. For this reason only propagation angles and phase speed calculations for traveling cross-flow waves were calculated using the surface mounted sensors. However, one micro-pitot sensor was used to measure spectral content near the surface mounted sensors. The spectral content of the micro-pitot was compared to the surface mounted sensors in order to validate that the micro-pitot could measure the desired instability once more are acquired </p>

A parametric study was performed to observe and quantify the effect of varying turbulence intensities on separation and performance in a compressor cascade at low Reynolds numbers. Tests were performed at 25° and 37.5° stagger angle, negative and positive angles of incidence up until the point of full stall, Reynolds numbers from 6 x 104 to 12.5 x 104, and turbulence intensities from approximately 0.7% – 8%. Additionally, oil flow techniques were combined with static tap data to visualize the boundary layer characteristics at various test conditions. The overall performance of the cascade was presented and evaluated through mass-averaged total pressure loss coefficients. The results of the study showed that the best efficiency (lowest pressure loss coefficient) was determined by separation characteristics for any angle of attack. While adding turbulence generally delayed separation, in some cases, adding turbulence to a separated airfoil resulted in decreased performance. Very similar separation characteristics were observed for the full range of Reynolds numbers and stagger, with the higher stagger setting giving slightly better performance. It was shown that a large percentage of total pressure losses can be recovered by applying the appropriate turbulence intensity at any angle of attack, which is relevant to possibilities for active control of such flows. / Master of Science

An investigation of ion flow control was performed to determine the effect of a positive, DC corona discharge on the boundary layer profile along a flat plate and to examine its ability to prevent separated flow in a low-speed compressor cascade. Flat plate tests were performed for two electrode configurations at free-stream velocity magnitudes of 2.5, 5, 7.5, and 10 m/s. Boundary layer velocity profile data was taken to measure the performance of the electrode pairs. Ion flow control was also tested in the compressor cascade for a stagger angle of 25° at angles of attack equal to 6° and 12°. The cascade tests were performed at free-stream velocities of 5 and 10 m/s. Static tap data was used to characterize separated flow behavior and the effect of ion flow control on flow reattachment. A computational model was developed using the commercial CFD software Fluent. This model simulates ion flow control as a body force applied to the flow through user-defined functions. The study showed that the corona discharge has the ability to increase near-wall velocities and reduce the thickness of the boundary layer for flow over a flat plate. Ion flow control successfully prevented trailing edge separation in a compressor cascade for angles of attack of 6° and 12°; however, the flow control scheme was not able to prevent leading edge separation for angle of attack equal to 12°. The ion flow control CFD model accurately predicted flow behavior for both the flat plate and cascade experiments. The numerical model was able to simulate the boundary layer velocity profiles for flat plate tests with good accuracy, and was also able to predict the flow behavior over a compressor blade. The model was able to show the trends of separated and reattached flow over the blade surface. / Master of Science

<p dir="ltr">The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) was used for a set of experiments studying the effect of isolated roughness elements on boundary-layer transition on a 7° half-angle cone. In quiet flow, the cone was tested at Reynolds numbers of 7.4 × 10e6 /m, 10.2 × 10e6 /m, and 13.0 × 10e6 /m. Tests were also completed at Re = 11.0 × 10e6 /m in noisy flow to examine the effects of freestream noise. The cone was set at both 0° and 6° angle of attack and an isolated, square trip oriented like a diamond with respect to the flow direction was attached before each set of runs. </p><p dir="ltr">Using infrared thermography and pressure transducers, the location of transition onset was estimated for each test. The results followed all expected trends: transition moved upstream as trip height increased, transition occurred earlier at higher freestream Reynolds numbers, and transition was significantly delayed in quiet flow compared to noisy flow. Mean flow solutions were generated to calculate correlation values commonly used to predict transition. Theexperimentaldatawasthenusedinconjunctionwiththesecorrelationvalues to identify a range of critical values that could be used to predict transition behavior. </p><p dir="ltr">Additionally, a z-type schlieren setup was developed for the BAM6QT. Various components were upgraded and standard procedures for aligning the system were developed. A new pulsed laser and high-speed camera were integrated into the system to enable schlieren imaging at up to 1.75M fps. The final configuration allows the schlieren system to be used for various applications with minimal adjustments, and has been utilized in many research projects in the BAM6QT.</p>

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