<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>

<div>A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel</div><div>at 6° angle of attack, extending several previous experiments on the growth and breakdown of</div><div>stationary crossflow instabilities in the boundary layer. </div><div><br></div><div>Measurements were made using infrared</div><div>imaging and surface pressure sensors. Detailed measurements of the stationary and traveling</div><div>crossflow vortices, as well as various secondary instability modes, were collected over a large</div><div>region of the cone.</div><div><br></div><div>The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was</div><div>adapted for application to crossflow work. It was demonstrated that the roughness elements were</div><div>the primary factor responsible for the appearance of the specific pattern of stationary streaks</div><div>downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness</div><div>insert was created with a high RMS level of normally-distributed roughness to excite the naturally</div><div>most-amplified stationary mode.</div><div><br></div><div>The nonlinear breakdown mechanism induced by each type of roughness appears to be</div><div>different. When using the discrete RIM roughness, the dominant mechanism seems to be the</div><div>modulated second mode, which is significantly destabilized by the large stationary vortices. This</div><div>is consistent with recent computations. There is no evidence of the presence of traveling crossflow</div><div>when using the RIM roughness, though surface measurements cannot provide a complete picture.</div><div>The modulated second mode shows strong nonlinearity and harmonic development just prior</div><div>to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal</div><div>band, leading to a peak in the heating simultaneously with the peak amplitude of the measured</div><div>secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic</div><div>heating pattern is reminiscent of experiments on a flared cone and earlier computations</div><div>of crossflow on an elliptic cone.</div><div><br></div><div>When using the distributed roughness there are several differences in the nonlinear breakdown</div><div>behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,</div><div>which is expected given computational results. Furthermore, the traveling crossflow waves become</div><div>very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there</div><div>appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability</div><div>under the upwelling of the stationary vortices. The traveling crossflow and the secondary</div><div>instability interact nonlinearly prior to breakdown.</div>

Thesis (Ph.D.) - University of Queensland, 2003. / Includes bibliography.

<div>A 7° half-angle cone with a highly-swept fin was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel at 0.0° angle of attack. Previous measurements of the surface heat transfer using temperature sensitive paint revealed heating streaks on the cone surface related to streamwise vortices generated by the fin shock. High-frequency measurements of the cone-surface pressure fluctuations revealed that transition occurs in the streak region at sufficiently-high freestream unit Reynolds numbers under quiet flow. In this work, high-resolution measurements of the surface heat transfer are obtained using infrared thermography and a polyether-ether-ketone wind-tunnel model. In addition, a novel model design made it possible to measure pressure fluctuations throughout the streak region on the cone surface.</div><div><br></div><div>A slender cone with a sharp nosetip and a fin swept back 75° with a 3.18 mm leading-edge radius served as the primary geometry for this work. Two laminar heating streaks</div><div>were measured on the cone surface. These travel along a line of nearly-constant azimuth. A hot spot develops in the streak farthest from the fin, which then moves upstream with increasing freestream Reynolds number. Downstream of this hot spot, the streaks begin to spread in azimuth. The heat transfer along the outer streak shows a threefold increase near the hot spot before decreasing back to nearly two times the laminar streak heating. The amplitude of the pressure fluctuations increases simultaneously with the heat transfer, reaching a peak of nearly 9% of the Taylor-Maccoll pressure for a 7° straight cone. Power spectral densities calculated from these fluctuations demonstrate spectral broadening, which is indicative of boundary-layer transition. Using surface-pressure-fluctuation and heat-flux measurements, transition onset was estimated to occur at an axial length Reynolds number of 2.2×10⁶. Pressure sensors that were rotated through the streak region showed that multiple instabilities amplify between the heating streaks, upstream of the transition onset location. Downstream of transition onset, the highest-amplitude instabilities are localized to the hot spot in the outer streak. The effect of freestream noise on transition was also investigated with this geometry. Under conventional noise levels, transition onset was estimated to occur at an axial length Reynolds number of 0.93×10⁶, and only one instability was measured in the streak region with a frequency similar to the second-mode instability.</div><div><br></div><div>Four configurations were tested to investigate the effect of fin sweep and nosetip bluntness under quiet flow. Fins with 70° and 75° sweep were each tested with nominally sharp and 1-mm-radius nosetips. Increasing fin sweep was shown to move the heating streaks on the cone closer to the fin and to decrease the peak-to-peak spacing of the streaks. In addition, transition onset occurred at lower freestream unit Reynolds numbers for the 70° sweep case. Increasing nosetip radius had little effect on the heating streaks, other than to delay the transition location. A blunt nosetip was shown to delay transition more for the 75° sweep fin as compared to the 70° fin. Similar instabilities were measured for all four of the configurations in this work. The frequency of the instabilities appears to be correlated with the peak-to-peak distance of the heating streaks, which can be viewed as an indirect measurement of the vortex diameter.</div><div><br></div><div>Lastly, the first quantitative measurements of heat transfer on the fin were made using the infrared thermography apparatus. Peak heating on the fin, not including the leading edge, is lower than peak heating rates on the cone. One broad heating streak was measured close to the corner, and smaller low-heating streaks were measured farther outboard. The heating within the streak closest to the corner was shown to agree well with a fully-laminar computed basic state, indicating that the flow on the fin is laminar up to at least 6.31×10⁶ m⁻¹. Using miniaturized Kulite sensors, pressure fluctuations were measured at twelve locations on the fin surface. No obvious conclusions could be drawn from these Kulite measurements, and there is no clear indication that transition occurs on the fin within the maximum quiet</div><div>freestream conditions.</div>

<div>Wind-tunnel experiments with a sharp 7-degree half-angle cone and a 33% scale Boundary Layer Transition (BOLT) model were performed in the Boeing/AFOSR Mach 6 Quiet Tunnel to investigate the effects of forward- and backward-facing steps on boundary-layer instability and transition. Each model was modified to include intentional steps just downstream of the nosetip. Experiments were performed at different freestream Reynolds numbers and varying step sizes. Infrared thermography was used to calculate surface heat transfer, and high-frequency pressure sensors were used to measure pressure fluctuations. A replica measurement technique was used to accurately measure step heights on the BOLT flight vehicle and the wind tunnel model.</div><div><br></div><div>A 7-degree half-angle cone was tested at 0-degree and 6-degree angles of attack. Step heights ranged from 0.610 mm to 1.219 mm. At a 0-degree angle of attack, no significant increases in heat transfer were observed with any of the forward- or backward-facing steps. However, a 250 kHz instability was measured with the forward-facing steps. Growth of the instability was similar to a second-mode. At a 6-degree angle of attack, an increase in heat transfer was observed on the windward ray with the forward-facing steps. Sharp increases in heating rates and increased pressure fluctuations were indications of boundary-layer transition. Elevated heating rates and pressure fluctuations were not measured with the backward-facing steps.</div><div><br></div><div>The BOLT model was tested at 0-degree, 2-degree, and 4-degree angles of attack and 2-degree and 4-degree yaw angles. Step heights ranged from 0.076 mm to 1.016 mm. At a 0-degree angle of attack and 0-degree yaw angle, thin wedges of heating were observed with the backward-facing steps. Instabilities were measured near these wedges of heating and are thought to be caused by a secondary instability. The effects of the steps were magnified on the windward side of the BOLT model at angles of attack. Wedges of heating were wider and more intense. At higher angles of attack, the onset of heating was further upstream. Sensors near and directly underneath the wedges of heating measured pressure fluctuations that were indicative of a turbulent flow. Wedges of heating were also observed at a 4-degree yaw angle, but only with the 1.016 mm backward-facing step.</div>

<p> Rising costs and volatility in aviation fuel and increased regulations resulting from climate change  concerns have driven gas turbine engine manufacturers to focus on reducing fuel consumption.  Implementing centrifugal compressors as the last stage in an axial engine architecture allows for  reduced core diameters and higher fuel efficiencies. However, a centrifugal compressor's  performance relies heavily on its stationary diffusion system. Furthermore, the highly unsteady  and turbulent flow field exhibited in the diffusion system often causes CFD models to fall short of  reality. Therefore, rapid validation is required to match the speed at which engineers can simulate  different diffuser designs utilizing CFD. One avenue for this is through the use of additive  manufacturing in centrifugal compressor experimental research. This study focused on implementing a new generation of the Centrifugal Stage for Aerodynamic  Research (CSTAR) at the Purdue Compressor Research Lab that utilizes an entirely additively  manufactured diffusion system. In addition, the new configuration was used to showcase the  benefits of additive manufacturing (AM) in evaluating diffusion components. Two diffusion  systems were manufactured and assessed. The Build 2 diffusion system introduced significant  modifications to the diffusion system compared to the Build 1 design. The modifications included changes to the diffuser vane geometry, endwall divergence, and increased deswirl pinch and vane  geometries. The Build 2 diffusion system showed performance reductions in total and static  pressure rise, flow range, and efficiencies. These results were primarily attributed to the changes  made to the Build 2 diffuser. The end wall divergence resulted in end wall separation that caused  increased losses. The changes to the diffuser vane resulted in increased throat blockage and lower  pressure rise and mass flow rate. In addition to the experimental portion of this study, a computational study was conducted to study  the design changes made to the Build 2 diffusion system. A speedline at 100% corrected rotational  speed was solved, and the results were compared to experimental data. The simulated data matched  the overall stage and diffusion system performance relatively well, but the internal flow fields of  the diffusion components, namely the diffuser, were not well predicted. This was attributed to  16 using the SST turbulence model over BSL EARSM. The BSL EARSM model more accurately  predicted the diffuser flow field to the SST model.  </p>

<p>Hypersonic flow over two generic missile configurations was investigated using CFD meth-</p> <p>ods. CFD results were compared with experimental results obtained by the hypersonic flight</p> <p>lab at Texas A&M University. Baseline RANS computations involving the missile configurations at a zero deg angle-of-attack were performed, along with computations at higher angles-of-attack. As the angle-of-attack was increased, complex vortex interactions were observed in the region between the fins. Increasing the angle-of-attack generally increased heating on the windward side of the missile geometries, especially on wall surface regions</p> <p>adjacent to the fin-root vortices. The results presented highlight observed fin region vortices and regions of intense heating on the body surface. DES simulations methods were also used to explore unsteady aspects of flow around the two generic missile configurations through time-accurate CFD simulations. Power spectral plots were generated to quantify the dominant frequencies of large-scale unsteadiness.</p>

<div>Many important measurements of low-amplitude instabilities related to hypersonic laminar-turbulent boundary-layer transition have been successfully performed with 1-MHz PCB132 pressure sensors. However, there is large uncertainty in measurements made with PCB132 sensors due to their poorly understood response at high frequency. The current work continues efforts to better characterize the PCB132 sensor with a low-pressure shock tube, using the pressure change across the incident shock as an approximate step input. </div><div> </div><div> New vacuum-control valves provide precise control of pre-run pressures in the shock tube, generally to within 1\% of the desired pressure. Measurements of the static-pressure step across the shock made with Kulite sensors showed high consistency for similar pre-run pressures. Skewing of the incident shock was measured by PCB132 sensors, and was found to be negligible across a range of pressure ratios and static-pressure steps. Incident-shock speed decreases along the shock tube, as expected. Vibrational effects on the PCB132 sensor response are significantly lower in the final section of the driven tube.</div><div> </div><div> Approximate frequency responses were computed from pitot-mode responses. The frequency-response amplitude varied by a factor of 5 between 200--1000 kHz due to significant resonance peaks. Measurements with blinded PCB132 sensors indicate that the resonances in the frequency response are not due to vibration. </div><div> </div><div> Using the approximate frequency response measured with the shock tube to correct the spectra of wind-tunnel data produced inconclusive results. Correcting pitot-mode PCB132 wind-tunnel data removed a possible resonance peak near 700 kHz, but did not agree with the spectrum of a reference sensor in the range of 11--100 kHz. </div>

<div>Good wind-tunnel flow quality characteristics are vital to using test data in the aerodynamic design process. Spatially uniform velocity profiles are required to avoid yaw and roll moments that would not be present in real flight conditions. Low turbulence intensity levels are also important as several aerodynamic properties are functions of turbulence intensity. When measuring mean flow and turbulence properties, hot-wire anemometry offers good spatial resolution and high-frequency response with a fairly simple operation, and the ability to make near-wall measurements. Using hot-wire anemometry, flow quality experiments were conducted</div><div>in a closed-circuit wind tunnel with a test section that has a cross section area of 1.2 m x 1.8 m (4 ft. x 6 ft.). The experiments included measurements of flow velocity and turbulence intensity variation over the test section cross-section, spatial and temporal temperature variation, and</div><div>boundary layer measurements. The centerline velocity and turbulence intensity were also measured for flow speeds ranging from 13 to 43 m/s.</div>

<div> Large thermal gradients can cause very nonintuitive effects in the flowfield, as flow motion and even a force (often referred to as a Knudsen thermal force) can be induced even with a freestream velocity of zero. These flows can be exploited on the microscale, where temperature gradients of 10E6K/m are achievable. These flows have been studied experimentally many times, and it has been shown that Knudsen forces have a bimodal relationship with pressure, where the peak is in the transitional flow regime. It has also been shown that these thermal gradients cause thermal diffusion, or species separation in a mixture.</div><div> </div><div> A MEMS based device called the Microscale In-Plane Knudsen Radiometric Actuator (MIKRA) was developed to use Knudsen forces to calculate pressure and gas composition. The direct simulation Monte Carlo (DSMC) method was used to analyze the device to calculate the device forces and calculate the flowfield. DSMC proved to be a reliable method of simulating these types of flows, as the force results agreed well with experiments, and the DSMC results matched the results of other numerical methods.</div><div> </div><div> N2 and H2O mixtures were also simulated, and it was shown that the force is sensitive to the composition. At the same pressure, the force is larger for mixtures dominated by N2. Heat flux is also larger for N2 dominated flows.</div>