<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>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>

<p dir="ltr">Instantaneous fuel regression rate within a solid fuel ramjet combustor was characterized using X-ray radiography and ultrasonic transducer measurements. Experiments were performed with cylindrical, center-perforated hydroxyl-terminated polybutadiene (HTPB) fuel grains at three mass fluxes (407-561 kg/m2-s) with consistent inlet total temperatures and chamber pressures. Ultrasonic transducer measurements demonstrated changes of web thickness ranging from 7.50-9.85 mm and regression rate measurements ranging from 1.35-1.74 mm/s. Local maxima of change in web thickness due to flow reattachment and erosive burning were consistently measured with the ultrasonic transducers. Changes in port radius on the order of 8-9 mm and regression rates of approximately 1.25 mm/s were deduced from the X-ray radiography images. Structure of the flow reattachment region was evident in measurements from the X-ray radiography images captured near the combustor entrance while images captured at the mid-length of the combustor exhibited more uniform fuel regression profiles. Ultrasonic measurements of change in web thickness were consistently greater in magnitude relative to X-ray radiography measurements. X-ray radiography imaging allowed for the more accurate measurement of fuel regression with the greatest axial spatial resolution while ultrasonic transducer measurements yielded the greatest radial spatial resolution. The change in web thickness calculated with weight-based techniques yielded smaller magnitude measurements of change in web thickness relative to X-ray radiography.</p><p dir="ltr">Time-dependent measurements of web thickness and regression rate along the port of aluminum-loaded and boron carbide-loaded, hydroxyl-terminated polybutadiene (HTPB) fuel grains were measured in a solid fuel ramjet combustor with X-ray radiography. The combustor was operated at three mass flux conditions, ranging from 397-532 kg/m2-s, with consistent chamber pressures and upstream-of-combustor total temperatures of 1313 kPa and 748 K, respectively. A cross-correlation-based edge detection scheme was used to extract the fuel grain edges within X-ray radiography images collected at 15 Hz. Cross-section photographs of the post-combustion fuel grain surfaces exhibited evidence of flow reattachment and large aft-end regression. Aluminized fuel grains exhibited average weight-based regression rates of 1.29-1.48 mm/s, and boron carbide-loaded fuel grains yielded average regression rates of 1.21-1.38 mm/s. Head-end X-ray measurements of change in port radius indicated flow reattachment, particularly for the bottom (theta = 180) edge of the fuel grain. The absolute maximum of change in port radius, which ranged between 8.56-10.31 mm for aluminized fuel grains and 8.22-9.40 mm for boron carbide-containing fuel grains, did not always coincide with the flow reattachment location. Time-averaged regression rate profiles measured with X-ray radiography were relatively uniform along the port axis but smaller in magnitude compared to the weight-based measurements; 1.17-1.35 mm/s for the aluminum-loaded fuel grains and 1.07-1.24 mm/s for the boron carbide-loaded fuel grains. Pre-ignition fuel regression, on the order of 1.5 mm, was determined to be the cause of the over-prediction of regression rate by weight-based measurements compared to X-ray measurements.</p><p dir="ltr">The weight-based average regression rates measured in tests conducted with the axisymmetric solid fuel ramjet test article in its various configurations were compared to quantify the effects of average port air mass flux, bypass air addition, carbon black addition, and metal particle addition on regression rate. Baseline tests without an aft-mixing section or bypass air addition fuel grains containing carbon black yielded a regression rate coefficient of a = 5.33E-2 and an exponent of n = 0.50 for p4 = 1179-1298 kPa. Including an aft-mixing section without bypass air addition yielded regression rates of 0.94-1.04 mm/s due to the increased residence time. Bypass air addition of 14\% bypass ratio reduced the regression rate to 0.83-0.92 mm/s, and 30% bypass ratio reduced the regression rate to 0.80-0.82 mm/s. For otherwise equal tests, adding carbon black to the fuel grain increased the regression rates from 0.76-0.78 mm/s to 0.83-0.92 mm/s (6-21%). Aluminized fuel grains exhibited an increase in regression rate coefficient over the baseline fuel grains from a = 5.33E-2 to a = 6.30E-2 (18%), but the regression rate exponent remained at n = 0.50. Boron carbide (B4C) addition reduced the regression rate exponent to n = 0.46 but increased the regression rate coefficient to a = 7.55E-2; a 42% increase.</p><p dir="ltr">A simplified solid fuel ramjet combustion model which includes (1) turbulent heat convection, (2) radiation, (3) radiation-coupled surface blowing, (4) unsteady sub-surface heat conduction, (5) solid fuel regression, (6) gas-phase combustion, and (7) fuel port hydrodynamics was developed for regression rate prediction over a range of combustor geometries and operating conditions. Turbulent convection was modeled with empirical correlations relating non-dimensional boundary layer transport numbers. Radiative heat transfer was estimated using modified empirical correlations for radiation in a slab hybrid rocket combustor. Hybrid rocket combustion theory was used to model surface blowing. The condensed-phase heat transfer was modeled by solving the unsteady, variable thermophysical property, regressing surface heat equation with an explicit time-integration, finite volume scheme on a non-uniform grid. A general Arrhenius expression was used to estimate the fuel regression rate. Chemical equilibrium calculations for a stoichiometric HTPB/air diffusion flame were used to model the gas-phase combustion. The port gas dynamics were modeled with compressible flow ordinary differential equations. The results of these individual physical processes were examined in detail for a high mass flux (G_air = 561 kg/m2-s) case. Experiments performed in the axisymmetric solid fuel ramjet combustor were simulated in the model, which yielded a lower regression rate versus mass flux exponent of n = 0.39 compared to the experimentally-obtained n = 0.50. A larger parameter sweep of the model yielded a mass flux exponent of n_1 = 0.30, a pressure exponent of n_2 = 0.04, and an inflow total temperature exponent of n_3 = 0.39. These exponents are less than those observed in other works, but the model successfully captured the relative influence of mass flux, chamber pressure, and inflow total temperature.</p><p dir="ltr">A combustion diagnostic consisting of X-ray radiography and thermocouples embedded within the fuel grain was successfully applied and demonstrated in a solid fuel ramjet slab combustor. One representative test condition with an air mass flowrate of 1 kg/s, an upstream-of-combustor static pressure of 560 kPa, and an upstream-of-combustor total temperature of 639 K was examined. Changes in web thickness of approximately 4 mm and steady-state regression rates of 0.35 mm/s were measured at the thermocouple locations. Condensed-phase temperature measurements yielded fuel grain surface temperatures of 820 K and temperature profiles which were compared to theoretical Michelson profiles. The Michelson profile closely matched the thermocouple-measured temperature profile at one axial location. Sub-surface conductive heat fluxes of 0.35 MW/m2, heat fluxes required to vaporize solid fuel of 0.60 MW/m2$, and surface heat fluxes of 0.95 MW/m2$ were estimated using the condensed-phase temperature profiles.</p>

<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>

<p dir="ltr">Experiments were performed in the Boeing/AFOSR Mach-6 Quiet Tunnel to study the effect of flight-derived discrete roughness elements repeated in an axisymmetric pattern near the nose of a sharp 7° cone. The aim of the roughness array was to simulate natural vehicle roughness and attempt to introduce a deterministic roughness pattern with the ability to cancel out the instabilities caused by the natural roughness. The cone was pitched at a 6° of attack to determine the three-dimensional flow field effects of the roughness elements. Tests were also ran at 0° of attack for comparison. Quiet flow testing included the designed-for freestream unit Reynolds number of 10.8x10⁶, and Reynolds numbers above and below. In noisy flow, comparable Reynolds numbers were also tested at to isolate the effects of noise in a conventional flow wind tunnel.</p><p dir="ltr">Infrared thermography and surface pressure sensors were used to document the behavior of the boundary layer. It was found that the roughness pattern was in general unsuccessful in controlling the added boundary layer instabilities as intended at 6° of attack, but it did create different instability amplitudes and heating patterns. Additionally, it was determined to reduce Mack's second-mode instability amplitudes at 0° of attack.</p><p dir="ltr">Additionally, work was done to document and characterize the roughness patterns found on samples of hypersonic glide vehicles PRIME (SV-5D or X-23) and ASSET (ASV-3). These samples were taken in the form of molded impressions of the surface which were able to be analyzed with an optical profilometer and considered for future experimental distributed roughness studies.</p>

<p dir="ltr">Force and moment measurements are valuable tools for evaluating designs in a wind tunnel environment. In fact, this type of research has been conducted ever since the earliest wind tunnels were in use. Load measurement techniques are complicated by hypersonic wind tunnel designs, which often have much shorter test times due to the immense stagnation pressures that are used. Previous research had been conducted once before in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University using a six-component moment balance. This initial testing utilized a balance with maximum load limits which far exceeded the loads experienced within the BAM6QT. Because of this, much of the data collected was imprecise. </p><p dir="ltr">Testing was conducted in the BAM6QT using three different balances - a five-component foil, five-component semiconductor, and six-component semiconductor balance. Data were taken for a variety of geometry configurations over a range of total pressures. All data were taken at 0 degree angle of attack. The two geometries used most commonly were the 1 inch diameter blunt nose-tip, 7 degree half-angle, 1.75 inch base diameter cone, with either a 20 degree or 30 degree curved ramp. An additional sharp nose-tip configuration was also used. Results for multiple load components were calculated during each run and compared between each balance type. Results were compared to the surface panel method results of CBAERO, which uses either modified Newtonian theory or the tangent cone method to compute loads. </p><p dir="ltr">Results between each balance type were similar and generally in good agreement. The semiconductor balance designs showed considerably less noise than the foil design. Results of CBAERO matched well with the balance data, with a baseline comparison of the plain blunt cone showing a maximum difference of 12% for the modified Newtonian theory. The more complicated ramp geometries, which exhibited regions of flow detachment, agreed surprisingly well with CBAERO results, despite the more complicated flow phenomena, which was unexpected. The best agreement was generally seen in the cases where the large 30 degree ramp was used, while the sharp nose-tip configuration produced the worst agreement. Overall, CBAERO proved valuable as an approximate method for determining the general magnitude of loads. The sting, used to mount the model in the wind tunnel, was found to drive the oscillation frequency of the model-sting system. The longer sting and less stiff balance used on the six-component system likely contributed to lower oscillation frequencies which affected the results for the pitching moment and normal force. The relationship between startup and running loads was also investigated and a startup-to-running load ratio of 5 to 20 was determined, depending on the load component and geometry.</p>

<p dir="ltr">The computational campaign reported in this thesis focuses on a series of experiments at Mach 2.85 carried out in the 1980s at NASA Ames Research Center on a set of cylinder / skewed flare configurations designed to produce highly three-dimensional shockwave / boundary-layer interactions in the absence of end-wall effects. Computations carried out in that era were unable to match the experimental results using the numerical techniques, turbulence models, and grid resolution available at the time. In the present work, newer Reynolds-averaged Navier-Stokes and detached eddy simulation methods have been applied to these flows, and relatively good agreement has been obtained with the experimental data. Difficulty in capturing the correct separation bubble size was encountered with initial detached eddy simulations, but the introduction of resolved turbulence via a boundary layer trip produced much better results. This thesis reports on results obtained for four inclination angles (0 deg, 5 deg, 10 deg, and 23 deg) of the skewed flare. Detached eddy simulation is seen to be an economical alternative to large eddy simulation for capturing many features of large-scale separation unsteadiness over long time intervals at true Reynolds number.</p>

<p dir="ltr">Prediction of boundary-layer transition is a critical element of hypersonic vehicle design</p><p dir="ltr">due to the impact transition has on boundary-layer separation, heat transfer, and aerodynamic</p><p dir="ltr">control. Transition is affected by many factors including surface roughness. The</p><p dir="ltr">roughness on a hypersonic vehicle can cause a boundary-layer to become turbulent. However,</p><p dir="ltr">there is a limited understanding of the impacts that roughness can have, and the conditions</p><p dir="ltr">under which it is important.</p><p dir="ltr">The rocket-sled track at Holloman Air Force Base was selected as a ground-test facility</p><p dir="ltr">for transition measurements. The present work is about understanding the mechanism of</p><p dir="ltr">transition on blunt ogives or blunt cones with moderate nose radii, as it appears that nosetip</p><p dir="ltr">roughness affects boundary-layer transition on the afterbody for moderate nose radii. A</p><p dir="ltr">single test-track shot is to be executed for a blunt ogive to determine if the test track can</p><p dir="ltr">make useful measurements of boundary-layer transition.</p><p dir="ltr">Initially, the present research used a boundary-layer solver to estimate target roughnesses</p><p dir="ltr">that would be applied to the nose tip. Preliminary analysis was conducted on test cases for</p><p dir="ltr">sharp cones and blunt cones. However, due to time constraints, the target roughnesses were</p><p dir="ltr">then estimated with a higher fidelity code by Brad Wheaton of JHU APL. Two separate</p><p dir="ltr">roughness targets were established for the upper and lower sides of the hemispherical nosetip.</p><p dir="ltr">The focus of this work then shifted to measurements of the roughness that was applied</p><p dir="ltr">by others to the hemisphere nose tip for a blunt ogive. Utilizing the Zygo ZeGage 3D optical</p><p dir="ltr">profiler, roughness scans were collected both directly under the profiler head and indirectly</p><p dir="ltr">using rubber molds. Profilometer measurements were also recorded. Twelve iterations were</p><p dir="ltr">completed to allow the polisher to develop appropriate procedures for applying the roughness,</p><p dir="ltr">given the material and curvature. The first five iterations involved roughness applied to</p><p dir="ltr">cylindrical-shaped test areas. After achieving the target roughnesses on these test areas,</p><p dir="ltr">the hemispherical ends of test specimens were then polished and measured until both the</p><p dir="ltr">rough and smooth halves met the roughness target. During this time, the three roughness measurement</p><p dir="ltr">techniques were refined until good agreement was reached between them. When the roughness-application and </p><p dir="ltr">roughness-measurement techniques were sufficiently mature,</p><p dir="ltr">the actual blunt-ogive nose tip was then polished until the roughness target was achieved.</p>

<p dir="ltr">The Boeing-AFOSR Mach-6 Quiet Tunnel (BAM6QT) was used for a set of experiments aiming to characterize the stability regimes of stagnation point injection in noisy and quiet flow across an array of different injected gases. Four gases were used in this experiment: air, helium, carbon dioxide, and argon. These gases were injected at varying thrust coefficients, ranging from 0.0516 to 0.5666, using a 7 degree half-angle cone with a 19 mm radius spherical nose and a single 1.93 mm-radius sonic jet in the center of the model. The primary data collected consists of schlieren images gathered at a sample rate of 76 kHz. These data were then analyzed using a shock tracking software to measure the physical locations of flow features as well as through spectral proper orthogonal decomposition (SPOD) to analyze specific modes in the flow.</p><p dir="ltr">Through this analysis, it was observed that three principle modes exist in stagnation point injection regardless of the injecting gas: a high frequency vortex-coupled mode, a low frequency Mach-shock-rigid mode, and a hybrid mode residing between these two modes. The first two modes were observed in all stability regimes, whereas the hybrid mode was only observed in the bifurcated regime. Furthermore, the unsteady regime was observed to be mostly characterized by this first, vortex-coupled mode. Conversely, the steady regime was observed to be driven by the Mach-shock-rigid mode instead. This transition was measured to occur as the thrust coefficient was increased.</p><p dir="ltr">This research also found that freestream noise resulted in an amplified and widened frequency range within the Mach-shock-rigid mode. This same freestream noise did not appear to have an impact on the other two principle modes; however, in some cases the noise produced in the Mach-shock-rigid mode due to this freestream noise did in fact mask the other principle modes.</p><p dir="ltr">Lastly, it was observed that the thrust coefficient, in and of itself, is not the sole indicator of stability in stagnation point injection. Across the different injected gases in this research, transition between the stability regimes did not in fact occur at a constant thrust coefficient value. Additionally, even within the same injected gas, this transition did not occur at the same thrust coefficient value between noisy and quiet runs—indicating an effect of freestream noise on stability.</p>