Direct experimental measurements of double-cone shock-wave/boundary-layer interaction dynamics at Mach 7.2 were collected using a schlieren-based shock-detection algorithm at the University of Texas at San Antonio. The points of flow separation and reattachment were spatially and spectrally evaluated at 300 kHz for time-resolved characterization of the unsteadiness of this flow field. The separation bubble was found to rapidly expand and contract about a large mean length. Spectral analysis of the velocity fluctuations of this motion revealed this behavior to be a turbulent phenomenon, with spectral energy trends consistent with the Kolmogorov theory for the turbulent energy cascade.

Accurate high-temperature measurements are essential for optimizing and validating chemical kinetic models. In high-enthalpy shock-tube experiments, even small temperature uncertainties can cause large errors in inferred rate constants. We introduce a laser absorption-based sensing technique-multi-species Boltzmann regression-based on a multi-variable regression of rovibrational state populations across multiple species to determine temperature in uniform flows. By exploiting the homogeneity of regression slopes across multiple species, this method provides a unified framework for inferring temperature and species mole fractions relative to two-color and single-species Boltzmann regression thermometry methods applied to the same absorption features. Validation experiments were performed in The University of Texas at San Antonio's optically-accessible high-enthalpy shock tube, targeting carbon dioxide and carbon monoxide as well as nitrous oxide and nitric oxide, with relevance to hydrocarbon combustion and nitrogen-based propellant chemistry. For non-reacting shock-heated carbon oxide mixtures, the proposed technique produced temperature uncertainties comparable to single-species Boltzmann regression thermometry and improved agreement with predicted temperatures relative to two-color methods, while for ethylene-oxygen oxidation comparable uncertainties were observed. However, for nitrous oxide thermal decomposition, multi-species Boltzmann regression (MSBR) did not achieve lower uncertainties than two-color thermometry, highlighting the importance of strategic line selection in sensing applications. To further examine this behavior, supplemental measurements of hydrogen cyanide and acetylene were performed in an ambient-temperature multipass optical gas cell, demonstrating reduced temperature uncertainties relative to two-color and single-species Boltzmann regression thermometry. A corresponding uncertainty analysis for both single- and MSBR shows that temperature uncertainty depends strongly on both experimental precision and the statistical distribution of the selected transitions' lower-state energies. These results demonstrate that multi-species Boltzmann regression is most advantageous when transitions from multiple species increase the statistical spread of lower-state energies, thereby improving regression leverage and temperature robustness, supporting optimization and validation of predictive combustion models.

Surface pressure distributions of a hypersonic shock-wave/boundary-layer interaction have been investigated at an average Re = 28.2 x 106m-1 in the Mach 7.2 Wind Tunnel facility located at the University of Texas at San Antonio. A ruthenium-based fast-responding pressure-sensitive paint, made in-house, was applied to a canonical wall-to-wall flat plate with a 33 degrees compression ramp. A static calibration of the paint's response was used to extract global surface pressure measurements of the hypersonic shock-wave/boundary-layer interaction. The 33 degrees ramp experienced high-pressure regions and "G & ouml;rtler-like" vortices were observed with a spacing of 3-6 599. The vortex structure spacing was further investigated using surface oil flow visualization, giving a spacing of 2-4 599. The separation region showed lower pressures than the ramp, and the lowest pressures recorded appeared from the leading edge of the model to the separation region. In this study, a statistical and spectral analysis of the images was performed globally and at discrete locations along the model. The spectral analysis showed energy peaks on the ramp over a low-frequency band of Strouhal numbers between St5 0.00029-0.00147. Modal analysis was also performed to further examine the energy content of the structures and sub-structures that appeared on the compression ramp.

Molecular tagging velocimetry, a minimally intrusive laser diagnostic, was used to find quantitative off-body velocity measurements in the wake of both strut-mounted and free-flight spheres in hypersonic flow. Acetone gas seeded in the flow was excited using the fourth harmonic of a pulse-burst Nd:YAG laser. The experimental tests were validated and compared to simulations using a continuous Galerkin flow solver and an adaptive mesh refinement process. Near the centerline in the wake of the sphere, velocities between −105 and 65 m/s were observed due to reverse flow in the viscous shear layer. Similarly, the simulated data showed negative velocity values in the shear layer region. Outside of the shear layer, average velocities of 730 m/s were observed in both the experimental and simulated results. The overall average uncertainty for the strut-mounted and free-flight cases was estimated to have an accuracy of ±4% (±35.5m/s or 4.9%). This non-intrusive technique provides a velocity map of the wake behind a sphere under hypersonic conditions.

Hypersonic flight can revolutionize commercial transport, defense, and space exploration. However, hypersonic vehicles face severe flow disturbances that can degrade their structural integrity by generating high temperatures and acoustic loads. In this work, we investigate phononic metamaterials (PMs) as a novel approach for stabilizing these disturbances. PMs are engineered materials that exhibit frequency ranges, known as band gaps, where wave propagation is significantly inhibited. Our findings show that these band gaps effectively mitigate unsteadiness in hypersonic shockwave/boundary-layer interactions (SWBLIs). Specifically, we analyze a bilayer PM subsurface exposed to a Mach 7.2 cylinder-induced SWBLI, comparing its performance to a rigid wall control. Spectral analysis reveals that the PM selectively attenuates disturbances and significantly stabilizes flow structures away from the fluid-PM interface. These results highlight PMs as a transformative technology for enhancing hypersonic vehicle resilience and enabling reusable hypersonic systems, paving the way for safer, more efficient hypersonic flight.

In this study, acetone molecular tagging velocimetry (MTV) was employed to measure velocity in the far wake of both a free-flight and strut-mounted sphere in the Mach 7 wind tunnel facility at the University of Texas at San Antonio. The far wake, defined here as the region beginning approximately two body diameters downstream of the sphere, was characterized using MTV to obtain a two-dimensional velocity field and velocity fluctuations. For the free-flight case, the 2D velocity map showed turbulent structures forming after the neck of the inner far wake, where velocities were measured to be similar to 100 m/s. Outside the inner turbulent wake velocities increased to similar to 600 m/s. For the strut-mounted case, all profiles converged to a velocity of 470 m/s, corresponding to a shock observed in the schlieren images, likely originating from the SWBLI formed between the model and the strut. Additionally, planar laser scattering (PLS) of condensed carbon dioxide particles was used to visualize the wake structures in the spanwise direction of the spherical model in free-flight. Spanwise PLS effectively showed cold flow regions where the bow shock is formed in the outer wake, as well as hot flow regions within the inner wake.

A joint experimental/numerical study is carried out on flat-plate supersonic turbulent boundary layers that evolve in space at a freestream Mach number, M-infinity, of 2.86. The considered range of the friction Reynolds or Von Karman number, delta(+), is approximately 1000-2500. Those Reynolds numbers ensure an adequate separation between the turbulence scales in the inner and outer region of the turbulent boundary layer and are attainable by Direct Numerical Simulations (DNS). Although supersonic turbulent boundary layers have been extensively scrutinized in the past, these studies have been isolated; the present proposed DNS approach designed in harmony with high-speed wind tunnel experiments at similar Reynolds numbers will contribute to a better understanding of the transport phenomena under the influence of wall cooling conditions in zero-pressure gradient (ZPG) flows. A baseline case under wall adiabatic conditions is also projected. The expected research outcomes are threefold: (i) low/high order turbulence statistics computation in DNS including the near wall and buffer regions, (ii) experimental fluid velocity profiles away from the wall and wall pressure/temperature are used as validating tools for numerical predictions, (iii) assessment of an implicit Large Eddy Simulation (iLES) scheme at higher Reynolds numbers.

Shock-wave/boundary-layer interactions (SWBLI) are a common problem in supersonic and hypersonic flows, leading to separation and unsteady flow effects. In practical hypersonic applications ablative materials are sometimes used for thermal protection, and the rough surface caused by ablation will change the nature of the SWBLI. In this work, we seek to better understand the impacts of surface roughness by extending previous research on both a rough and smooth surfaced flat plate compression ramp using planar laser scattering (PLS). Streamwise imaging of a smooth flat plate has highlighted structures moving through the boundary layer into the reattachment shock, and skimming images of the flat plate showed long structures persistent in the flow as well as finger-like structures near the flat plate, around the recirculation zone. Surface skimming PLS at different heights within the boundary layer showed varying spanwise structures. The rough flat plate showed a turbulent boundary layer, and distributed turbulent structures across the entire test section with the skimming PLS data. Proper orthogonal decomposition on both smooth and rough flat plate cases showed similar structures under the reattachment shock, although differences are noted upstream due to the boundary layer state.

We report a novel "cone-ray" model of background-oriented schlieren (BOS) imaging that accounts for depth-of-field effects. Reconstructions of the density field performed with this model are far more robust to the blur associated with a finite aperture than conventional reconstructions, which presume a "thin-ray" pinhole camera. Our model is characterized and validated using forward evaluations based on simulated and experimental BOS measurements of buoyancy-driven flow and hypersonic flow over a sphere. Moreover, we embed the model in a neural reconstruction algorithm, which is demonstrated with a total variation penalty as well as the compressible Euler equations. Our cone-ray technique dramatically improves the accuracy of BOS reconstructions: the shock interface is well-resolved in all our tests, irrespective of the camera's aperture setting, which spans f-numbers from 22 down to 4.

A non-intrusive laser diagnostic known as molecular tagging velocimetry was used to find quantitative off-body velocity measurements in the wake of a sphere in the Mach 7 Ludwieg Tube Wind Tunnel located at The University of Texas at San Antonio. Acetone gas seeded in the flow was excited using the 4th harmonic of a pulse-burst Nd:YAG laser. The experimental results were used to validate and compare to simulations using a continuous Galerkin flow solver. Both the experimental and simulated results agreed on negative velocities in the viscous shear layer. Near the centerline of the sphere, average velocities of -200 m/s were observed due to reverse flow in the recirculation region. Outside of the shear layer, velocities of 800 m/s were observed in both the experimental and simulated results. The overall average uncertainty for the strut-mounted and free-flight case was estimated to be +/- 27 m/s. The flow features of the wake were visualized using schlieren imaging, and the experimental results agreed well with the simulated results.

An optically based experimental approach for estimating detonation cell size of premixed gas phase fuel–oxidizer mixtures in an optically accessible linear detonation tube is presented. Detonation wave fronts propagating through undiluted near-stoichiometric ethylene–oxygen mixtures in the circular detonation tube were visualized and recorded using CH* chemiluminescence imaging near 430 nm at 100 kHz for initial mixture pressures up to 22 kPa. The chemiluminescence imaging, coupled with high-speed videography, is shown to capture cellular detonation structures as small as 1.6 mm in width. The measured cell sizes increase as the initial fill pressure decreases, corroborating well-established relationships between detonation cell sizes and initial reactant pressures. The optically based method is validated against conventional soot foil measurements performed simultaneously with multiple detonations at various initial fill conditions. Both chemiluminescence images and soot foil measurements are compared to previously published cell size trends for undiluted fuel–oxygen detonations, demonstrating reasonable agreement with the established methods. Paired with the optically accessible detonation channel, the high-speed chemiluminescence imaging technique offers a passive estimation of detonation cell size for the range of conditions investigated with a faster experimental turnaround time relative to conventional methods.

High-speed global surface temperature distributions and heat flux measurements on the Initial Concept 3.X vehicle (IC3X) model were investigated at the UTSA Mach 7 wind tunnel, examining angles of attack of 0° and 5° at a freestream unit Reynolds number (Re) ~7 × 106 m−1. A ruthenium-based, fast-responding, temperature-sensitive paint (fast-TSP) prepared in-house was applied to a 7.1% scale model of the vehicle. Static calibration was performed to convert the intensity measurements into surface temperature values. The surface temperatures and derived heat flux fields conformed to the predicted trends, which was corroborated by Schlieren flow visualization. Notably, the average surface temperature variation was identified to range from 6 to 34 K at a 0° angle of attack and from 11 to 44 K at a 5° angle of attack, with the most pronounced gradient detected at the stagnation point. Additional measurements provided a detailed thermal assessment of the model, including estimations of the stagnation point heat flux, the convective heat transfer coefficient, and the modified Stanton number. Statistical and time series analyses of the data collected revealed the absence of prevailing unsteady phenomena, suggesting that the tested design geometry is well suited for hypersonic flight applications. These experimental outcomes not only shed light on the aerothermodynamics experienced during high-speed flight but also underscore the effectiveness of fast-TSP in capturing both quantitative and qualitative thermal data.

The lifetime and intensity techniques are two distinct methods used to measure the surface pressure of wind tunnel models with pressure-sensitive paint (PSP). When exposed to 405 nm light for extended periods, both the luminescence and the lifetime ratio were observed to decrease, affecting perceived pressure values. To help further the implementation of unsteady pressure-sensitive paint (uPSP) for wind tunnel tests, this paper describes the efforts to characterize how photodegradation affects the intensity of the paint signal and its lifetime ratio, as well as the intensity calibration's response and how these changes vary with different pressures. It was found that the paint photodegrades at different rates depending on the partial pressure of oxygen during excitation. It was also found that the calibration curves exhibit greater changes when exposed to excitation light in higher pressures, both indicating photodegradation due to photodecomposition and photo-oxidation processes. These effects are observed on models in a subsonic 14" x 14" wind tunnel as well as a large-scale 11-by 11-ft transonic wind tunnel. The work provides analysis into the photodegradation behavior of PSP and uPSP, helping to reduce the error in pressure measurements obtained through these techniques.

Hypersonic flight holds immense promise across commercial aviation, defense, and space exploration. Yet, many technical challenges hamper the widespread use of hypersonic vehicles. Of particular concern are the interactions between shock-waves and turbulent boundary layers. These interactions increase structural risk by inducing high temperatures, amplifying vibration loads, and increasing turbulent drag. To address this technical challenge, we investigate using phononic metamaterials (PMs) as a passive flow control mechanism. PMs are artificial periodic materials that exhibit frequency ranges, called bandgaps, at which wave propagation is highly attenuated. We aim to reduce the amplitude of turbulent fluctuations by designing PMs with bandgaps at frequency ranges where shock-wave/boundary-layer interactions (SWBLIs) are most active (i.e. ~𝟏 𝒌𝑯𝒛). We present the design, analysis, fabrication, and testing of a bilayer phononic metamaterial tailored for controlling hypersonic SWBLIs. Experimental results showcase the material's ability to reduce turbulent fluctuations, providing experimental evidence on applying phononic metamaterials for flow control in hypersonic flight.

Experiments with sonic and Mach 3 transverse jets in crossflow (JICF) were conducted in the UTSA Mach 7 wind tunnel. A test apparatus based on a flat plate with a sharp leading edge and interchangeable jet inserts is described. Schlieren imagery was collected at 20 kHz for a JPR of 870 to 400 for the sonic jet and 40 to 20 for the Mach 3 jet. Modal analysis of the schlieren imagery using spectral proper orthogonal decomposition (SPOD) allowed investigation of JICF dynamics such as bow shock flapping and separation shock movement. An image-based shock tracking algorithm was used to investigate the separation shock foot dynamics. Acetone planar laser-induced fluorescence was used to visualize the streamwise centerline JICF for the Mach 3 jet at acquisition rates of 10 kHz and 100 kHz.

Hypersonic flight holds immense promise across commercial aviation, defense, and space exploration. Yet, many technical challenges hamper the widespread use of hypersonic vehicles. Of particular concern are the interactions between shock-waves and turbulent boundary-layers. These interactions increase structural risk by inducing high temperatures, amplifying vibration loads, and increasing turbulent drag. To address this technical challenge, we investigate using phononic metamaterials (PMs) as a passive flow control mechanism. PMs are artificial periodic materials that exhibit frequency ranges, called bandgaps, at which wave propagation is highly attenuated. We aim to reduce the amplitude of turbulent fluctuations by designing PMs with bandgaps at frequency ranges where shock-wave/boundary-layer interactions (SWBLIs) are most active (i.e. similar to 1 kHz). We present the design, analysis, fabrication, and testing of a bilayer phononic metamaterial tailored for controlling hypersonic SWBLIs. Experimental results showcase the material's ability to reduce turbulent fluctuations, providing experimental evidence on applying phononic metamaterials for flow control in hypersonic flight.