This work presents a methodology for designing waverider geometries that incorporate leading edge rounding and parameterized surface perturbations, and evaluates their performance using two distinct aerodynamic prediction approaches: a reduced-order analytical/empirical model and a GPU-based Euler CFD solver, both applied with an empirical approach to estimate the viscous contributions. Geometry optimization was performed using a global pattern search algorithm, and optimal trajectories were computed using a direct collocation method to assess the impact of aerodynamic modeling fidelity on mission performance. Cross-comparisons between these aerodynamic performance prediction methods revealed that CFD-based optimization produces geometries that capture off-design flow phenomena that are absent from analytical-based optimizations. As a result, CFD-optimized waveriders demonstrated higher average lift-to-drag ratios than those derived from the analytical model. The analytical trajectories had lower angle-of-attack control and degraded in accuracy for vehicles with high pressure losses. These findings highlight the limitations of traditional analytical methods and underscore the feasibility of using modern GPU-accelerated CFD for resource-intensive tasks such as vehicle optimization and trajectory analysis.

Laminar-to-turbulent transition delay is a key challenge in hypersonic boundary-layer flows. Unstable disturbances-most prominently the first and second modes-trigger the onset of turbulence and pose a fundamental technological barrier to hypersonic transport. While existing control strategies target the second mode, simultaneous mitigation of the first mode has long appeared physically impossible. A new flow-control concept is introduced in which phase relations between wall pressure and velocity fluctuations are tailored using subsurface phonon engineering to control both modes concurrently. The outcome is substantial drag reduction and alleviation of the extreme thermal loads associated with turbulence.

Surface roughness significantly impacts transition to turbulence, especially over high-speed, blunt geometries where surface ablation is necessary to mitigate heat loads during atmospheric entry. Inspired by sand-grain roughness experiments performed by Hollis (2017), we perform the first direct numerical simulation (DNS) of a blunt cylinder in Mach 6 cross-flow with roughness elements distributed along the entire surface. Such simulations aimed to uncover the precise means by which laminar-turbulent transition occurs given the limited measurements attainable from experiments and non-existent high-fidelity simulations. Element heights were held fixed at approximately 35% boundary layer thickness, while the relative phasing between streamwise rows was varied. All configurations exhibited convective instabilities driving the transition process, with the mode type being set by the roughness configuration. A fundamental sinuous streak mode dominated the aligned roughness element case, whereas both the staggered and randomly phased cases saw 2D T-S waves dominating. These instability waves, when grown to sufficient amplitude, triggered the steady streaks seeded by the underlying roughness pattern to begin forming hairpin vortices and breakdown occurred soon thereafter. The roughness arrangement was found to dramatically influence the degree to which the waves were destabilised, as well as the strength of the underlying steady streaks, thereby combining to dictate the position along the surface where LTT occurred. Finally, exogeneous forcing was not required for the T-S dominated cases as the acoustics generated by the turbulence in the subsonic flow excited T-S waves on the other side - a feedback mechanism hitherto unknown.

In the present paper, several boundary-layer stability codes are compared, based on hypersonic, high-enthalpy boundary-layer flows around two different blunted cones: a 7 deg and a 10.9 deg half-angle blunted cone. One part of the code-to-code comparison, which is conducted between seven different codes, assesses the influence of different numerical approaches for the second-mode instability. The focus of the paper characterizes the role of the high-temperature gas effects on the second mode. The experimental test cases for the comparison are provided by two high-enthalpy wind-tunnel facilities: the High Enthalpy Shock Tunnel (HIEST) of Japan Aerospace Exploration Agency (JAXA) and the High-Enthalpy Laboratory Munich (HELM) of the University of Bundeswehr Munich. Analyses between the various computational stability codes and experimental measurements are performed, quantifying the high-temperature gas effects and the influence of uncertainties (e.g., small mean-flow or wall-temperature variations). A comparison between air and N2 test gases is also performed.

This study introduces a computational framework for predicting structural deformation caused by high-speed water droplet impacts while accounting for the effects of droplet aerobreakup within the shock layer. This investigation is significant because no prior computational studies have specifically addressed the damage resulting from a deformed raindrop impact at extreme hyper-velocities. The framework employs a sequential, one-way coupling of a higher-order computational fluid dynamics (CFD) multiphase solver and a peridynamics (PD)-based structural mechanics approach. CFD simulations resolve the droplet-shock interaction and provide aerodynamically deformed droplet geometries, which are subsequently used in PD simulations of the impact event. Simulations are conducted over a wide range of Mach numbers (1.5-8), and predicted crater depths are compared with available experimental data at lower speeds. The results demonstrate that the droplet breakup history imposes a first-order effect on both surface deformation and material damage: slightly deformed droplets produce crater depths comparable to spherical droplets at lower velocities, while moderately deformed droplets produce the greatest crater depths, highly deformed droplets yield the largest crater diameters, and all damage metrics decrease significantly for severely deformed droplets. Variations in peak force, impulse, and impact duration further indicate that droplet morphology governs the temporal distribution of loading during impact. These results demonstrate that accounting for droplet aerobreakup is essential for accurately predicting high-speed liquid impact damage.

The intense thermal loads experienced by hypersonic vehicles operating within the atmosphere can be mitigated using transpiration cooling, a technique that injects cold fluids along the vehicle's surface. However, this process can significantly impact flow stability, potentially triggering an early transition to turbulence. The temperature of the injected fluid, as it mixes with the external airflow, is highly dependent on the injection mechanisms and flow conditions. This study focuses on the destabilizing effects of gas injection for various injection gas temperatures. It is well-established that changes in wall temperature can have either stabilizing or destabilizing effects on flow instabilities, depending on the dominant instability mechanisms, such as first or second Mack modes. To evaluate the influence of injection gas temperature on the stability of a high-enthalpy boundary layer flow, air and CO2 injections are considered. The analysis considers different blowing rates and injection gas temperatures between 300 K and 1,000 K. The paper includes mappings of heat flux versus N-factors for different injection scenarios, and energy budget analysis is employed to elucidate the differences in the stability characteristics of the flow.

The accurate modeling of hypersonic environments, including coupled ablation, is a challenging problem due to the complex flow physics, numerical accuracy, and robustness required for these simulations. The simulation mesh needs to be designed carefully, and structured high-aspect ratio stretched grids are typically used to properly capture the high wall-normal gradients. The mesh generation process proves to be a cumbersome and time-consuming process for realistic and complex vehicles, creating a severe bottleneck to the current CFD workflow. This work highlights the development of the Cartesian Higher-Order Adaptive Multi-Physics Solver (CHAMPS) near-body Cartesian grid solver, capable of automatic volume mesh generation, and its accurate and robust use in simulating coupled ablation for steady-state graphite ablation by interfacing with the Kentucky Aerothermodynamics and Thermal Response System Material Response (KATS-MR) solver. Validation on a Mach 8.7 cylinder and a Mach 22.88 axisymmetric Mars 2020 capsule showcases the challenges associated with shock-capturing on non-shock-aligned Cartesian grids. A verification study of the CHAMPS–KATS coupled ablation framework is performed for steady-state graphite ablation on a Mach 25.2, 9 deg blunt cone, followed by the validation of the solver for a blunt cone in the NASA Ames Interaction Heating Facility arcjet.

Hypersonic vehicles operating within the atmosphere are subjected to extreme thermal loads, which can be mitigated using transpiration cooling - a technique that injects coolant gas or liquid through the vehicle surface. While this heat mitigation strategy can be effective in reducing surface temperatures, it can also alter boundary layer stability and potentially trigger an earlier transition to turbulence. Since the surface temperature and blowing characteristics of a transpiring coolant depend on both the external flow field and the properties of the porous material, this study investigates, for the first time, the effects of gas injection through a realistic porous structure on high-speed boundary layer stability. The gas transport within the material and the thermal response are resolved using a high-fidelity coupled CHAMPS-KATS fluid-material interaction solver. The resulting surface blowing velocity and temperature distributions are then used as boundary conditions for a separate linear stability analysis. Air injection is evaluated at multiple mass flux levels under both ground-based wind tunnel and stratospheric flight conditions over a moderately blunt cone geometry. LST and time-spectral analysis shows a strong effect on second-mode growth with a stabilization in the injection region and destabilization downstream of the injection region.

Wall-modeled large-eddy simulations (WMLES) of the Aerospatiale A-airfoil flow at Reynolds number of 1.0x107, angle of attack of 13.3 deg, and Mach number of 0.15 are used as a test bed for exploring how to handle the thin laminar boundary layer and how it transitions to turbulence in a WMLES context. It is shown that application of the wall model everywhere with an unsuitable exchange location can lead to excessive boundary-layer growth and up to 10% underprediction of the lift coefficient. Several numerical approaches are examined to handle the laminar region of the airfoil flow within the WMLES framework. A transition sensor is implemented to locally detect where the boundary layer is laminar or turbulent and then used to dynamically activating or deactivating the wall model. The sensitivity of the predictions to the mesh resolution and the exchange location is explored.

The effects of grid-boundary nonalignment within an immersed-boundary framework are investigated in the context of wall-modeled turbulent flows. In a previous study by Ganju et al. [1], it was shown that ghost cell reconstruction is essential for accurately simulating boundary layer flows using finite-difference-based immersed boundary methods (IBM). The study hypothesized that inaccuracies in the IBM wall-modeled large eddy simulation approach are primarily driven by the convective terms. The current paper is a continuation of this work. To isolate and quantify these effects, a set of flat plate boundary layer test cases were run over a range of inclination angles using a wall-modeled RANS-based IBM approach with a prescribed eddy-viscosity. Significant variations in the skin friction and inner-scale velocity profiles were observed, with a high degree of sensitivity to the immersed boundary treatment. The non-conservative nature of the boundary treatment was found to be the dominant error mechanism, and this effect was found to be amplified by error-mixing which arises from the use of differentiation operators not aligned with the boundary in highly anisotropic regions. A hybrid conservative boundary treatment was subsequently developed and tested, and found to essentially remove these nonalignment effects.