This study investigates the Mach-6 compression ramp flows at a unit Reynolds number of 6.37×106 m−1 in a Ludwieg tube tunnel, with ramp angles ranging from 4 to 10 deg. Global stability analysis indicates that global instability is absent in all cases. Resolvent analysis identifies three convective instability modes: Mack’s first mode, second mode, and a streak mode, each exhibiting distinct features across different cases. A combination of schlieren visualization, Rayleigh scattering photography, surface pressure measurements, and infrared thermography is employed to reveal the convective instability characteristics and the associated laminar-to-turbulent boundary-layer transition mechanisms. It is found that the transition point significantly shifts upstream when flow separation occurs at the ramp corner for ramp angles exceeding 4 deg. The shear-layer mode and Mack’s second mode are measured by pressure sensors. The spatial evolution of detailed flow structures within the reattached boundary layer, such as staggered wavelike structures, Λ vortices, and streamwise streaks, is visualized. Analysis of the evolution and breakdown of these structures reveals that the nonlinear evolution of the first mode predominantly drives boundary-layer transition in the 4-deg-ramp case, referred to as the oblique transition. Görtler instability is significantly amplified near the separation and reattachment regions, with its strength increasing with the ramp angle. Consequently, it plays a more critical role in the transition process at higher ramp angles and dominates in the 10-deg-ramp case. For moderate ramp angles, both the first mode and Görtler instability are pronounced in the transition process.

Small unmanned air vehicles (SUAVs) frequently perform rapid tilting maneuvers to transition between flight conditions or avoid obstacles. This study investigates the aerodynamic effects of three specific transitions: hovering to forward flight, climbing to forward flight, and forward flight to hovering. The analysis reveals that rapid maneuvers cause aerodynamic loads such as thrust and torque to deviate significantly from steady-state values while also inducing marked oscillations in the loads and substantial wake deformation. Due to its nonnegligible inertia, the wake cannot instantaneously adapt to the propeller’s position and orientation throughout the maneuver. This lag results in wake stretching or compression along the axial direction, which modifies the induced velocity on the blades and affects the resulting aerodynamic forces. These effects become more pronounced as the maneuver’s rapidity increases. The analyses are carried out using a midfidelity vortex particle method (VPM) solver applied to a low-Reynolds-number SUAV propeller. This meshless, Lagrangian approach efficiently captures complex vortical structures and wake interactions at low computational cost. The propeller blades are modeled using a lifting-line approach integrated into the VPM solver, making it an efficient tool for investigating dynamic propeller effects in SUAV performance evaluation and design.

A recent tractable gradient-based approach that exploits a new two-point topological derivative is extended here to encompass multifidelity modeling. Handling multidisciplinary responses during structural topology optimization is important to a field that has grown rapidly in the last decade, spurred by additive manufacturing. The cellular division method for topological design is more general than conventional optimality criterion methods used most often in industrial structural topology design; however, it is a discrete approach that is daunting without topological derivatives, due to the curse of dimensionality. Results are presented to address that issue for benchmark test cases using the multifidelity approach with two-point topological derivatives to drive cellular development.

Airborne towed-antenna systems have significant application potential in scenarios such as very-low-frequency communications, owing to their flexible deployment and high survivability. The system considered herein consists of an unmanned aerial vehicle (UAV) and a flexible towed antenna and exhibits pronounced nonlinearity and strong bidirectional coupling. To address the difficulty of providing a unified description of the two-way interactions between the UAV and the towed body, a coupling-oriented dynamic modeling approach for a UAV sling-loaded towed-antenna system is proposed. First, the towed antenna is discretized into multiple elements, and its flexible dynamics are formulated using the absolute nodal coordinate formulation. Subsequently, a six-degree-of-freedom rigid-body dynamic model of a quadrotor UAV is established. A bidirectional coupling mechanism incorporating kinematic constraints and force feedback is then constructed: the UAV center of mass is equivalently mapped to the top-end node of the antenna to transmit the UAV pose to the flexible body, while the equivalent towing load acting on the UAV is obtained by synthesizing the nodal external forces and inertial terms of the antenna. In this manner, a closed-loop force/pose coupling model at the UAV/antenna connection point is achieved. Finally, based on the developed coupled model, a model predictive control trajectory-tracking controller is designed, and comparative simulations against a proportional/integral/derivative controller are conducted.

Vortical structures and the associated flowfield on finite-span, cantilevered NACA 0015 wings were observed experimentally. The effects of semi-aspect ratios of 1, 2, and 4, sweep angles of 0, 15, 30, and 45 deg, at angles of attack of 12, 18, and 22 deg were explored at a Reynolds number of 600. On each wing, there are three regions along the span: a wall region, a tip region, and wake region between the two. At low semi-aspect ratios, end effects played a significant role by mitigating the shedding, resulting in a small separated region. At higher semi-aspect ratios, the wings ends had less of an influence on the overall flowfield, allowing the developments of separation and shedding in the wake. For the unswept wings, secondary structures of alternating rotation developed as a result of the interaction between the separated flow and the end regions. In the swept cases, an outboard spanwise velocity component developed, which interacted with the separated region forming secondary structures and a large dominant swirl vortex along the span. Both time-averaged and instantaneous flowfields were used to understand the formation of the structures. The present study shows the complexity of three-dimensional flow separation even on simple geometries.

Multifidelity uncertainty propagation combines the efficiency of low-fidelity models with the accuracy of a high-fidelity model to construct statistical estimators of quantities of interest. It is well known that the effectiveness of such methods depends crucially on the relative correlations and computational costs of the available computational models. However, the question of how to automatically tune low-fidelity models to maximize performance remains an open area of research. This work investigates automated model tuning, which optimizes model hyperparameters to minimize estimator variance within a target computational budget. Focusing on multifidelity trajectory simulation estimators, the cost-versus-precision tradeoff enabled by this approach is demonstrated in a practical, online setting where upfront tuning costs cannot be amortized. Using a real-world entry, descent, and landing example, it is shown that automated model tuning largely outperforms hand-tuned models even when the overall computational budget is relatively low. Furthermore, for scenarios where the computational budget is large, model tuning solutions can approach the best-case multifidelity estimator performance where optimal model hyperparameters are known a priori. Recommendations for applying model tuning in practice are provided, and avenues for enabling adoption of such approaches for budget-constrained problems are highlighted.

Conical–cylindrical assembled shells are typical structural designs of launch vehicles. In addition to thermal stresses from the high-temperature environment, these structures are subjected to external aerodynamic loads during service, leading to a nonlinear stress state in the shells. However, current studies on the vibration behaviors of shells under coupled thermal–aerodynamic loads primarily rely on the linear solution framework, and they are predominantly limited to single-shell structures. To tackle the limitation, a novel theoretical scheme for the vibration of conical–cylindrical assembled shells under thermal–aerodynamic loads, considering geometric nonlinearity, is developed for this study. The original nonlinear governing equation is transformed by applying the perturbation method and the quasi-linearization method, and the obtained nonhomogeneous and homogeneous equations are solved with the precise integration method with and without dimensional expanding, respectively. Additionally, the flutter, thermal buckling, and free vibration analyses are also achieved by adjusting the involved aerodynamic, thermal, and vibration terms in the present scheme. Through extensive solution comparisons across different design parameters involving material properties, geometric parameters, and boundary and loading conditions, the developed scheme demonstrates excellent accuracy. Furthermore, the effects of key parameters on the vibration behaviors are quantitatively investigated. The summarized findings may help facilitate the structural designs of conical–cylindrical assembled shells.

Experiments were conducted in the University of Tennessee Space Institute’s Mach 7 Ludwieg tube, examining the impact of laser–surface interactions on the transitioning boundary layer on a 91-6 flared cone with Re∞= 8.0 million per meter. Laser–surface interactions were created using a pulse burst laser system operating with a 1064 nm output at repetition rates of 50, 100, 200, and 400 kHz. Global flowfield effects were visualized using high-speed schlieren imaging, and high-response surface pressure sensors were used to monitor changes in boundary-layer spectral content. Results show that for both weak and strong laser–surface interactions (as defined in this work), a significant reduction in boundary-layer fluctuations at higher frequencies occurs, which is also accompanied by a shift to larger turbulent structures and a thickening of the boundary layer. In strong interactions, it is also shown that significant separation can occur and that wraparound effects can even cause boundary-layer thickening and transition impacts on the nonincident side of the model.

Flow-induced vibration is a major contributor to structural damage in aircraft. This work aims to mitigate such vibrations using reinforced structures. Unlike conventional reinforcement approaches, the structures developed here are engineered to align the aerodynamic mode with the structural mode in the spatial/wave-number domain, thereby minimizing flow-induced vibration. To achieve this, a multi-iteration design method is first introduced to reduce structural vibration by tuning the structural natural frequencies through adjusting stiffener dimensions so that the structural mode approaches the minimum-vibration condition under a given wall-pressure excitation. However, the optimal reinforcement scheme is difficult to obtain within a limited number of iterations. To further enhance the efficiency of reinforcement design, an improved method is developed that accelerates the design process by directly determining reinforcement dimensions using generated velocity contours for certain wall-pressure fluctuation modes. Evaluation of the vibration-suppression performance indicates that the newly developed methods significantly reduce flow-induced vibration and effectively control structural stress at the natural frequency as compared with conventional reinforcement strategies.

Periodic unsteady flow problems often exhibit computational inefficiency when conventional time-domain methods are employed. This phenomenon can be attributed to the lengthy transition phase before reaching the ultimate periodic state. To address this problem, the harmonic balance method (HBM) is employed for aeroacoustic applications through the implementation of convective monopole sources, frequency-domain gust excitation, and nonreflective boundary conditions. A set of test cases is considered to evaluate this method, including monopole sources with uniform background flows, the pitching oscillations of airfoils, leading-edge noise of gust–airfoil interactions, and the response of gust–cascade interactions. The results show good agreement with analytical and experimental data, with normalized root-mean-square errors typically below 10% across most considered cases. The findings from this research demonstrate a comprehensive range of potential applications for HBM in the fields of fluid dynamics and aeroacoustics.