Search Results (52)

In our previous study, the transition behavior of a flared cone–swept fin configuration was investigated under an angle of attack (AoA) of 0°. To further explore the role of AoA in complex three-dimensional geometries with strong fin–body interactions, wind tunnel experiments were conducted at Ma = 9.3, Re = 1.36 × 10⁷/m, with AoA ranging from −6° to 6°. Global surface temperature distributions were obtained using temperature-sensitive paint (TSP), while localized heat flux and pressure fluctuations were captured using thin-film thermocouples and high-frequency pressure sensors. The results show that varying AoA shifts the location of high heat flux between the upper and lower surfaces of the flared cone and induces a switch from streamwise to separation vortices. The windward side exhibits stronger disturbance responses than the leeward side. The junction region between the flared cone and the near-horizontal surface is highly sensitive to AoA variations, consistently exhibiting pronounced second-mode instabilities. These findings provide experimental support for understanding transition mechanisms under the combined effects of shock/boundary layer interaction (SBLI), crossflow, and adverse pressure gradients, with implications for transition prediction and thermal protection system design. Full article

Aerodynamic ablation of external control surfaces and structural complexity in hypersonic reentry vehicles (HRVs) pose significant challenges for maneuverability and system reliability. To address these issues, this study develops a novel bank-to-turn (BTT) control strategy integrating a single internal moving mass with differential ailerons, eliminating reliance on ablation-prone elevators/rudders while enhancing internal space utilization. A coupled 7-DOF dynamics model explicitly quantifies inertial-rolling interactions induced by the moving mass, revealing critical stability boundaries for roll maneuvers. To ensure robustness against aerodynamic uncertainties, aileron failures, and high-frequency mass-induced disturbances, a dynamic inversion controller is augmented with an L1 adaptive layer decoupling estimation from control for improved disturbance rejection. Monte Carlo simulations demonstrate: (1) a 20.6% reduction in roll-tracking error (L₂-norm) under combined uncertainties compared to dynamic inversion control, and (2) a 72% suppression of oscillations under aerodynamic variations. Comparative analyses confirm superior transient performance and robustness in worst-case scenarios. This work offers a practical solution for high-maneuverability hypersonic vehicles, with potential applications in reentry vehicle design and multi-actuator system optimization. Full article

Computational fluid dynamics (CFD) has aided the development, design, and analysis of hypersonic airbreathing propulsion technologies, such as scramjets. The complex flow field in a scramjet isolator has been the subject of intense interest and study for several decades. Many features of this flow field also occur in supersonic wind-tunnel nozzles and diffusers. Computational analysis of these topics has frequently provided immense insight into the actual functionality and performance. Research presented in this work supports scientific investigation and understanding of a less-researched topic, which is shock–shock interference and interaction with the boundary layer in supersonic internal flows, as well as the passive control of its adverse effects to prevent the onset of unstart in a scramjet isolator. This computational investigation is conducted on a backpressured isolator and a modified three-dimensional shock-tube to represent a scramjet isolator with ram effects provided by high-pressure gas and high-speed flow provided by a supersonic inflow. Computational results for the backpressured isolator have been validated against available measured time-averaged wall pressure data. The modified shock-tube provided an opportunity to study the shock–shock interference and shock–boundary-layer interaction effects that would occur in a scramjet isolator or a ram-accelerator when the high-speed flow from the inlet interacted with the shock produced due to the combustor pressure traveling and meeting in the isolator. An assessment of wall cooling effects on these phenomena is presented for both the backpressured isolator and the modified shock-tube. Full article

This study presents a design approach for a uniform porous surface to control laminar-to-turbulent transition in hypersonic boundary layers. The focus is on suppressing the Mack second mode, which is a dominant instability in hypersonic boundary layers. The Mack second mode is acoustic-wave-like in the ultrasonic frequency range and can be effectively attenuated by porous surfaces. Previous studies have explored porous surfaces, either by targeting a specific frequency or by adopting geometrically complex configurations for various frequencies. In contrast, the present study proposes a porous surface design that effectively stabilizes the Mack second mode over a wide frequency range, while maintaining structural simplicity. In addition, this porous surface design incorporates constraints associated with practical fabrication to enhance manufacturability. The absorption characteristics of porous surfaces are evaluated with an acoustic impedance model, and the stabilization performance is assessed with linear stability theory. The proposed porous surface design is compared with a conventional design method that focuses on the Mack second mode with a single frequency. Consequently, the proposed design methodology demonstrates robust and consistent suppression of the Mack second mode in a broad frequency range. This approach improves both stabilization performance and manufacturability with a uniform porous surface, contributing to its practical application in high-speed vehicles. Full article

Computational fluid dynamics (CFD) simulations play a strong role in the design and development of aerospace and defense vehicles, including high-speed applications where testing under the correct operational conditions is not yet viable. In this study, metrics for the onset of isolator unstart are identified. An assessment of the variance of operating variables and their impact on metrics for the onset of isolator unstart and operability metrics was performed, utilizing a nozzle–isolator assembly from NASA Langley Research Center as a demonstration case. The effects of increasing backpressure ratio and decreasing inflow Mach number on these metrics and the underlying contributions of shock physics were investigated in detail. A major conclusion from this study is that both inflow Mach number and backpressure ratio can strongly impact pseudo shock train and shock–boundary layer interactions inside the isolator, but inflow Mach number has a stronger impact than the backpressure ratio. The research presented in this paper demonstrates that the isolator performance can shift from start to unstarted and operable to inoperable with a small variance in operating conditions. Another important insight presented in this research is that the length of the pseudo shock train and the Mach stem height change discontinuously with both the backpressure ratio and the inflow Mach number. Therefore, the length of the pseudo shock train and height of the Mach stem are strong indicators of the onset of unstart, which is an important consequence for instrumentation and closed-loop adaptive feedback control system design for scramjet flight operations. Full article

Powder bed fusion using a laser beam (PBF-LB) offers a suitable alternative to manufacturing Ti65 with intricate geometries and internal structures in hypersonic aerospace applications. However, issues such as undesirable surface roughness, defect formation, and microstructural inhomogeneity remain critical barriers to its wide application. In this study, a coupled discrete element method–computational fluid dynamics (DEM-CFD) model was utilized to investigate the spreading behavior of Ti65 powder in a multi-layer PBF-LB process. The macro- and microscopic characteristics of the powder beds were systematically analyzed across different layers and regions under various spreading velocities. The results show that the packing density and uniformity of the powder beds in multi-layer PBF-LB of Ti65 powder improves as the number of solidified layers increases. Poor bed quality is observed in the first two layers due to a strong boundary effect, while a stable and denser powder bed emerges from the fourth layer. The presence of a previously solidified region strongly influences its neighboring unsolidified areas, enhancing density in the upstream region and causing looser packing downstream. Additionally, due to the existence of a solidified region, the height of the powder bed progressively decreases along the spreading direction. Full article

The shock wave/boundary layer interaction phenomenon in hypersonic inlets, affected by background waves, may induce the formation of multiple separation zones. Existing theories prove insufficient in explaining the underlying flow mechanisms behind complex phenomena arising from multi-separation zone interactions, which necessitates further investigation. To clarify the governing factors in multi-separation zone interactions, this study developed a simplified dual-separation-zone model derived from inlet flow field characteristics. A series of numerical simulations were conducted under an incoming flow at Mach 3 to systematically analyze the effects of internal contraction ratio, the influencing locations of expansion waves, and incident shock wave intensity on the mergence and re-separation of dual separation zones. The results demonstrate that both the expansion wave impingement position and incident shock intensity significantly influence specific transition points in dual-separation-zone flow states, though they do not fundamentally alter the evolutionary patterns governing the merging/re-separating processes. Furthermore, increasing incident shock intensity leads to the expansion of separation zone scales and prolongation of the dual-separation-zone merging distance. Full article

Large-eddy and direct numerical simulations generate vast data sets that are challenging to interpret, even for simple geometries at low Reynolds numbers. This has increased the importance of automatic methods for extracting significant features to understand physical phenomena. Traditional techniques like the proper orthogonal decomposition (POD) have been widely used for this purpose. However, recent advancements in computational power have allowed for the development of data-driven modal reduction approaches. This paper discusses four applications of deep neural networks for aerodynamic applications, including a convolutional neural network autoencoder, to analyze unsteady flow fields around a circular cylinder at Re = 100 and a supersonic boundary layer with Tollmien–Schlichting waves. The autoencoder results are comparable to those obtained with POD and spectral POD. Additionally, it is demonstrated that the autoencoder can compress steady hypersonic boundary-layer profiles into a low-dimensional vector space that is spanned by the pressure gradient and wall-temperature ratio. This paper also proposes a convolutional neural network model to estimate velocity and temperature profiles across different hypersonic flow conditions. Full article

Recent trends in aeroelastic analysis have shown a great interest in understanding the role of shock boundary layer interaction in predicting the dynamic instability of aircraft structural components at supersonic and hypersonic flows. The analysis of such complex dynamics requires a time-accurate fluid-structure interaction solver. This study focuses on the development of such a solver by coupling a finite-volume Navier-Stokes solver for fluid flow with a finite-element solver for structural dynamics. The coupled solver is then verified for the prediction of several panel instability cases in 2D and 3D uniform flows and in the presence of an impinging shock for a range of subsonic and supersonic Mach numbers, dynamic pressures, and shock strengths. The panel deflections and limit cycle oscillation amplitudes, frequencies, and bifurcation point predictions were compared within $10\%$ of the benchmark results; thus, the solver was deemed verified. Future studies will focus on extending the solver to 3D turbulent flows and applying the solver to study the effect of turbulent load fluctuations and shock boundary layer interactions on the fluid-structure coupling and structural dynamics of 2D panels. Full article

During the operation of hypersonic vehicles, a reciprocal coupling effect is manifested between the inlet and the combustion chamber. This results in an unavoidable non-uniformity of conditions at the combustion chamber’s entrance, which, in turn, influences the fuel mixing within the chamber. The present study employed the Reynolds-averaged Navier–Stokes (RANS) equations to perform a numerical simulation of an X-51-like vehicle, with a focus on examining the impact of isolation section length and multi-injection strategies on the fuel mixing characteristics within the combustion chamber under conditions of non-uniform inflow. The findings indicated that a supersonic non-uniform inlet triggers incident shock waves, leading to a non-uniform pressure distribution across the flow section. Moreover, the position of injection was found to be pivotal in regulating penetration depth and mixing efficiency. The incident shock wave, bow shock, and boundary layer separation shock interacted with each other to increase local pressure. The coupling of high and low pressures generated an adverse pressure gradient that led to boundary layer separation, which further enhanced fuel penetration depth. Full article

This review examines studies aimed at the organization of energy (non-mechanical) control of high-speed flow/flight using spatially multi-component plasma structures and combined energy deposition. The review covers selected works on the experimental acquisition and numerical modeling of multi-component plasma structures and the use of sets of actuators based on plasma of such a spatial type for the purposes of control of shock wave/bow shock wave–energy source interaction, as well as control of shock wave–boundary layer interaction. A series of works on repetitive multiple laser pulse plasma structures is also analyzed from the point of view of examining shock wave/bow shock wave–boundary layer interaction. Self-sustained theoretical models for laser dual-pulse, multi-mode laser pulses, and self-sustained glow discharge are also considered. Separate sections are devoted to high-speed flow control using combined physical phenomena and numerical prediction of flow control possibilities using thermal longitudinally layered plasma structures. The wide possibilities for organization and applying spatially multi-component structured plasma for the purposes of high-speed flow control are demonstrated. Full article

The issue of hypersonic boundary layer transition prediction is a critical aerodynamic concern that must be addressed during the aerodynamic design process of high-speed vehicles. In this context, we propose an advanced mesoscopic method that couples the gas kinetic scheme (GKS) with the Langtry–Menter transition model, including its three high-speed modification methods, tailored for accurate predictions of high-speed transition flows. The new method incorporates the turbulent kinetic energy term into the Maxwellian velocity distribution function, and it couples the effects of high-speed modifications on turbulent kinetic energy within the computational framework of the GKS solver. This integration elevates both the transition model and its high-speed enhancements to the mesoscopic level, enhancing the method’s predictive capability. The GKS-coupled mesoscopic method is validated through a series of test cases, including supersonic flat plate simulation, multiple hypersonic cone cases, the Hypersonic International Flight Research Experimentation (HIFiRE)-1 flight test, and the HIFiRE-5 case. The computational results obtained from these cases exhibit favorable agreement with experimental data. In comparison with the conventional Godunov method, the new approach encompasses a broader range of physical mechanisms, yielding computational results that closely align with the true physical phenomena and marking a notable elevation in computational fidelity and accuracy. This innovative method potentially satisfies the compelling demand for developing a precise and rapid method for predicting hypersonic boundary layer transition, which can be readily used in engineering applications. Full article

A two-point cylindrical-focused laser differential interferometer (2P-CFLDI) system and a conventional Z-type Schlieren were used to measure the hypersonic turbulent boundary layer on a flat plate at Mach number Ma = 6 and Reynolds number Re = 1.08 × 10⁶ m⁻¹. The boundary layer thickness at the measurement location and the noise radiation angle were obtained by post-processing the Schlieren image. The 2P-CFLDI data underwent cross-correlation analysis to calculate the mean convective velocities at different heights and compared with previous experimental and numerical results. The experimentally measured mean convective velocities agree with the trend of available DNS and experimental results. The mean convective velocity near the wall is significantly larger than the local mean velocity and is the main noise source region. Further filtering treatment shows that the convective velocity of the disturbed structure decreases gradually with the increase in the disturbance scale. The differences between convective velocities at different scales are significantly larger outside the boundary layer than inside the boundary layer, which is in agreement with the findings of the previous hot wire experiments. Near the wall, large-scale disturbances mainly determine the localized mean convective velocity, which are the main source of noise radiation for the hypersonic turbulent boundary layer. Full article

The double-wedge configuration is a typical characteristic shape of the rudder surface of high-speed aircraft. The impact of the shock wave/boundary layer interaction and the shock wave/shock wave interaction resulting from the double wedge on aircraft aerodynamics cannot be ignored. The aerodynamic performance of the aircraft would be seriously affected. Accordingly, to reduce the wave drag, and to relieve the thermal load and pressure load, flow control is required for the shock wave/shock wave interaction and the shock wave/boundary layer interaction induced by the double-wedge configuration. In this paper, double-wedge shock wave/shock wave interaction is controlled by a high-energy surface arc discharge array and observed by high-speed schlieren flow field measurement at Mach 8. The 30-channel discharge array is set on the primary wedge plane, and actuation is generated. Hypersonic V shock wave/shock wave interaction is effectively controlled by the shock wave array induced by the high-energy surface arc discharge array, which makes the shock wave/shock wave interaction structure disappear or intermittent. The potential control mechanism is to reduce strong shock wave interaction by transforming the type of shock wave interaction. Therefore, the ability of plasma array actuation to control complex shock wave/shock wave interaction is verified, which provides a new method for hypersonic shock wave/shock wave interaction control. Full article

Temperatures within the boundary layers of high-enthalpy hypersonic flows can soar to thousands or even tens of thousands of degrees, leading to significant real gas phenomena. Although there has been significant research on real gas effects on hypersonic boundary layer stability, their impact on the boundary layer’s receptive stage is still poorly understood. Most aerodynamic boundary layers in flight vehicles are three-dimensional. Because of complex geometry and significant crossflow effects, the crossflow mode in three-dimensional boundary layers is crucial in hypersonic vehicle design. In this study, a linear stability analysis (LST) accounting for chemical nonequilibrium effects (CNE) and its adjoint form (ALST) is developed to investigate the real gas effects on the stability and receptivity of stationary crossflow modes. The results indicate that real gas effects significantly influence the receptivity of stationary crossflow modes. Specifically, chemical nonequilibrium effects destabilize the crossflow modes but reduce the receptivity coefficients of the stationary crossflow modes. The Mach number effect was also investigated. It was found that increasing the Mach number stabilizes the stationary crossflow modes, but the receptivity coefficients increase. As the Mach number progressively rises, these effects alternately dominate, leading to a non-monotonic shift in the transition position. Full article