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We report a kind of coiled-up meta-liners with nonuniform cross sections (CMNC), which can efficiently attenuate low-frequency sound waves under grazing flow with a deep subwavelength thickness (e.g., ∼λ/17 at 500 Hz). At a grazing flow Mach number of 0.26, the average transmission loss of the meta-liner is 12.6 dB at 500–1000 Hz, which is twice as much as that of a double-degree-of-freedom acoustic liner of the same size. Physically, the nonuniform cross-sectional distribution and significant cross-sectional area ratio enhances vortex shedding, thus resulting in severe acoustic energy dissipation. The excellent low-frequency acoustic attenuation performance of CMNC is investigated thoroughly with experimental, theoretical, and numerical methods. This work provides an avenue for low-frequency noise reduction in grazing flow scenarios (e.g., in a high bypass ratio turbofan engine). [ABSTRACT FROM AUTHOR]

We are concerned with the long time existence of smooth solutions to full compressible Navier-Stokes equations in 3D bounded domains with Dirichlet boundary conditions on the velocity and temperature. The smallness restriction on the initial data or time interval is unnecessary. Under the condition that the limiting system admits a reasonably smooth solution on a given time interval, we verify that the compressible system admits the smooth solution on the same interval as well, provided that the Mach number is sufficiently small. Moreover, as the Mach number goes to zero, the solution of the compressible system converges uniformly to that of the incompressible one. We utilize some geometric techniques to overcome difficulties introduced by non-slip boundary conditions. [ABSTRACT FROM AUTHOR]

Efficient combustion schemes are increasingly vital as scramjet expansion into higher Mach numbers emerges as a leading trend. Thus, in the current work, an LES investigation on mixing improvement caused by a two-strut injector has been carried out for a Mach 2.5 model H 2 fueled scramjet combustor. The research focuses on understanding the flow field, flame lift-off characteristics and combustion stabilization in the two-strut combustor. The simulation results indicate that the two-strut configuration exhibits a broader pattern of temperature fluctuation, implying a more efficient spread of combustion downstream compared to the single-strut configuration. Analysis of Fast Fourier Transform (FFT) pressure data reveals inherent instability in the combustion processes for two-strut under supersonic air inflow conditions. The findings also demonstrate the significant improvement in mixing performance achieved by the two-strut design compared to conventional struts. The primary mechanism is a two-strut's ability to create significant streamwise vorticity, which can improve the mixing of H 2 fuel with supersonic air. According to the mixing and total pressure loss plots, a two-strut produces a rapid complete mixing at 0.182 m at the expense of a minute rise in total pressure loss. Finally, the entropy change plot reveals that the two-strut configuration exhibits a reduced entropy change compared to the single-strut configuration. This is mainly due to enhanced air-fuel mixing, efficient shock wave management, and enhanced flame stabilization in the two-strut setup. These factors collectively contribute to reduced irreversibilities and disorder in the flow, leading to lower entropy within the system. [Display omitted] • LES investigation has been carried out for a Mach 2.5 model H 2 fueled scramjet combustor. • For two-strut, the shock train is unstable and affected by the recirculation zone. • Two-strut arrangement illustrates a more extensive pattern of temperature variation. • The two-strut-induced mixing improvement is more apparent. • The two-strut configuration shows lower entropy change compared to single-strut. [ABSTRACT FROM AUTHOR]

Theoretical studies indicate that electric potential saturation in a polar cap ionosphere is regulated by the limited reconnection rate at the magnetopause associated with field‐aligned currents (FACs), predicting the possible saturation of FACs under large interplanetary electric field (IEF). However, recent statistical studies have shown that observed FACs increase linearly with IEF. To reconcile this disparity, numerical experiments are conducted to explore the response of FACs under large IEF. Results show that FACs may exhibit either linear or saturated responses, depending on the Alfvén Mach number. The magnetospheric closure paths of region 1 currents differ significantly between the linear and saturated conditions. In the linear case, essential parts of the region 1 current extend toward dayside and connect to the magnetopause currents, whereas these parts of currents almost disappear in the saturation state. Saturation usually occurs under sub‐Alfvénic solar wind conditions, so the observed FACs generally grow linearly with IEF. Plain Language Summary: Field‐aligned currents (FACs) are highly responsive to fluctuations in the solar wind (SW) and play a crucial role in the coupling between the SW, magnetosphere, and ionosphere. Theoretical models propose that FACs may saturate under high interplanetary electric field (IEF). However, observations from satellites reveal a linear correlation between FACs and IEF. To address this disparity, we conducted comparative global simulations and scrutinized FAC magnitudes and configurations. Our findings unveil distinct structural disparities in FACs between linear and saturated conditions, and show that the FACs saturate when the magnetosheath magnetic pressure is much greater than the thermal pressure, usually under sub‐Alfvénic solar wind conditions. Key Points: Both saturation and linear variations of field‐aligned current may occur with increasing interplanetary electric field (IEF)Region 1 (R1) field‐aligned current structures are different for linear and saturation cases even under the same IEFR1 current saturates when the magnetosheath magnetic pressure is much larger than thermal pressure, possibly under sub‐Alfvénic solar wind [ABSTRACT FROM AUTHOR]

We introduce A strophysical H ybrid- K inetic simulations with the flash code (⁠|$\tt {AHKASH}$|⁠) – a new Hybrid particle-in-cell (PIC) code developed within the framework of the multiphysics code flash. The new code uses a second-order accurate Boris integrator and a predictor–predictor–corrector algorithm for advancing the Hybrid-kinetic equations, using the constraint transport method to ensure that magnetic fields are divergence-free. The code supports various interpolation schemes between the particles and grid cells, with post-interpolation smoothing to reduce finite particle noise. We further implement a |$\delta f$| method to study instabilities in weakly collisional plasmas. The new code is tested on standard physical problems such as the motion of charged particles in uniform and spatially varying magnetic fields, the propagation of Alfvén and whistler waves, and Landau damping of ion acoustic waves. We test different interpolation kernels and demonstrate the necessity of performing post-interpolation smoothing. We couple the turbgen turbulence driving module to the new Hybrid PIC code, allowing us to test the code on the highly complex physical problem of the turbulent dynamo. To investigate steady-state turbulence with a fixed sonic Mach number, it is important to maintain isothermal plasma conditions. Therefore, we introduce a novel cooling method for Hybrid PIC codes and provide tests and calibrations of this method to keep the plasma isothermal. We describe and test the 'hybrid precision' method, which significantly reduces (by a factor |$\sim 1.5$|⁠) the computational cost, without compromising the accuracy of the numerical solutions. Finally, we test the parallel scalability of the new code, showing excellent scaling up to 10,000 cores. [ABSTRACT FROM AUTHOR]

Solar activity dynamics are explored through an in-depth analysis of the interplay between sunspot numbers and critical magnetohydrodynamic parameters − specifically Alfvén Mach number and Magnetosonic Mach number − over the past four solar cycles (SC). Our investigation reveals a robust negative correlation between SSN and both Alfvén Mach number and Magnetosonic Mach number, shedding light on the intertwined nature of solar magnetic phenomena and magnetohydrodynamic processes. Significant temporal synchronicities are unveiled, elucidating compelling alignments between specific features of Alfvén Mach number and Magnetosonic Mach number and the peaks and troughs of SSN throughout the solar cycles. This temporal coherence underscores the complex interplay between solar magnetic activity and the broader dynamics of magnetohydrodynamic phenomena, providing deeper insights into solar cycle behavior. To enhance our understanding and predictive capabilities, we deploy an optimized LSTM+model for forecasting Alfvén Mach number and Magnetosonic Mach number in the ongoing solar cycle, SC-25. Rigorous validation of the model's accuracy is achieved through meticulous examination of prediction results for SC-24, affirming the reliability and robustness of our predictive framework. Furthermore, the anticipated timing of the first appearance to peak and the overall peak of SSN in SC-25 is calculated as 2 Jun. 2023 ± 34 days and 16 Jan. 2025 ± 27 days, respectively. Notably, these projections suggest the possibility of a double peak phenomenon in SC-25, characterized by comparable intensity levels around 160. [ABSTRACT FROM AUTHOR]

This research studied the performance of a spindle‐controlled steam ejector using models such as ideal gas and wet steam under various operating conditions based on computational fluid dynamics (CFD). A wet steam model incorporating nonequilibrium condensation was employed to simulate the complex flow phenomena within the ejector. The structure of the flow, entrainment ratio (Er), and shock wave characteristics of the steam ejector were examined in two different models. Results indicate that the spindle position has a substantial impact on steam ejector performance. For the ideal gas and wet steam models, the optimal spindle position (SP‐5) at a .1 MPa motive pressure achieves the highest entrainment ratios (Er) of 1.01 and 1.042, respectively. However, an ejector with a fixed geometry achieves Er values of only .517 and .549 for the ideal gas and wet steam models, under identical working conditions. This represents a substantial improvement of 89.8% over the fixed‐geometry ejector. The wet steam model consistently predicts 2%–4% higher Er values compared with the ideal gas model across all spindle positions. The study also reveals that increasing the motive pressure from .1 to .3 MPa reduces Er by up to 45.8% at the optimal spindle position, with the shock train length extending to 35% of the mixing chamber at .3 MPa. These findings offer insights for improving the design and optimization of variable‐geometry steam ejectors, potentially increasing efficiency in industrial applications. [ABSTRACT FROM AUTHOR]

In this paper, we consider a singular limit problem for a diffuse interface model for two immiscible compressible viscous fluids. Via a relative entropy method, we obtain a convergence result for the low Mach number limit to a corresponding system for incompressible fluids in the case of well‐prepared initial data and same densities in the limit. [ABSTRACT FROM AUTHOR]

We investigate the numerical performance of a Discontinuous Galerkin (DG) hydrodynamics implementation when applied to the problem of driven, isothermal supersonic turbulence. While the high-order element-based spectral approach of DG is known to efficiently produce accurate results for smooth problems (exponential convergence with expansion order), physical discontinuities in solutions, like shocks, prove challenging and may significantly diminish DG's applicability to practical astrophysical applications. We consider whether DG is able to retain its accuracy and stability for highly supersonic turbulence, characterized by a network of shocks. We find that our new implementation, which regularizes shocks at subcell resolution with artificial viscosity, still performs well compared to standard second-order schemes for moderately high-Mach number turbulence, provided we also employ an additional projection of the primitive variables on to the polynomial basis to regularize the extrapolated values at cell interfaces. However, the accuracy advantage of DG diminishes significantly in the highly supersonic regime. Nevertheless, in turbulence simulations with a wide dynamic range that start with supersonic Mach numbers and can resolve the sonic point, the low-numerical dissipation of DG schemes still proves advantageous in the subsonic regime. Our results thus support the practical applicability of DG schemes for demanding astrophysical problems that involve strong shocks and turbulence, such as star formation in the interstellar medium. We also discuss the substantial computational cost of DG when going to high order, which needs to be weighted against the resulting accuracy gain. For problems containing shocks, this favours the use of comparatively low DG order. [ABSTRACT FROM AUTHOR]

The propagation of sound in a uniform flow with rigid body swirl has been considered in the approximation of low swirl Mach number M 2 ≪ 1 ,where the swirl Mach number M = Ω r / c 00 is specified by the angular velocity Ω , stagnation sound speed c 00 and radial distance r; in this case the mass density and sound speed in the mean flow are constant, and the convected wave equation is solved in terms of Bessel functions. In this paper the restriction on swirl Mach number is relaxed from M 2 ≪ 1 to M 4 ≪ 1 , thus keeping O ( M 2 ) terms in the mean flow, mass density and sound speed, that become non-uniform; the acoustic vortical wave equation is no longer the convected wave equation, and is extended to this case and solved in terms of generalized Bessel functions. The radial dependence to second order in the swirl Mach number is specified by a generalized Bessel differential equation for decoupled acoustic or vortical modes and for acoustic-vortical waves to first order in the swirl Mach number. The solutions are obtained: (i) for finite radius in terms of generalized Bessel and Neumann functions, that determine the radial wavenumbers, natural frequencies and normal eigenfunctions for cylindrical or annular ducts with rigid or impedance wall boundary conditions; (ii) asymptotically for large radius in terms of generalized Hankel functions, specifying the growth of amplitude with radius either as a power law or as an exponential of one-half of the square of the swirl Mach number. Thus the compressible uniform mean flow with rigid body rotation is spatially unstable in the radial direction; it is also unstable in time for cut-on modes with real axial wavenumbers and cut-off modes with imaginary axial wavenumbers. Compared with acoustic waves the acoustic-vortical waves have more modes, some with complex rather than real eigenvalues leading to instabilities. [ABSTRACT FROM AUTHOR]

For propeller-driven Mars airplanes operating at low Reynolds numbers, the speed of the rotor tips may reach transonic. To date, only a few studies have investigated the transonic buffet in flow past low-Reynolds-number airfoils. In this study, direct numerical simulations of high-speed flow (M=0.2 , 0.6, and 0.8) past a NACA 0012 airfoil are performed with a low Reynolds number of Rec=23,000 , where transonic buffet is observed at M=0.8. We first investigated the effects of Mach number on the aerodynamic performance and flow fields. Dynamic mode decomposition (DMD) and linear stability analysis (LSA) are used to analyze the flow instability mechanisms of the transonic buffet. Results showed that the multiple high-frequency oscillations are related to the vortex shedding at the trailing edge of the airfoil, and the low-frequency oscillation is caused by Type C shock motions. Both of them are confirmed to be self-sustained in feedback cycles. [ABSTRACT FROM AUTHOR]

This study investigates the impact of upstream energy deposition as an active flow control method on the performance of a two-dimensional scramjet inlet through numerical simulations. An in-house inviscid finite volume solver was employed for the simulations, with boundary reconstruction of the scramjet achieved using a ghost-cell based immersed boundary method. The effects of different energy spot locations, various energy strengths, and different freestream Mach numbers on the scramjet performance were thoroughly analyzed. Four inlet parameters, namely, mass capture ratio (μ), total pressure ratio (TPR), kinetic energy efficiency (η K E), and flow distortion index (DI), were used to assess the scramjet performance. The analysis revealed that energy deposition significantly reduced flow distortion, resulting in noticeable improvements in mass flux and total pressure ratio. However, kinetic energy efficiency exhibited minimal change. The study provides comprehensive insights into the impact of energy deposition on scramjet performance, highlighting its potential for enhancing inlet characteristics. • Effects of upstream energy deposition on scramjet inlet performance are numerically investigated. • Scramjet performance is assessed using four aerothermodynamic parameters. • Upstream energy deposition significantly reduces flow distortion and improves mass flux. • The location and strength of the deposited energy strongly influence the performance parameters. [ABSTRACT FROM AUTHOR]

Diffraction is a fundamental phenomenon that occurs when blast or shock waves pass over sudden discontinuous surfaces. It generates a complex flow field consisting of diffracted waves, expansion waves, slipstream, contact surface, and an unstable shear layer, in addition to emitting acoustic waves. In this study, we investigated the diffraction of a blast wave passing over a series of grooved structures with different aspect ratios and geometrical shapes (rectangular, circular, and triangular) using high-speed shadowgraph images. The blast wave Mach number considered in our investigation is 1.34. The grooves feature leading-edge geometrical variations such as rectangular, circular arc, and wedge shapes positioned at various lateral locations from the exit of the shock tube. The aspect ratios of the rectangular grooves vary from 0.33, 0.5, and 0.67. The circular and triangular grooves have an aspect ratio of 0.33. The trajectories and velocities of the primary vortex are calculated by tracking the location of the vortex in the shadowgraph images. Our observations revealed that a large portion of the incident blast wave is abducted inside the groove as the aspect ratio increases in rectangular grooves, resulting in better attenuation of the blast wave. The grooves, which have circular shapes, produced weaker diffraction, which resulted in delayed and weak primary vortex. The triangular grooves produced the strongest primary vortex and have the highest attenuation characteristics among other grooves. The strength and trajectory of the primary vortex formed over the grooves strongly depend on the aspect ratio and the curvature of the leading edge for a given Mach number. Vortices generated from rectangular and triangular grooves exhibit considerable strength and longevity. [ABSTRACT FROM AUTHOR]

A three-dimensional numerical study is performed to investigate the effect of inlet mass flow rate on flow field characteristics and pressure gain of rotating detonation combustor. The results show that the mass flow rate has an important effect on the behavior of the detonation wave. When the inlet mass flow rate is increased, the flow Mach number of inlet throat increases and flow develops to a close critical state. After ignition, two counter-rotating detonation waves will collide in the combustor and the number of collisions decreases with the increase of inlet mass flow rate. The strong wave eventually engulfs the weak wave and develops into a single-wave mode. When the inlet mass flow rate is reduced, the detonation wave velocity and refill height of premixed gas increase. Deflagration area of flow field reduces and the stability is improved. In addition, when the inlet mass flow rate is reduced, the total pressure loss caused by expansion wave decreases. The pressure gain and fuel-based specific impulse are increased. However, increasing the total pressure gain also improves the static pressure fluctuation ratio of inlet throat. The phenomenon of high-pressure feedback under small flow rate condition is also more prominent. [ABSTRACT FROM AUTHOR]

In this paper, a novel two-stage hydrocarbon fuel heating and delivery system was used to more accurately simulate the state of fuel in the cooling channel of the regenerative cooling engine. Mildly cracked (<10%) experiments of n-dodecane were conducted by using the new heating system and pyrolysis products were sampled and analyzed. The ignition delay time and laminar flame speed of the cracked products were different from those of the surrogate gas fuel. When the temperature is 1100 K, the ignition delay time of the cracked products are ~50% higher than that of surrogate gas fuel. The laminar flame speed of the cracked products is about 30% lower than that of surrogate gas fuel at the equivalence ratio of 1.0. The supersonic combustion experiments of mildly cracked and different equivalence ratios n-dodecane were also employed at Mach number 3.0 of the isolator entrance, with total temperature ~1550 K and total pressure ~1.5 MPa. The static pressure along the axis of combustor was measured, while the velocity and temperature of the airflow at the exit of combustor were also measured by the tunable diode laser absorption spectroscopy (TDLAS). The distributions of airflow velocity, static temperature, and total temperature along the combustor at different conditions were analyzed through a one-dimensional analysis method. The deviations of velocity and temperature of the airflow are less than 4% and −6%, respectively, which indicates the calculation is in great accordance with the TDLAS detections. The experimental data will be helpful for validating the simulation of supersonic combustion, and studying the influence of mildly cracked fuel on supersonic combustion characteristics. [ABSTRACT FROM AUTHOR]

The terrestrial bow shock is the boundary where supersonic solar wind slows down abruptly near the magnetopause. The shock front geometry could be modulated by surface waves to form rippled structures, which impact the acceleration process of the solar wind particles. However, the rippled structures are hard to be identified unambiguously due to the similar signatures in single‐spacecraft observations between rippled shocks and reforming shocks. Here, we utilize the four‐spacecraft observations from the MMS mission to investigate an event of quasi‐perpendicular bow shock crossing. The periodic oscillations of shock normal directions and normal velocities support the scenario of surface wave propagation in the tangential direction. We also reconstruct the shock profile along the normal direction, and its monotonic shape further excludes the occurrence of shock reformation. These ripples are found to modulate the reflected ions and whistler wave packets, which adds to the complexity of the bow shock plasma environments. Plain Language Summary: Collisionless shocks are ubiquitous structures in the space and astrophysical plasma environments where significant energy conversion occurs between electromagnetic energy, kinetic energy, and thermal energy. These structures often exhibit nonstationary signatures, from which important information on the associated plasma dynamics can be extracted. In this paper, we report the spacecraft observations of a bow shock crossing with periodic enhancements of the magnetic field strength. Such signatures have been previously attributed either to the rippled shock structures or to the shock reformation process. These two potential processes are examined via simultaneous observations at four different locations, which shows the periodic oscillations of the local shock normal directions indicative of rippled rather than reforming shocks. We also show that the rippled shocks could spatially modulate the reflected ions and whistler wave packets, so they can also be detected periodically by the spacecraft. This phenomenon has been reported on other planets across a wide range of Mach numbers, indicating that the shock ripples, universal in the plasma universe, play a critical role in impacting particle dynamics and shock evolution processes. Key Points: Periodic magnetic strength enhancements are observed during the bow shock crossing of the MMS spacecraftMulti‐spacecraft analysis supports the scenario of shock ripples rather than the shock reformation processThe surface waves alter the shock geometry, modulating the reflected ions and whistler wave packets [ABSTRACT FROM AUTHOR]

The accurate computation of different turbulent statistics poses different requirements on numerical methods. In this paper, we investigate the capabilities of two representative numerical schemes in predicting mean velocities, Reynolds stress and budget of turbulent kinetic energy (TKE) in low Mach number flows. With concerns on numerical order of accuracy, dissipation and dispersion properties, a high-order upwind scheme with relatively good dispersion and dissipation and a second-order non-dissipative central scheme with perfect dissipation but poor dispersion are adopted for this comparative study. By carrying out a series of numerical simulations including Taylor-Green vortex, turbulent channel flow at R e τ = 180 and turbulent flow over periodic hill at R e b = 10595 , it can be obtained that although the high-order upwind scheme lacks perfection of dissipation in high wave number range, it still demonstrates superior predictive capability compared with the second-order non-dissipative central scheme, especially with relatively coarse grids. Finally, by taking the high-order upwind scheme as a suitable selection for turbulence simulation, the turbulent flow over a 30P30N multi-element airfoil is investigated as an application study. After briefly comparing the simulated profiles and spectrum with reference experimental results as validation, the budget of TKE is analyzed to locate the dominant flow structures and regions. It is found that the production and dissipation terms behave in a "monopole" pattern in the locations with strong shears and wakes. Whereas the advection and diffusion terms show an "inward" pattern and an "outward" pattern, which indicate the spatial transport of TKE between the center of the shear layer and nearby locations. [ABSTRACT FROM AUTHOR]

The aim of this work is to develop a numerical computation program allowing designing new contours of a supersonic axisymmetric nozzle having two expansions at the throat, named by DEN (Dual Expansion Nozzle). This new nozzle gives a uniform and parallel flow at the exit section, to improve considerably the performances compared to the conventional Minimum Length Nozzle (MLN), and the Best Performances Nozzle (BPN). The present nozzle has a two unknowns external and central body curved walls. Each of them is started by an initial expansion angle to give a uniform and horizontal flow at the exit section. Two others transition regions are calculated in parallel with the contours points to give the desired exit Mach number. The walls are determined point by point by the High Temperature Method of Characteristics (HT MOC) model. The resolution of the four compatibility and characteristics equations is done numerically by the finite difference predictor corrector algorithm. The validation of the results is controlled by the convergence of the calculated critical sections ratio to that given by the theory. The design depends on four parameters, where MLN and BPN become special cases of DEN. A comparison is made with MLN , since it is currently used in the aerospace propulsion and with BPN aiming to improve their performances. The comparison is made for the same critical mass flow rate. The results demonstrate a remarkable reduction up of 45 %, and 52 % in the mass of DEN when the exit Mach number M E = 3.00 and the stagnation temperature T 0 = 2000 K. The application is made for air and for future aerospace missiles in order to improve their trajectory parameters. The chosen example demonstrates an improvement of 13 % and 16 % on the missile range compared, respectively to MLN , and BPN. [ABSTRACT FROM AUTHOR]

To suppress the adverse effect of the gear windage phenomenon in the high-speed aeronautic industry, a shroud as an effective alternative strategy is usually to enclose gears to reduce the windage behaviors of high-speed gears. To deeply understand these no-load power losses, this paper proposes a new simulation methodology based on the Lattice Boltzmann method to study the windage losses of a shrouded spur gear and conducts a series of numerical studies. The models reproduce a shroud spur gear varying radial and axial clearances to evaluate the influence of casing walls on windage losses. The simulation results were then compared with experimental values, showing a satisfactory agreement. Furthermore, a torque containment factor integrating the air compressibility at high Mach numbers is introduced to represent the reduction in torque (windage power losses) for the shrouded gear compared to the free case, and the theoretical formulae for predicting windage power losses are further improved for better applicability as the tight shroud approaches the gear during the preliminary design stage. [ABSTRACT FROM AUTHOR]

The reaction zone structure of spontaneous reaction wave at steady subsonic and supersonic speeds is discussed using numerical simulation and theoretical approaches. The main objective of this study is to develop the mathematical model of the spontaneous reaction wave propagating at a steady speed and to clarify the relationship between the reaction zone structure and non-dimensional parameters. A one-dimensional direct numerical simulation (DNS) of the spontaneous reaction wave propagation is performed, and the modeling assumptions are discussed based on the numerical results. The non-dimensional parameters, the Damköhler number and a parameter related to compressibility defined by the Mach number of the spontaneous reaction wave velocity, are introduced into the modeling. The mathematical model is developed using asymptotic techniques, and the solution of the reaction zone structure is analytically derived using Newton’s method. The relationship between reaction zone structure of spontaneous reaction wave and non-dimensional parameters is discussed based on the analytical solutions. [ABSTRACT FROM AUTHOR]

In high-speed flow applications, increased aero-thermodynamic loads, localized turbulent regions, instability, excessive drag and pressure losses are among the negative effects of the interaction between shock waves and boundary layers and the subsequent separation of the boundary layer. These drawbacks emphasize how crucial it is to look into ways to control the boundary layer separation caused by shock waves. In order to determine whether using a pressure feedback technique to regulate boundary layer separation on a two-dimensional forward-facing step (FFS) is effective, a numerical study is conducted. The pressure feedback technique utilizes a channel to remove fluid from the separated area and reintroduce it into the mainstream flow, by leveraging the pressure contrast as the motivating factor. Ansys Fluent 2021 simulation software is used to conduct a numerical study investigating shock wave boundary layer interaction over a FFS at a supersonic Mach number of 2.5. The present study aims to explore the effectiveness of implementing a pressure feedback channel near the FFS as a strategy for controlling flow separation induced by the step. Turbulence effects are simulated using a k-omega model, and the study employs a finite-volume method with upwind flux difference splitting and second-order upwind flow discretization. The simulation results indicate the efficacy of the pressure feedback mechanism in reducing flow separation, thus enhancing its role in separation control. Integration of an 80mm channel results in a notable 42.85% decrease in separation bubble size, while the longer 90mm channel yields an even more substantial 52.38% reduction. [ABSTRACT FROM AUTHOR]

The work studies scenarios of self-ignition of hydrogen-air mixtures after a shock wave reflects from a destructible barrier made of sand or a solid wall. The experiments were carried out using a shock tube. The transverse dimensions of the diagnostic section were 40 × 40 mm. The initial pressure of the hydrogen-air mixture varied from 10 kPa to 50 kPa. The molar excess of hydrogen (equivalence ratio) was 0.3, 0.4 and 0.5. Typical oscillograms of pressure, a photomultiplier tube, and the results of high-speed visualization using the Schlieren method are presented. When the shock wave interacted with the destructible barrier, a thermodynamic unloading of shock-compressed mixture occurred in the direction of the scattering of fragments of the barrier. It was shown that under certain conditions, self-ignition of hydrogen could be prevented by using a destructible barrier. In this case, the prevention was recorded in a narrow range of Mach numbers of the incident shock wave, which are characterized by the occurrence of the so-called "weak" (or "mild") ignition of the hydrogen-air mixture. The conditions for strong ignition behind the reflected shock wave, as well as during the interaction of the reflected shock wave with the flame front behind the incident shock wave, were determined. The results of the presented experimental work are aimed at ensuring explosion safety when working with hydrogen in a confined space. [Display omitted] • The interaction of a shock wave with a sand barrier was studied. • Scenarios for self-ignition behind a reflected shock wave were determined. • Self-ignition can be prevented by placing a sand barrier instead of a solid wall. • The pressure ratios were determined during self-ignition. [ABSTRACT FROM AUTHOR]

The article in the Journal of Fluid Mechanics explores dynamic stall onset and vortex shedding in low-Reynolds-number flows, focusing on the leading edge suction parameter (LESP) and boundary enstrophy flux (BEF) as predictive criteria. High-resolution simulations of an airfoil at two Reynolds numbers reveal various unsteady flow phenomena, with BEF effectively identifying stall onset and vortex shedding events. The study underscores the significance of these criteria in comprehending low-Reynolds-number unsteady flows and their implications for flow control. The research, supported by NSF and US Air Force grants, sheds light on the dynamics of vortex shedding on airfoils at low Reynolds numbers. [Extracted from the article]

The study examines the Richtmyer–Meshkov and Kelvin–Helmholtz instabilities on a shock-accelerated inclined single-mode interface through shock-tube experiments. It finds that the Richtmyer–Meshkov instability drives early-time amplitude growth, while the Kelvin–Helmholtz instability contributes to later-stage growth. The research demonstrates that nonlinear flow features emerge faster with larger angles and stronger shocks. The study validates theoretical models for predicting interface morphology and amplitude evolution in the early stages of these instabilities, shedding light on the coupling of Richtmyer-Meshkov and Kelvin-Helmholtz instabilities in shock tube experiments. [Extracted from the article]

The article presents a new wall-pressure spectral model for non-equilibrium boundary layers, addressing the limitations of existing semi-empirical models in strong pressure gradient conditions. By analyzing high-fidelity simulations and experimental data, the study identifies optimal scaling variables for accurate wall-pressure spectrum predictions. The proposed model, incorporating local mean velocity profiles, shows improved performance in capturing the effects of pressure gradients on wall-pressure statistics in turbulent boundary layer flows. Overall, the research provides valuable insights into the impact of Reynolds number and pressure gradients on wall-pressure spectra, enhancing our understanding of non-equilibrium boundary layer dynamics. [Extracted from the article]

The effect of dust is studied on the parallel propagation of a solitary hydromagnetic wave in a magnetized, cold, collision-free plasma. The fully nonlinear analysis is based on the Adlam–Allen method. The system of nonlinear differential equations was derived using the assumption of massive dust, namely, that the dust velocity is constant and parallel to the background magnetic field. Self-consistent solutions obtained from the model for the lighter particles dynamics and for the resulting electric and magnetic fields are presented, together with the validity range of Mach numbers as a function of dust parameters. In particular, supersonic solutions are possible in scenarios tending to two-component plasmas (positively charged ions and negatively charged dust). The stationary solitary pulse in terms of particle densities, magnitude of transverse magnetic field, transverse electron–ion velocities, and bipolar electric field is presented. The magnitude of these amplitudes is maximum at the center of wave and minimum at the boundary. Furthermore, it is shown that all the solitary wave amplitudes are decreases with dust parameter. [ABSTRACT FROM AUTHOR]

The flow loss caused by the fan blades in a turbofan engine with a large bypass ratio is significant, and the wake considerably affects the inlet flow of downstream components. Surfaces with bird feather-like convergent–divergent (C–D) riblets have been proven to modulate the boundary layer flow by inducing counter-rotating rolling modes; however, the effects of these surfaces on the total pressure loss and wake turbulence of transonic compressor cascades remain unexplored. In this study, the effects of C–D-riblet surfaces on the total pressure loss and wake turbulence of a transonic compressor cascade were experimentally investigated using a five-hole probe and hot-wire measurements. The flow-loss-control effects of C–D-riblet surfaces with different characteristic lengths were also analyzed. The most significant reduction in the area-averaged total pressure loss (11.23%) was achieved using a C–D-riblet surface with a characteristic length of 30 μm at a Mach number of 0.94; this total pressure loss reduction corresponded to an increase in the mean velocity and a decrease in the turbulence intensity of the wake profile. Furthermore, the effectiveness of the loss control varied significantly with the spreading position of the C–D riblets. The optimal control effect was observed in the divergence-line region, and the control was slightly less effective as the measurement position neared the convergence line. This paper demonstrates the promising potential of using C–D riblets to achieve flow loss control in transonic compressor cascades. [ABSTRACT FROM AUTHOR]

The gradient of flow parameters in a transonic compressor cascade flow field varies significantly, especially in the region of shock waves, which causes a significant challenge to its high-precision flow field prediction. In this study, the position query-guided cross-modal flow field prediction model (PGCM) is proposed to effectively predict the flow field parameter distribution of a transonic compressor cascade. The PGCM utilizes the self-attention mechanism for the global and deep geometric feature extraction of configurations, which contributes to an in-depth understanding of the spatial relationships between coordinate points within the flow field, accurately capturing and analyzing the structural complexity of a compressor cascade flow. In addition, the PGCM integrates the cross-attention mechanism that establishes correlations between different input sequences, which enhances the performance of the model in querying and interpreting flow parameters at specific coordinates. The flow field prediction models are developed to predict the flow parameter distributions of different cascade geometries at Mach numbers of 0.78 and 0.93, respectively. The validation results indicate that the PGCM performs significantly better than the existing convolutional neural network and vision transformer, especially in the prediction of the pressure coefficient Cp distribution. The PGCM is adaptable to the variation of flow conditions and geometrical configurations efficiently and accurate in predicting the flow field of a compressor cascade. This paper demonstrates the promising potential of conducting the multi-modal information fusion to enhance the capability of flow field prediction. [ABSTRACT FROM AUTHOR]

Complex aerodynamic interference has long posed challenges to the safe separation of multi-body systems. Based on a new conceptual ground horizontal-sliding takeoff approach, the rocket sled system, computational fluid dynamics (CFD) simulations are conducted to study the separation between the payload and the rocket booster, with the Advisory Group for Aerospace Research and Development Model B representing the payload. The motion trajectory of the payload and the unsteady interaction flow mechanisms during separation are analyzed by altering the payload's layout, center of gravity (CoG), and sliding Mach number. The results show that positioning the payload closer to the front along the axis increases the likelihood of collisions after separation. At a sliding velocity of M a = 3 , the payload tends to exhibit a nose-down attitude when l c / l p < 0.64. Conversely, when l c / l p > 0.64 , the pitching angle becomes excessively large, which increases the risk of stalling or somersaulting. Furthermore, the CoG needs to move forward to reduce the growth rate of the pitching angle when the separation Mach number is increased. These studies can provide valuable references for the development of separation technology for the rocket sled system. [ABSTRACT FROM AUTHOR]

The supersonic inflow passes through the shock train in the isolator of the scramjet to complete deceleration and pressurization, followed by combustion and energy release, providing strong thrust. When the back pressure generated by combustion is disturbed forward, the location of shock train leading edge (STLE) will also change accordingly. Once it moves to the entrance of the isolator, it will cause unstart. Accurately detecting STLE in the isolator of a scramjet is crucial for controlling the shock train and preventing the inlet from unstart. Therefore, based on the sparse reconstruction of compressive sensing and sensor compression coupling, a supersonic flowfield reconstruction model (POD-STLE) based on proper orthogonal decomposition (POD) was constructed to reconstruct the supersonic flowfield and detect the location of STLE in the supersonic isolator. The experiments were conducted on the shock oscillation under variable Mach numbers and back pressures, to construct the experimental dataset. Combining supersonic flowfield reconstruction and matrix decomposition, different sensor layouts were constructed, which can ensure accuracy and stability while saving sensor resources. The POD-STLE was applied to the flowfield reconstruction of the supersonic isolator, and the location of STLE was detected under variable and constant conditions, ultimately achieving the expected reconstruction effect and detection accuracy. This study provides a new research method for detecting the location of STLE in the supersonic isolator of a scramjet and provides technical for exploring supersonic flowfield. [ABSTRACT FROM AUTHOR]

The vortex ring is usually analytically modeled in computational studies involving vortex ring and shock wave interaction. However, incompressibility assumptions of the isolated vortex ring model mostly limits its applicability to the lower Mach number regime. To overcome this limitation, a compressible vortex ring is generated using a pulse jet exiting a circular nozzle, and its interaction with a shock wave is analyzed. The shock–vortex interaction and resulting sound wave generation is simulated by solving the compressible axisymmetric Navier–Stokes equations. A characteristics-based filter is used to isolate acoustic and hydrodynamic disturbances in the flow field. A reconfigured computational domain is used to negate the aftereffects of the nozzle on the flow field. The strong interaction of compressible vortex rings at higher Mach numbers with different shock wave Mach numbers in the flow field is studied including resultant acoustic wave generation. [ABSTRACT FROM AUTHOR]

This paper presents a novel framework for the active control of transonic airfoil flutter using synthetic jets through deep reinforcement learning (DRL). The research, conducted in a wide range of Mach numbers and flutter velocities, involves an elastically mounted airfoil with two degrees of freedom of pitching and plunging oscillations, subjected to transonic flow conditions at varying Mach numbers. Synthetic jets with zero-mass flux are strategically placed on the airfoil's upper and lower surfaces. This fluid–structure interaction (FSI) problem is treated as the learning environment and is addressed by using the arbitrary Lagrangian–Eulerian lattice Boltzmann flux solver (ALE-LBFS) coupled with a structural solver on dynamic meshes. DRL strategies with proximal policy optimization agents are introduced and trained, based on the velocities probed around the airfoil and the dynamic responses of the structure. The results demonstrate that the pitching and plunging motions of the airfoil in the limited cycle oscillation (LCO) can be effectively alleviated across an extended range of Mach numbers and critical flutter velocities beyond the initial training conditions for control onset. Furthermore, the aerodynamic performance of the airfoil is also enhanced, with an increase in lift coefficient and a reduction in drag coefficient. Even in previously unseen environments with higher flutter velocities, the present strategy is achievable satisfactory control results, including an extended flutter boundary and a reduction in the transonic dip phenomenon. This work underscores the potential of DRL in addressing complex flow control challenges and highlights its potential to expedite the application of DRL in transonic flutter control for aeronautical applications. [ABSTRACT FROM AUTHOR]

A local preconditioning high-order discontinuous Galerkin (DG) method is introduced to enhance the efficiency and accuracy of the density-based approach for low-speed flows. The method is based on the time-derivative preconditioning technique typically employed in finite volume methods. Traditional approximation methods that only precondition low-order derivatives lead to incompatibility of the equations. To extend the preconditioning technique to the DG method, we employ an analytic preconditioning mass matrix derived from the DG formulation of the preconditioned Navier–Stokes equations. Modified numerical flux functions and a self-adaptive local velocity truncation parameter are employed for DG systems with preconditioning. We demonstrate the efficiency and accuracy of the preconditioned DG method using explicit and implicit schemes for steady flows and a dual-time scheme for unsteady flows. The method has been implemented and used to compute a variety of flow problems, including inviscid and laminar flows, utilizing various degrees of polynomial approximations. Computations with and without preconditioning are performed to analyze the influence of spatial discretization on the accuracy and convergence of the DG solutions at low Mach numbers. Numerical results underscore the enhanced accuracy and convergence properties of the proposed method for low-speed flows, indicating significant performance improvements over traditional methods. [ABSTRACT FROM AUTHOR]

Aero-optical steady-distortion reduces the far-field quality of the forward-looking laser beam of an airborne turret. It has a complex relation with the viewing angle of the laser beam and the flow around the turret at different flight speeds. In this paper, the aero-optical steady-distortion, including the fixed tip/tilt and steady-lensing wavefront, for a hemisphere-on-cylinder turret with a conformal window is analyzed at different freestream Mach numbers (M a ∞ = 0.3 – 1.7) and window angles. Computational fluid dynamics and particle-based ray-tracing methods are combined to simulate the steady-state flow over the turret and its aero-optical effect. The scaling laws of the fixed tip/tilt and steady-lensing wavefront for the forward-looking laser beam are summarized under these subsonic, transonic, and supersonic flows. The results show that the scaling factor M a ∞ 2 is only applicable to the wavefront distortion caused by low-velocity flows, while the relevant scaling factors M a ∞ 2 / [ 1 + (γ + 1) M a ∞ 2 (γ + 1) / (γ − 1) ] and (γ − 1) / (2 γ) M a ∞ γ / (γ − 1) are effective across subsonic/transonic and supersonic flows. The fixed tip/tilt is discussed separately, and its scaling factor and scaling law differ from the steady-lensing wavefront. In addition, the fixed tip/tilt and steady-lensing wavefront within these flow regimes are only related to the viewing angle, not the elevation angle. [ABSTRACT FROM AUTHOR]

The wave dynamics play a crucial role in the operation characteristics of the rotating detonation engine. We conducted numerical simulations of a rotating detonation combustor (RDC) using multicomponent reactive Navier–Stokes equations coupled with a discrete phases model. The RDC in this research employs a configuration with multiple coaxial injectors supplying oxygen-enriched air and kerosene spray at room temperature. To accurately identify and analyze waves within the RDC, we proposed a three-dimensional transient detonation wave detection method based on the combined parameters of normal Mach number and heat release rate in the flow field. Two typical wave modes, referred to as single-wave mode and counter-waves mode, are identified and then selected to conduct a detailed wave dynamics analysis. The general wave behavior is discussed, and velocity deficit is compared for these two wave modes. For the single-wave mode, intermittent micro-explosions are observed generating retonation waves periodically in the unburnt pockets behind the rotating detonation shock front. For the counter-waves mode, we analyzed the collision process of the two waves and the coupling/decoupling of the shock front with the detonative heat release zone, revealing the reason for significant velocity deficits in this wave mode. This research demonstrates that micro-explosions intermittently occur in the multiphase RDC in both single-wave and counter-waves modes and generate micro explosion shock waves periodically, which influence the complicated wave dynamics behavior. [ABSTRACT FROM AUTHOR]

When optical imaging-guided aircraft flies at hypersonic speeds in the atmosphere, the optical window withstands severe aerodynamic heating. Conducting the thin film resistance thermometer measurements in a hypersonic gun wind tunnel with a Mach number of 7.1 and total temperature of 670 K, the study investigates the effect of nozzle pressure ratio (NPR = film exit static pressure/nearby mainstream static pressure) on supersonic (Mach 2.43) film cooling for the hypersonic optical window. By combining the flow information near the window obtained using the three-dimensional compressible Reynolds-averaged Navier–Stokes method, the study reveals the mechanism of the effect of NPR on film cooling. The results indicate that increasing NPR can enhance the momentum of the unit volume film and improve the film's ability to resist mainstream mixing. Moreover, the film with a large NPR can better maintain its own momentum, leading to an increase in the film effective cooling length and film cooling effectiveness. The film effective cooling length corresponding to the unit mass flow rate of the cooling gas increases with the increase in NPR. It verifies the nonlinear relationship between the film cooling performance and the coolant mass flow rate, indicating the additional benefits of increasing NPR on the film cooling performance. Through research, it is found that increasing NPR can increase the film thickness, thereby enhancing its ability to isolate the mainstream. Moreover, as NPR increases, the cooling film expands, objectively leading to the widening of the film flow channel, allowing the Mach number of the supersonic film to increase continually. This further reduces the static temperature of the film in the flow field, thereby enhancing its cooling capability for the mainstream. [ABSTRACT FROM AUTHOR]

While the flow mechanisms of two-dimensional supersonic cooling films have been studied in-depth, this paper used the nanoparticle planar laser scattering and particle image velocimetry techniques to investigate the flow of supersonic conical cooling films at different angles of attack (AOAs). The mainstream Mach number was M a ∞ = 3.8 , and the supersonic conical cooling film was tangentially injected via a precisely calibrated M a j = 2.8 annular nozzle. Initially, the streamwise boundary layer transition process without cooling film injection was analyzed. The boundary layer transition on the leeward side occurred prematurely, whereas on the windward side, the transition process was notably delayed. Subsequently, the supersonic conical cooling film flow was qualitatively and quantitatively evaluated from the perspectives of turbulent structures, and the time-averaged and statistical characteristics of the velocity field. On the windward side, as the ratio of static pressure decreased, the effective cooling length also decreased with an increase in AOA. On the leeward side, at a small positive AOA, the supersonic conical cooling film mixed with the low-energy fluid within the thickened inner layer of the mainstream boundary layer, which mitigated the growth rate of the mixing layer and ultimately enhanced the effective cooling length. With a further increase in AOA, the supersonic conical cooling film experienced the three-dimensional detrimental effects of crossflow-separation vortices and downwash mainstream on the leeward surface, resulting in a decrease in the effective cooling length. [ABSTRACT FROM AUTHOR]

Quantifying the degree of universality in compressible turbulence is challenging due to the existence of different modes and their complex interactions. For a restricted family of flows, Donzis and John [Phys. Rev. Fluids 5, 084609 (2020)] showed that universal behavior is indeed observed in compressible turbulence if the ratio of dilatational to solenoidal root mean square (rms) velocities (δ = u ′ d / u ′ s ) is incorporated as a scaling parameter along with the traditional turbulent Mach number ( M t = u ′ / 〈 c 〉 , where u ′ is the rms velocity and 〈 c 〉 is the mean speed of sound). In this paper, we argue for the generality of those results by analyzing a wide range of compressible turbulent flows spanning a variety of flow configurations and setups to assess the degree of universal behavior. These include, among others, reacting flows, flows with solenoidal, thermal, and dilatational forcing, and flows with mean shear and bulk viscosity. We also performed new direct numerical simulations, which include turbulence in situations where vibrational modes of constitutive molecules are not in thermal equilibrium. Collectively, we offer the largest comparison across studies in the literature to date. We find that despite the wide range of forcing conditions and physical processes, universality holds across all these turbulent flows to a very satisfactory degree when both δ and Mt are considered as intrinsic compressibility parameters. The statistics investigated here—single-point statistics up to order four—are chosen such that they represent different ranges across the spectrum of dynamically relevant turbulence scales. We discuss the applicability of the purposed universal behavior for other key statistics in these turbulent flows, including two-point statistics and inhomogeneity effects, and the perspective it opens for modeling them. [ABSTRACT FROM AUTHOR]

The tip region of transonic turbine blades (exit Mach number of 0.95) exhibits complex flow characteristics with the coexistence of multiple shock wave systems and multiscale vortices. Based on a validated detached Eddy simulation method, unsteady flow features such as the spatial-temporal dynamic evolution of tip leakage vortices (TLV) and the periodic buffet/oscillation of shock trains inside the tip gap are revealed and discussed systematically under different incidence angles (i) and heights of tip clearance. Moreover, the distinction of tip flow structures between the subsonic and transonic conditions is also manifested. Results indicate that the wandering behavior of the TLV is influenced by both the swirling strength of the vortex itself and the interaction of adjacent secondary vortices. The TLV under a tip gap of 5% blade height (h) exhibits "binary and bimodal" wandering characteristics both along the pitchwise and spanwise direction, whereas under the 1%h case, only the pitchwise wandering is prominent. The dominant characteristic frequency of the TLV wandering under different conditions falls within the spectrum range of 2.2–2.4 kHz. As for the shock trains inside the tip clearance (τ), coherently movement back and forth along the pitchwise direction with varying amplitudes can be observed, where the magnitude within the τ = 1 % h exceeds that observed in the τ = 5 % h , depending on the intensity of the shock waves. Notably, significant shock wave oscillations are present throughout the range of the chord length (c) within the τ = 1 % h , whereas within the tip gap of 5%h, shock wave systems exhibit more pronounced oscillations predominantly near the trailing edge. [ABSTRACT FROM AUTHOR]

Hyperloop is proposed as the next generation of sustainable high-speed transport. Recently, an increasing body of literature has been amassed on Hyperloop aerodynamics, however, the vast majority of this work is numerical. Experimentally, there are few relevant studies and none are suitable for validating computational approaches. This paper presents three benchmark cases to provide a framework for computational research and to address this significant gap. Benchmark 1 provides experimental data from existing work on a projectile traveling at Mach 1.1 in ground effect. This incorporates many of the flow characteristics of a Hyperloop system, including (i) transonic Mach numbers, (ii) wall confinement, and (iii) shock formation/reflection. These experimental data are compared to Computational Fluid Dynamics simulations with a very good match seen. Next, Benchmark 2 is proposed which extends these simulations toward a baseline Hyperloop pod design operating in an axisymmetric low-pressure tube environment. This is achieved in stages by adding a full tube, scaling up the domain, reducing the air pressure, and introducing a baseline pod design. It is shown that the enclosed tube environment causes the most significant change in aerodynamic characteristics via flow choking. Nevertheless, a number of aerodynamic similarities remain, compared to Benchmark 1. Finally, Benchmark 3 is proposed to explore the impact of ground clearance of the pod. This aspect has a significant influence on the flow by deflecting the wake and the downstream shock pattern. Furthermore, the drag, downforce, and pitching moment are all found to increase with lower ground clearances. [ABSTRACT FROM AUTHOR]

Control of self-sustained oscillation and noise reduction poses a significant challenge. The present study employs Implicit Large Eddy Simulation at a Mach number of 0.85 to investigate the influence of a porous cavity floor on flow dynamics. By substituting the solid floor with porous media, the fundamental pressure–velocity relationship within the medium is established according to Darcy's law. Findings reveal marked suppression of wall pulsations, accompanied by a 10 dB decrease in sound pressure levels. The porous medium induces blowing and suction effects, effectively modulating large-scale re-circulation and mitigating shear layer instability, thereby approximating free mixing layer characteristics and suppressing cavity flow oscillations. At an optimized porosity for maximum noise reduction, altering back pressure at the cavity floor induces a transition in the local flow regime from suction-dominated to blowing-dominated state. Excessive reduction of back pressure promotes suction; conversely, increased pressure intensifies blowing, further attenuating feedback mechanisms and enhancing noise reduction. To explore noise reduction mechanisms, mode decomposition analyses demonstrate the efficacy of porous media in disrupting large-scale coherence structures within shear layer and redistributing energy from dominant modes to a broader frequency spectrum that engages smaller flow structures. This energy reallocation mechanism contributes to the mitigation of cavity flow noise and deepens insights into the role of porous media in flow modulation and noise control. [ABSTRACT FROM AUTHOR]

In recent years, nonequilibrium flows have been frequently encountered in various aerospace engineering and micro-electro-mechanical systems applications. To understand nonequilibrium physics, multiscale effects, and the dynamics in these applications, a reliable multiscale scheme for all flow regimes is required. Following the direct modeling methodology, the adaptive unified gas-kinetic scheme employs discrete velocity space to accurately capture the nonequilibrium physics, recovering the original unified gas-kinetic scheme (UGKS). By adaptively employing continuous distribution functions based on the Chapman–Enskog expansion, it efficiently handles near-equilibrium flow regions. The two regions are dynamically coupled at the cell interface through the fluxes from the discrete and continuous gas distribution functions, thereby avoiding any buffer zone between them. In this study, an implicit adaptive unified gas-kinetic scheme (IAUGKS) is constructed to further enhance the efficiency of steady-state solutions. The current scheme employs implicit macroscopic governing equations and couples them with implicit microscopic governing equations within the nonequilibrium region, resulting in high convergence efficiency in all flow regimes. To validate the efficiency and robustness of the IAUGKS, a series of numerical tests were conducted for high Mach number flows around diverse geometries. The current scheme can capture the nonequilibrium physics and provide accurate predictions of surface quantities. In comparison with the original UGKS, the velocity space adaptation, unstructured discrete velocity space, and implicit iteration significantly improve the efficiency by one or two orders of magnitude. Given its exceptional efficiency and accuracy, the IAUGKS serves as an effective tool for nonequilibrium flow simulations. [ABSTRACT FROM AUTHOR]

For understanding many real-world problems involving rarefied hypersonic, micro-, and nanoscale gas flows, the primary method may be the direct simulation Monte Carlo (DSMC). However, its computational cost is prohibitive in comparison with the Navier–Stokes–Fourier (NSF) solvers, eclipsing the advantages it provides, especially for situations where flow is in the near continuum regime or three-dimensional applications. This study presents an alternate computational method that bypasses this issue by taking advantage of data-driven modeling and nonlinear coupled constitutive relations. Instead of using numerical solutions of higher-order constitutive relations in conventional partial differential equation-based methods, we build compact constitutive relations in advance by applying deep neural network algorithms to available DSMC solution data and later combine them with the conventional finite volume method for the physical laws of conservation. The computational accuracy and cost of the methodology thus developed were tested on the shock wave inner structure problem, where high thermal non-equilibrium occurs due to rapid compression, for a range of Mach numbers from 2 to 10. The simulation results obtained with the computing time comparable to that of the NSF solver showed almost perfect agreement between the neural network-based combined finite volume method and DSMC and original DSMC solutions. We also present a topology of DSMC constitutive relations that allows us to study how the DSMC topology deviates from the NSF topology. Finally, several challenging issues that must be overcome to become a robust method for solving practical problems were discussed. [ABSTRACT FROM AUTHOR]

The uncertain factors in wind tunnel experiments, especially the interference effects, will cause the flow around the test model to differ from that under the free flow condition. The wind tunnel wall interference effects are challenging to be eliminated, especially in transonic flow regions where classical linear correction methods are ineffective. To address this issue, an efficient wind tunnel wall interference correction framework based on data assimilation is proposed for steady and shock buffet flows around the airfoils in the transonic region in this work. The ensemble Kalman filter (EnKF) is used to find the combination of inflow conditions, which minimizes the differences between the pressure coefficients measured in the wind tunnel experiment and those estimated by computational fluid dynamics (CFD). The polynomial chaos expansion method is used to propagate uncertainty in the forecast step of the EnKF in order to improve efficiency. Typical wind tunnel wall interference correction problems are studied, including two steady cases: the transonic flow around the Royal Aircraft Establishment 2822 (RAE2822) and the National Aerospace Laboratory 7301 (NLR7301) airfoils, and one unsteady case: the transonic shock buffet on the National Aeronautics and Space Administration supercritical (phase 2)-0714 (NASA SC(2)-0714) airfoil. It is demonstrated that with the corrected Mach number and angle of attack, the estimated CFD results are in better agreement with the measured experimental data than traditional linear methods and the computational efficiency is improved compared to the original EnKF. [ABSTRACT FROM AUTHOR]

This work aimed to investigate combustion mode conversions by rapid variation of the fuel flow rate in a cavity-based scramjet combustor. The experiments were carried out on a direct-connected facility with an inflow condition of Mach number 2.52, a total pressure of 1.35 MPa, and a total temperature of 1650 K. The fuel injector consisted of two injection ports: fuel was continuously injected from one port while the other controlled the fuel flow for mode conversions by switching it on or off. Simultaneous schlieren and CH* imaging techniques were used to characterize the dynamics of combustion mode conversions. It was recognized that the combustion modes characterized by different flow field structures and heat release distributions can be classified into three types: the shear-layer mode, the transition mode, and the jet-wake mode. During the combustion mode conversion, the mixing region of the transverse jet and air became thicker with the increase in fuel flow rate, and the gradient of the flow field density and the flame area increased, making the flame more likely to propagate upstream. The combustion suppression induced by rapid fuel addition was observed at low equivalence ratios. It was speculated that the weak heat supply was insufficient to provide adequate heat for the rapid ignition of the added fuel. Furthermore, it was found that the flame-flow matching process with frequent flame propagation upstream occurred during the combustion mode conversion. This process was attributed to the mismatch between the increasing heat release and the original flow field structure. [ABSTRACT FROM AUTHOR]

When a shock wave crosses a density interface, the Richtmyer–Meshkov instability causes perturbations to grow. Richtmyer–Meshkov instabilities arise from the deposition of vorticity from the misaligned density and pressure gradients at the shock front. In many engineering applications, microscopic surface roughness will grow into multi-mode perturbations, inducing mixing between the fluid on either side of an initial interface. Applications often have multiple interfaces, some of which are close enough to interact in the later stages of instability growth. In this study, we numerically investigate the mixing of a three-layer system with periodic zigzag (or chevron) interfaces, calculating the dependence of the width and mass of mixed material on properties such as the shock timing, chevron amplitude, multi-mode perturbation spectrum, density ratio, and shock mach number. The multi-mode case is also compared with a single-mode perturbation. The Flash hydrodynamic code is used to solve the Euler equations in three dimensions with adaptive grid refinement. Key results include a significant increase in mixed mass when changing from a single-mode to a multi-mode perturbation on one of the interfaces. The mixed width is mainly sensitive to the density ratio and chevron amplitude, whereas the mixed mass also depends on the multi-mode spectrum. Steeper initial perturbation spectra have lower mixed mass at early times but a greater mixed mass after the reflected shock transits back across the layer. [ABSTRACT FROM AUTHOR]

In this paper, the spatiotemporal evolution and the instability of the turbulence mixing layer in a supersonic ejector are deeply studied based on large eddy simulation. Time-averaged distributions and transient results are discussed based on mixed fluids. In particular, the general flow characteristics are identified through the examination of the spatiotemporal evolution of temperature, entropy, Mach number, and density gradient fields. A thin shear layer is gradually formed, and several ring vortices are effectively captured at the lip of the nozzle. Surprisingly, the annular vortex gradually evolves into a V-shaped vortex with time evolution, which reveals that the transition of turbulence occurs in the first half of the mixing chamber. It is found that the supersonic mixing layer exhibits slow growth initially, and a significant increase in velocity is effectively captured at the throat position based on time averaged velocity profiles. A similar trend is also evident in the transition region of the supersonic ejector throat. The mixing of motive flow and suction flow is effectively enhanced by a shock train generated within the diffusion chamber. A recirculation region is created, which leads to a narrowing of the potential core of the jet, and the symmetry of the flow is broken due to the shock location. It is further found that the instability of the main jet is the dominant factor in the fluid mixing process by studying the power spectrum of the pressure signal and characteristic frequency inside the mixing chamber. It further provides important physical insights into understanding the spatiotemporal evolution of internal turbulence and the mixing mechanism in the supersonic steam ejector. [ABSTRACT FROM AUTHOR]

The study utilizes a two-temperature model to analyze non-equilibrium in normal shocks within hypersonic flows in polyatomic rarefied gases. Derived from the extended second law of thermodynamics, this model separates translational and internal temperatures in polyatomic gases, providing a more accurate depiction of non-equilibrium gas flow compared to classical theories like the Navier–Stokes and Fourier (NSF) system. Notably, the analysis reveals that the two-temperature model incorporates an additional contribution to the heat flux due to the gradient of the dynamic temperature, resulting in improved accuracy, especially for high Mach numbers. Results show that the model gives satisfactory shock density and temperature profiles up to Mach 10, with very good agreement observed up to Mach 6.1 compared to the classical NSF model. We conduct an order of magnitude analysis on the dynamic temperature and heat flux gradients appearing in the new constitutive equation using the Mott-Smith method. This analysis highlights the impact of these terms on accurately modeling polyatomic gas behavior in high-speed flows. The effects of bulk viscosity and incoming temperature on shock profiles are also investigated, contributing to a better understanding of shock wave structures in polyatomic gases and their implications for hypersonic flow dynamics. [ABSTRACT FROM AUTHOR]

This article presents a numerical study on the 2D aerodynamic characteristics of an airfoil with a morphed camber. The operational regime of the main rotor blade of the IAR 330 PUMA helicopter was encompassed in CFD simulations, performed over an angle of attack range of α = [ − 3 ° ;   18 ° ] , and a Mach number of M = 0.38 . Various degrees of camber adjustment were smoothly implemented to the trailing-edge section of the NACA13112 airfoil, with a corresponding chord length of c = 600 mm at the Reynolds number, R e = 5.138 × 10 6 , and the resulting changes in static lift and drag were calculated. The study examines the critical parameters that affect the configuration of the morphing airfoil, particularly the length of the trailing edge morphing. This analysis demonstrates that increasing the morphed camber near the trailing edge enhances lift capability and indicates that the maximum lift of the airfoil depends on the morphed chord length. The suggested approach demonstrates potential and can be implemented across various categories of aerodynamic structures, such as propeller blade sections, tails, or wings. [ABSTRACT FROM AUTHOR]

The Mach number effect on the Richtmyer–Meshkov instability (RMI) evolution of the shocked V-shaped N 2 / SF 6 interface is numerically studied in this research. Four distinct Mach numbers are taken into consideration for this purpose: M s = 1.12 , 1.22 , 1.42 , and 1.62. A two-dimensional space of compressible two-component Euler equations is simulated using a high-order modal discontinuous Galerkin approach to computational simulations. The numerical results show good consistency when compared to the available experimental data. The computational results show that the RMI evolution in the shocked V-shaped N 2 / SF 6 interface is critically dependent on the Mach number. The flow field, interface deformation, intricate wave patterns, inward jet development, and vorticity generation are all strongly impacted by the shock Mach number. As the Mach number increases, the V-shaped interface deforms differently, and the distance between the Mach stem and the triple points varies depending on the Mach number. Compared to lower Mach numbers, higher ones produce larger rolled-up vortex chains. A thorough analysis of the Mach number effect identifies the factors that propel the creation of vorticity during the interaction phase. Moreover, kinetic energy and enstrophy both dramatically rise with increasing Mach number. Lastly, a detailed analysis is carried out to determine how the Mach number affects the temporal variations in the V-shaped interface's features. [ABSTRACT FROM AUTHOR]