Search

This paper presents a numerical study conducted to analyze the aerodynamic performance of supersonic missiles consisting of a cylindrical body and four flat-plate rear fins arranged uniformly, equipped with conical and ogive heads. Computational Fluid Dynamics (CFD) simulations were performed using the ANSYS Fluent 17.1 solver, along with the Gambit grid generation software. The objective was to compare the aerodynamic characteristics of these two head designs in terms of drag, lift, and stability at supersonic speeds. Various flow parameters, including Mach number and angle of attack, were investigated to comprehensively assess the performance of the missile configurations. The results indicate clear differences in the aerodynamic behavior of conical and ogive heads. Specifically, there was a 2–11 percent increase in the lift coefficient of the conical heads compared to the ogive heads, and an increase in the drag coefficient of both conical and ogive heads.

Abstract Aerodynamic test rigs are necessary for experimental testing of turbomachines and investigation of possible ways to improve machine performance. Existing installations give higher losses and do not work efficiently at off-design operating modes. An operating part with an adjustable radial diffuser was designed in order to determine the characteristics of turbine annular cascades. Experimental studies and computational verifications showed satisfactory results at various operating modes. The regulated backwall of the radial diffuser ensured supersonic velocity values behind the cascade and overall stable operation in a wide range of Mach numbers up to 1.3 when using compressors of comparatively small capacity. The optimum positions of the regulated backwall were determined, which provided a deep vacuum behind the cascade, as well as 1.5 times Mach number increase compared to the turbine cascade without a diffuser. Changing the inlet channel geometry at supersonic modes leads to an increase in the diffuser efficiency. Additionally, it was determined that the use of the turbine vane cascade in the test rig flow path is not necessary during calculation studies, but instead an axisymmetric vaneless converging area can be applied and give satisfactory results as well as reduce the time spent on calculations. The computational model can be used to optimize the design of an aerodynamic tunnel outlet area.

There is a complex transition phenomenon in the flow field of a typical hypersonic vehicle, which has a significant impact on the performance of the vehicle. The effects of Mach number and wall temperature on the transition of HyTRV were studied by numerical simulation methods. The self-developed software of the research group was used to carry out numerical calculations. The range of Mach number was 3~8, and the range of wall temperature was 150~900 K. Firstly, the hypersonic corrections of the γ-$\mathop R\limits^ \sim $eθt transition model and the SST turbulence model were carried out. The pressure gradient coefficient correction and the high-speed cross-flow correction were introduced into the γ-$\mathop R\limits^ \sim $eθt transition model, and the compressibility corrections of the closure coefficients β* and β of the SST turbulence model were carried out. Then, the grid independence verification was carried out, and the modified numerical method and software platform were confirmed by comparing with experimental results. Finally, the effects of Mach number and wall temperature on the transition law of the HyTRV boundary layer were studied. The results show that the transition area is mainly concentrated on both sides of the upper surface and the center line of the lower surface. With the increase of the incoming Mach number, the starting position of transition on the upper and lower surfaces is greatly backward, and the turbulent zone is greatly reduced, but it still exists. At the same time, the friction coefficient of the laminar flow zone on the upper surface increases continuously, and the friction coefficient of the turbulent zone on the lower surface decreases. As the wall temperature increases, the starting position of transition on the upper and lower surfaces shifts forward, then rapidly shifts backward, and finally the turbulent zone almost disappears.

In this paper, we present the results of the explosive dispersal of particles in high-speed environments. We carry out Euler–Lagrange numerical simulations of a source at quiescent ambient conditions as well as moving at Mach numbers of 3 and 6. Particle volume fractions of 0%, 1%, and 4.5% are presented. The detonation profile is computed with the Jones–Wilkins–Lee equation of state using a reactive burn model. Non-static cases provide a framework to consider the effect of a bow shock and pre-existing high-speed flow conditions on the dispersal process. We also compute averages of both static and dynamic pressures, as well as impulse density histories on virtual probe planes to characterize the momentum of the flow and particles that would deposit on a target. Results suggest that the presence of the particles can have a substantial effect on the pressure average of the virtual target planes. [ABSTRACT FROM AUTHOR]

We investigate the driving of MHD turbulence by gravitational contraction using simulations of an initially spherical, isothermal, magnetically supercritical molecular cloud core with transonic and trans-Alfvénic turbulence. We perform a Helmholtz decomposition of the velocity field, and investigate the evolution of its solenoidal and compressible parts, as well as of the velocity component along the gravitational acceleration vector, a proxy for the infall component of the velocity field. We find that (1) In spite of being supercritical, the core first contracts to a sheet perpendicular to the mean magnetic field, and the sheet itself collapses. (2) The solenoidal component of the turbulence remains at roughly its initial level throughout the simulation, while the compressible component increases continuously, implying that turbulence does not dissipate towards the centre of the core. (3) The distribution of simulation cells in the B– ρ plane occupies a wide triangular region at low densities, bounded below by the expected trend for fast MHD waves (B ∝ ρ, applicable for high-local Alfvénic Mach number M A) and above by the trend expected for slow waves (B ∼ constant, applicable for low local M A). At high densities, the distribution follows a single trend |$B \propto \rho ^{\gamma _{\rm eff}}$|⁠ , with 1/2 < γeff < 2/3, as expected for gravitational compression. (4) The mass-to-magnetic flux ratio λ increases with radius r due to the different scalings of the mass and magnetic flux with r. At a fixed radius, λ increases with time due to the accretion of material along field lines. (5) The solenoidal energy fraction is much smaller than the total turbulent component, indicating that the collapse drives the turbulence mainly compressibly, even in directions orthogonal to that of the collapse. [ABSTRACT FROM AUTHOR]

The ability of streamwise-travelling waves of spanwise velocity to reduce the turbulent skin-friction drag is assessed in the compressible regime. Direct numerical simulations are carried out to compare drag reduction in subsonic, transonic and supersonic channel flows. Compressibility improves the benefits of the travelling waves, in a way that depends on the control parameters: drag reduction becomes larger than the incompressible one for small frequencies and wavenumbers. However, the improvement depends on the specific procedure employed for comparison. When the Mach number is varied and, at the same time, wall friction is changed by the control, the bulk temperature in the flow can either evolve freely in time until the aerodynamic heating balances the heat flux at the walls, or be constrained such that a fixed percentage of kinetic energy is transformed into thermal energy. Physical arguments suggest that, in the present context, the latter approach should be preferred. This provides a test condition in which the wall-normal temperature profile more realistically mimics that in an external flow, and also leads to a much better scaling of the results, over both the Mach number and the control parameters. Under this comparison, drag reduction is only marginally improved by compressibility. [ABSTRACT FROM AUTHOR]

Richtmyer-Meshkov (RM) instability at a single-mode interface impacted by a cylindrical divergent shock with low to moderate Mach numbers is investigated experimentally. The motion of an unperturbed interface is first examined to obtain the background flow. The shocked interface moves uniformly at the early stage, but later decelerates. The stronger the incident shock, the larger the interface deceleration, which is reasonably predicted by a one-dimensional model considering the effect of postshock non-uniformity. Such a deceleration greatly inhibits the growths of harmonics of an initially perturbed interface and, consequently, the divergent RMinstability presents very weak nonlinearity from early to late stages. Particularly, higher-Mach-number cases present weaker nonlinearity due to larger deceleration there. This abnormal linear growth regime is reported for the first time. Benefiting from this, the incompressible linear model holds validity at all stages of divergent RM instability. It is also found that compressibility inhibits the initial growth rate, but produces a weak influence on the subsequent instability growth. [ABSTRACT FROM AUTHOR]

In this work, we employ well-established relations for compressible turbulent mean flows, including the velocity transformation and algebraic temperature-velocity (TV) relation, to systematically improve the algebraic Baldwin-Lomax (BL) wall model for high-speed zero-pressure-gradient air boundary layers. Any new functions or coefficients fitted by ourselves are avoided. Twelve published direct numerical simulation (DNS) datasets are employed for a priori inspiration and a posteriori examination, with Mach numbers up to 14 under adiabatic, cold and heated walls. The baseline BL model is the widely used one with semilocal scalings. Three targeted modifications are made. First, we employ a total-stress-based transformation (Griffin et al., Proc. Natl Acad. Sci. USA, vol. 118, issue 34, 2021, e2111144118) to the inner-layer eddy viscosity for improved scaling up to the logarithmic region. Second, we utilize the van Driest transformation in the outer layer based on the compressible defect velocity scaling. Third, considering the difficulty in modelling the rapidly varying and singular turbulent Prandtl number near the temperature peak in cold-wall cases, we design a two-layer strategy and use the TV relation to formulate the inner-layer temperature. Numerical results prove that the modifications take effect as designed. The prediction accuracy for mean streamwise velocity is notably improved for diabatic cases, especially in the logarithmic region. Moreover, a significant improvement in mean temperature is realized for both adiabatic and diabatic cases. The mean relative errors of temperature to DNS for all cases are down to 0.4%in the logarithmic wall-normal coordinate and 3.4% in the outer coordinate, around one-third of those in the baseline model. [ABSTRACT FROM AUTHOR]

This study investigates the impact of molecular thermal fluctuations on compressible decaying isotropic turbulence using the unified stochastic particle (USP) method, encompassing both two-dimensional (2-D) and three-dimensional (3-D) scenarios. The findings reveal that the turbulent spectra of velocity and thermodynamic variables follow the wavenumber (k) scaling law of k(d-1) for different spatial dimensions d within the high wavenumber range, indicating the impact of thermal fluctuations on small-scale turbulent statistics. With the application of Helmholtz decomposition, it is found that the thermal fluctuation spectra of solenoidal and compressible velocity components (us and uc) follow an energy ratio of 1: 1 for 2-D cases, while the ratio changes to 2: 1 for 3-D cases. Comparisons between 3-D turbulent spectra obtained through USP simulations and direct numerical simulations of the Navier-Stokes equations demonstrate that thermal fluctuations dominate the spectra at length scales comparable to the Kolmogorov length scale. Additionally, the effect of thermal fluctuations on the spectrum of uc is significantly influenced by variations in the turbulent Mach number. We further study the impact of thermal fluctuations on the predictability of turbulence. With initial differences caused by thermal fluctuations, different flow realizations display significant disparities in velocity and thermodynamic fields at larger scales after a certain period of time, which can be characterized by 'inverse error cascades'. Moreover, the results suggest a strong correlation between the predictabilities of thermodynamic fields and the predictability of uc. [ABSTRACT FROM AUTHOR]

For the problem of ventral support interference in a transonic-speed wind tunnel with the blended-wing-body aircraft NPU-BWB-300 installed, the numerical simulation method based on Reynolds-averaged Navier–Stokes (RANS) equations is used to study the influence law of aerodynamic characteristic interference with the variation of Mach numbers and angles of attack. Moreover, the characteristics of ventral support interference for blended-wing-body aircraft and conventional aircraft are compared. The relevant mechanism of the generation and change of ventral support interference is revealed by employing analysis of the body surface pressure, the shock wave of the strut, and the separation area between the strut and the aircraft. The aerodynamic characteristic interference obtained from the numerical simulation is linearized based on the principle of the least square method. Afterward, a numerical simulation correction method of ventral support interference in the transonic-speed wind tunnel for the blended-wing-body aircraft is developed. Finally, the test results after the corrections of ventral support interferences in the transonic-speed wind tunnel for NPU-BWB-300 are obtained, which is significant for the evaluation of current aerodynamic performances and subsequent optimization designs. [ABSTRACT FROM AUTHOR]

Introduction: The Kelvin-Helmholtz Instability (KHI) is an interface instability that develops between two fluids or plasmas flowing with a common shear layer. KHI occurs in astrophysical jets, solar atmosphere, solar flows, cometary tails, planetary magnetospheres. Two applications of interest, encompassing both space and fusion applications, drive this study: KHI formation at the outer flanks of the Earth's magnetosphere and KHI growth from non-uniform laser heating in magnetized direct-drive implosion experiments. Here, we study 2D KHI with or without a magnetic field parallel to the flow. We use both the GAMERA code, which solves the compressible Euler equations, and the STRATOSPEC code, which solves the Navier-Stokes equations under the Boussinesq approximation, coupled with the magnetic field dynamics. GAMERA is a global three-dimensional MHD code with high-order reconstruction in arbitrary nonorthogonal curvilinear coordinates, which is developed for a large range of astrophysical applications. STRATOSPEC is a three-dimensional pseudo- spectral code with an accuracy of infinite order (no numerical diffusion). Magnetized KHI is a canonical case for benchmarking hydrocode simulations with extended MHD options. Methods: An objective is to assess whether or not, and under which conditions, the incompressibility hypothesis allows to describe a dynamic compressible system. For comparing both codes, we reach the inviscid incompressible regime, by decreasing the Mach number in GAMERA, and viscosity and diffusion in STRATOSPEC. Here, we specifically investigate both single-mode and multi-mode initial perturbations, either with or without magnetic field parallel to the flow. The method relies on comparisons of the density fields, 1D profiles of physical quantities averaged along the flow direction, and scale-by-scale spectral densities. We also address the triggering, formation and damping of filamentary structures under varying Mach number or Atwood number, with or without a parallel magnetic field. Results: Comparisons show very satisfactory results between the two codes. The vortices dynamics is well reproduced, along with the breaking or damping of small-scale structures. We end with the extraction of growth rates of magnetized KHI from the compressible regime to the incompressible limit in the linear regime assessing the effects of compressibility under increasing magnetic field. Discussion: The observed differences between the two codes are explained either from diffusion or non-Boussinesq effects. [ABSTRACT FROM AUTHOR]

Context. The spatial distribution of dark matter halos carries cosmological and astrophysical information. Cosmological information can be considered to be contained in the connection between halo main properties and the large-scale halo bias, while the astrophysical information would be encoded in the scaling relations between halo properties. The combination of these two contributions leads to the effect of secondary halo bias. Aims. Our goal is to measure the signal of secondary halo bias as a function of a variety of intrinsic and environmental halo properties and to characterize its statistical significance as a function of cosmological redshift. Methods. Using fixed and paired N-body simulations of dark-matter halos – the UNIT simulation – with masses above ∼1011 M⊙h−1 identified over a wide range of cosmological redshifts (0 < z < 5), we explored the behavior of the scaling relations among different halo properties. We included novel environmental properties based on the halo distribution as well as the underlying dark-matter field. We implemented an object-by-object estimator of large-scale effective bias and tested its validity against standard approaches. With a bias assigned to each tracer, we performed a statistical analysis aimed at characterizing the distribution of the bias and the signal of the secondary halo bias. Results. We show how the halo scaling relations linking direct probes of the halo potential well do not depend on the environment. On the contrary, links between the halo mass and the so-called set of secondary halo properties are sensitive to the cosmological environment, mainly to under-dense regions. We show that the signal of secondary bias is derived statistically from secondary correlations beyond the standard link to the halo mass. Conclusions. We show that the secondary bias arises through nonlocal and/or environmental properties related either to the halo distribution or to the properties of the underlying dark-matter field. In particular, properties such as the tidal field (a measure of the anisotropy of the density field) and the local Mach number (a measure of the local kinetic temperature of the halo distribution) generate the signals of the secondary bias with the highest significance. We propose applications of the assignment of individual bias for the generation of mock catalogs containing the signal of secondary bias, as well as a series of cosmological analyses aimed at mining large galaxy datasets. [ABSTRACT FROM AUTHOR]

Recent observations have revealed detailed structures of radio relics across a wide range of frequencies. In this work, we performed three-dimensional magnetohydrodynamical (3D MHD) simulations of merger shocks propagating through a turbulent magnetized intracluster medium. We then employed on-the-fly Lagrangian particles to explore the physical processes behind the origination of radio substructures and their appearance in high and low-frequency observations. We employed two cosmic-ray (CR) electron acceleration models, with a fresh injection of electrons from the thermal pool and the re-acceleration of mildly relativistic electrons. We used the relative surface brightness fluctuations, δSν, to define a "degree of patchiness." First, we found that patchiness is produced if the shock's surface has a distribution of Mach numbers, rather than a single Mach number. Second, radio relics appear patchier if the Mach number distribution consists of a large percentage of low Mach numbers (ℳ ≲ 2.5). Furthermore, as the frequency increases, the patchiness also becomes larger. Nevertheless, if radio relics are patchy at high frequencies (e.g., 18.6 GHz), they necessarily will also be patchy at low frequencies (e.g., 150 MHz). Then, to produce noticeable differences in the patchiness at low and high frequencies, the shock front should have a Mach number spread of σℳ ≳ 0.3 − 0.4. Finally, the extent of the patchiness depends on the Mach number distribution as well as the CR acceleration model. We propose δSν as a potential tool for extracting merger shock properties and information on particle acceleration processes at shocks in radio observations. [ABSTRACT FROM AUTHOR]

Context. Hadronic interactions between cosmic rays (CRs) and interstellar gas have been probed in γ rays across the Galaxy. A fairly uniform CR distribution is observed up to a few hundred parsecs from the Sun, except in the Eridu cloud, which shows an unexplained 30–50% deficit in GeV to TeV CR flux. Aims. To explore the origin of this deficit, we studied the Reticulum cloud, which shares notable traits with Eridu: a comparable distance in the low-density region of the Local Valley and a filamentary structure of atomic hydrogen extending along a bundle of ordered magnetic-field lines that are steeply inclined to the Galactic plane. Methods. We measured the γ-ray emissivity per gas nucleon in the Reticulum cloud in the 0.16–63 GeV energy band using 14 years of Fermi-LAT data. We also derived interstellar properties that are important for CR propagation in both the Eridu and Reticulum clouds, at the same parsec scale. Results. The γ-ray emissivity in the Reticulum cloud is fully consistent with the average spectrum measured in the solar neighbourhood, but this emissivity, and therefore the CR flux, is 1.57 ± 0.09 times larger than in Eridu across the whole energy band. The difference cannot be attributed to uncertainties in gas mass. Nevertheless, we find that the two clouds are similar in many respects: both have magnetic-field strengths of a few micro-Gauss in the plane of the sky; both are in approximate equilibrium between magnetic and thermal pressures; they have similar turbulent velocities and sonic Mach numbers; and both show magnetic-field regularity with a dispersion in orientation lower than 10°–15° over large zones. The gas in Reticulum is colder and denser than in Eridu, but we find similar parallel diffusion coefficients around a few times 1028 cm2 s−1 in both clouds if CRs above 1 GV in rigidity diffuse on resonant, self-excited Alfvén waves that are damped by ion-neutral interactions. Conclusions. The loss of CRs in Eridu remains unexplained, but these two clouds provide important test cases to further study how magnetic turbulence, line tangling, and ion-neutral damping regulate CR diffusion in the dominant gas phase of the interstellar medium. [ABSTRACT FROM AUTHOR]

The primary objective of this paper is to study the detonation re-initiation through multiple reflections upon its entry into a confined space containing liquid fuel within a flow system, which holds significant implications for two-phase pulse detonation engines (PDE) and rotating detonation engines (RDE). Utilizing the Eulerian-Lagrangian method, this study investigates the disparities in the diffraction, quenching, and re-initiation processes of liquid n-heptane/air detonation upstream and downstream. Additionally, it explores the interaction between shock waves and droplets as well as determines the lean flammability limit for re-initiation within a flow system. The findings suggest that the incoming flow significantly contributes to the upstream decoupled detonation re-initiation, as it intensifies the leading shock, thereby expediting droplet fragmentation and evaporation. However, the compressive effect of the incoming flow is insufficient to initiate upstream detonation independently. To achieve re-initiation, it is still necessary for the detonation wave to be reflected by the wall. Moreover, the incoming flow expands the re-initiation boundaries, thereby facilitating the re-initiation of upstream detonation. As the equivalence ratio decreases, a higher inflow Mach number is required to achieve upstream re-initiation. • Upstream and downstream detonation re-initiation are compared. • Interaction between droplets and leading shock upstream and downstream are illustrated. • The detonation re-initiation boundaries beneath an incoming flow are found. [ABSTRACT FROM AUTHOR]

Gaps represent a common feature on the surface of supersonic vehicles, and their presence engenders localized thermal phenomena, including elevated heat flux and thermal invasion. These occurrences significantly imperil the aircraft's thermal stability and safety. This study employs numerical simulation to investigate supersonic streamwise gap flow and offers a comprehensive analysis of the impact of geometric and leakage on the distribution of wall heat flux. The computational results manifest a high alignment between regions of high heat flux and high Mach number. Consequently, these findings underscore the pivotal role of convective heat transfer governing the distribution of heat flux within the gap. Most notably, the sensitivity of wall heat flux is most pronounced with respect to gap width. A reduction in gap width yields a substantial reduction in wall heat flux, with this alteration primarily stemming from its impact on the shear interactions between the supersonic mainstream and the vortices within the gaps. Furthermore, changes in gap length influence the invasion characteristics of the supersonic mainstream, thereby instigating consequential shifts in the distribution of heat flux density along the inner surfaces of the gaps. In contrast, variations in gap depth exert a comparatively modest impact on the distribution of heat flux density. The introduction of leakage effects at the base of the gaps results in a significant intensification of supersonic mainstream penetration and concomitant convective heat transfer in the bottom exit region. This phenomenon engenders localized zones characterized by elevated heat flux density, thus severely compromising the aircraft's thermal sealing performance. • Thermal environments within streamwise gaps in supersonic flow are investigated. • Influences of geometries on the distribution of wall heat flux are analyzed. • Mechanisms by which changes in vortex affect the thermal environment are revealed. • Effects of potential bottom leakage on thermal invasion are evaluated. [ABSTRACT FROM AUTHOR]

In a gaseous medium, dynamical friction (DF) reaches a maximum when the orbital speed of a (point-like) perturber moving on a circular orbit is close to the sound speed. Therefore, in a quasi-steady state, eccentric orbits of perturbers approaching the sound barrier (from below) should rapidly circularize as they experience the strongest drag at pericentre passage. To investigate this effect, we extend the solution for circular DF in a uniform gaseous medium to eccentric Keplerian orbits. We derive an approximation to the steady-state DF force, which is valid for eccentricities as high as e  = 0.9 in a limited range of Mach number around the transition to supersonic regime. We validate our analytical result with 3D simulations of the gas density response. Although gaseous DF generally dissipates orbital energy, we find that it can be directed along the motion of the perturber near pericentre passage when the eccentricity is e ≳ 0.9. We apply our results to compute the long-time evolution of the orbital parameters. Most trajectories tend to circularize as the perturber moves into the supersonic regime. However, orbits with eccentricities e ≳ 0.8 below the sound barrier experience a slight increase in eccentricity as they loose orbital energy. Possible extensions to our analytical approach are also discussed. [ABSTRACT FROM AUTHOR]

Accurate prediction of Mach numbers is of paramount importance in wind tunnel experiments. While ensemble prediction models have shown promising performance compared with single-model approaches, existing ensemble methods often prioritize ensemble generation while neglecting the critical aspect of ensemble fusion. This oversight has led to suboptimal precision in Mach number prediction. In this paper, we address this issue by proposing a novel approach that leverages metalearning for ensemble fusion. By incorporating metalearning techniques, our framework effectively combines the strengths of multiple base learners to enhance predictive accuracy. Extensive experiments conducted on a comprehensive data set demonstrate the superiority of our proposed approach over traditional ensemble methods. The results highlight the significance of ensemble fusion in achieving high-precision Mach number prediction in wind tunnel experiments. This research contributes to the advancement of ensemble modeling by emphasizing the importance of fusion techniques and provides valuable insights for improving ensemble prediction models in various domains. [ABSTRACT FROM AUTHOR]

Numerical simulation of an underexpanded rectangular sonic jet controlled with a pair of normally injected supersonic minijets on mixing promotion was carried out using a commercial computational fluid dynamics (CFD) solver. Results from three-dimensional (3D) Reynold's-averaged Navier Stokes simulations exhibited very good agreement with experimental results for all the cases studied. Detailed analysis was carried out to quantify the influence of mass flow rate ratio and orientation of the air-tab injected along minor axis and diagonal axis, and pressure gradient at nozzle exit on the jet mixing. Centerline Mach number decay, density gradient contours, generation and propagation of streamwise vortices, axis switching, average shear layer thickness, and mass entrainment are used to quantify the jet mixing. It was found that fluidic injection along diagonal axis generates asymmetrical counter-rotating streamwise vortices causing rapid jet spread along the minor axis, resulting in axis switching and enhanced mass entrainment, leading to a shorter core and rapid characteristic decay. Although injection along the minor axis is capable of reducing shock strength, the decay caused by it is only marginal. Effectiveness of the fluidic injection diminishes with increases in the favorable pressure gradient. A maximum of about 53% reduction in jet core length was caused by injection along diagonal (IAD), corresponding to a mass flow rate ratio of 4%. [ABSTRACT FROM AUTHOR]

Supersonic flow over a half-angle wedge (θ = 15°) with an upstream Mach number of 2.0 was investigated using 2D Euler equations where sea level conditions were considered. The investigation employed the Steger–Warming flux vector splitting (FVS) method executed in MATLAB 9.13.0 (R2022b) software. The study involved a meticulous comparison between theoretical calculations and numerical results. Particularly, the research emphasized the angle of oblique shock and downstream flow properties. A substantial iteration count of 2000 iteratively refined the outcomes, underscoring the role of advanced computational resources. Validation and comparative assessment were conducted to elucidate the superiority of the Steger–Warming flux vector splitting (FVS) scheme over existing methodologies. This research serves as a link between theoretical rigor and practical applications in high-speed aerospace design, enhancing the efficiency of aircraft components. [ABSTRACT FROM AUTHOR]

Purpose: This study aims to study the interaction effects between a rectangular supersonic jet with a flat wall computationally using wall length as a parameter. The purpose of this study is to investigate the effect of change in wall length on supersonic core length (SCL) reduction, jet deflection and jet decay behavior. Design/methodology/approach: The design Mach number and aspect ratio at the rectangular exit were 1.8 and 2, respectively. To study the wall length effects on jet-wall interactions, wall length (Lw) was varied as 0.5Dh, 1Dh, 2Dh, 4Dh and 8Dh, where Dh was the hydraulic diameter of the nozzle exit. The flat wall with the matching width of the rectangular exit section of a supersonic nozzle was placed at the nozzle exit such that the supersonic jet grazed past the wall. The studies were carried out at over-expansion [nozzle pressure ratio (NPR) = 4], near optimum expansion (NPR = 6) and under-expansion (NPR = 8) levels. Findings: Results indicated that significant reduction in wall-bounded SCL was noticed in the range of 0.5Dh ≤ Lw ≤ 1Dh for both over-expansion and under-expansion conditions. At Lw ≥ 4Dh, SCL got enhanced at NPR = 4 and 6 but had a negligible effect at NPR = 8. Practical implications: Thrust vector control, noise reduction and easy take-off for high-speed aircraft. Originality/value: The effect of change in flat wall length on interaction characteristics of a rectangular supersonic jet was not studied before in terms of SCL reduction and jet decay behavior. [ABSTRACT FROM AUTHOR]

Thermal boundary conditions of the turbine disk cavity system are of great importance in the design of secondary air systems in aero-engines. This study aims to investigate the complex heat transfer mechanisms of a rotating turbine disk under high-speed conditions. A high-speed rotating free-disk model with Dorfman empirical solutions is developed to evaluate the heat transfer performance considering various factors. Specifically, the influence of compressibility, variable properties, and heat dissipation is determined using theoretical and numerical analyses. In particular, a novel combined solution method is proposed to simplify the complex heat transfer problem. The results indicate that the heat transfer performance of a free disk is primarily influenced by the rotating Mach number, rotating Reynolds number, Rossby number, and wall temperature ratio. The heat transfer temperature and Nusselt number of the free disk are strongly correlated with the rotating Mach number and rotating Reynolds number. Analysis reveals that heat dissipation is a critical factor affecting the accurate evaluation of the heat transfer performance of the turbine disk. Thus, the combined solution method can serve as a reference for future investigations of flow and heat transfer in high-speed rotating turbine disk cavity systems in aero-engines. [ABSTRACT FROM AUTHOR]

This paper reports a numerical study of a novel methodology for passive suppression of deep cavity noise by means of strategically designed and arrangements of multiple elastic panels and examines its underlying aeroacoustic–structural interaction physics. The study is conducted with a freestream, at Mach number 0.09 and Reynolds number of 4 × 104 based on the cavity length, past a two-dimensional cavity by means of direct aeroacoustic simulation coupled with a panel dynamic solver in monolithic fashion. For each cavity-panel configuration, the fluid-loaded panel natural frequencies are harmonized with the characteristic aeroacoustic processes of the original/modified cavity aeroacoustic feedback loop. This promotes panel aeroacoustic-structural resonance for absorption of feedback flow and acoustic fluctuation energy for achieving less eventual cavity noise. The most effective configuration gives a remarkable noise power reduction by 15 dB from a rigid cavity. Inadvertently, it reduces cavity drag by almost 15%. Simultaneous reduction of both cavity noise and drag is unprecedented among similar attempts in the literature. In-depth spatiotemporal analyses of aeroacoustic–structural interaction results elucidate the intricate interplay between cavity flow, panel vibration responses, and cavity acoustic modes, leading to noise reduction in all cavity-panel configurations studied. Essentially, the vertical panel acts to curtail the efficacy of coupling between growing shear layer and cavity acoustic modes whose sustenance is further impeded by an acoustically induced resonant panel at the cavity bottom. The proposed methodology is confirmed to be feasible yet effective, which holds great potential for fluid-moving applications in which a quiet and energy-efficient cavity configuration is desired. [ABSTRACT FROM AUTHOR]

The impact of porous boundaries on skin friction and heat transfer in compressible turbulent channel flows is explored through direct numerical simulations. Utilizing the volume-averaged Navier–Stokes equations for the porous media, we examine flows at Mach numbers M b = 0.3 and M b = 1.5 , with friction Reynolds numbers ranging from 176 to 220, across three permeability levels: 0 (impermeable), 1.6 × 10 − 5 , and 1.0 × 10 − 4 . Our findings reveal a dual impact of porous boundaries on skin friction: a direct reduction due to averaged slip velocity at the interface and an indirect increase from amplified Reynolds shear stress due to enhanced wall-normal permeability. This permeability weakens the wall-blocking effect, leading to intensified wall-normal velocity fluctuations and consequently higher Reynolds stress. Additionally, the Mach number significantly influences flow structures, particularly at higher permeabilities, where it disrupts near-wall streaks at M b = 0.3 but has a diminished effect at M b = 1.5. Thermal analysis confirms the validity of the generalized Reynolds analogy for the mean temperature in these flows, with the turbulent Prandtl number approximating 1, in line with the strong Reynolds analogy. Moreover, a strong spatial correlation between heat flux and wall shear stress fluctuations underscores the intertwined dynamics of momentum and thermal transport at the porous–fluid interface. [ABSTRACT FROM AUTHOR]

Shockwave boundary layer interaction is studied for oblique shocks generated from a shock generator with a varying flow deflection angle. A two-dimensional shock generator with a wedge-shaped geometry is fixed inside the test section and mounted to the wind tunnel's top wall. The shock generator is made to pitch up and down so that the net flow deflection angle changes because of the pitch motion. As a result, the shock strength changes for a varying flow deflection angle of the shock generator with a fixed incoming Mach number of the flow. The imposed adverse pressure gradient on the boundary layer by this incident shock changes, which influences the shape and structure of the shockwave boundary layer interaction. Three different shockwave boundary layer interaction structures are observed with high-speed schlieren from the experiments conducted for the current study. Schematics for these structures' types are illustrated, and a model is presented to build the transition criteria for the flow deflection angles for which the transition would take place from one structure to the other. The two transition criteria, one for type 1 to type 2 transition and the other for type 2 to type 3 transition, are validated with the observed results from the experiments. Assumptions made for the transition criteria are also validated with experimental results. [ABSTRACT FROM AUTHOR]

The current work is undertaken to examine the capability of a recently developed ICFM (internal conical flow M) basic flowfield [Musa et al., AIAA J. 61, 1–16 (2023)] to design hypersonic internal waverider intakes. The osculating axisymmetric flows concept is employed alongside the streamline-tracing method to design three internal waverider intakes with different contraction ratios, i.e., 6.0, 9.0, and 25. The three intakes have similar capture and throat shapes and entrance areas. Then, the starting characteristics of designed intakes were investigated at different design and off-design conditions. The selected design condition is the flight at Mach 6.0 and 28 km altitude. Viscous computations have been performed at Mach 6.0, 5.0, 4.0, 3.9, 3.8, 3.5, and 3.0 for the intake with a contraction ratio of 6.0, and at Mach 6.0, 5.0, 4.7, and 4.0 for the intake with contraction ratio of 9.0. The intake with a contraction ratio of 25 is simulated at Mach 6.0 and 5.0. Successful intake starting has been achieved up to Mach 3.8 for the first intake and up to Mach 4.7 for the second intake. The intake with a contraction ratio of 25 attained an unstart state in the design condition due to a very high contraction ratio. Thus, the starting limit of internal waverider intakes designed using ICFM basic flowfield is found to be controlled by Van Wie empirical limit. It is concluded that the ICFM basic flowfield can effectively design high-performance fixed-geometry intakes with overboard spillage, demonstrating excellent performance and startability across a wide range of Mach numbers. [ABSTRACT FROM AUTHOR]

Turbulent kinetic energy (TKE) is a quantity of primary importance in shock–turbulence interaction (STI). Linear interaction analysis (LIA) serves as a theoretical method to predict the amplification of TKE in STI. LIA is constrained by low-amplitude fluctuations and neglects nonlinear effects. In this paper, we explore the nonlinear amplification of the kinetic energy of fluctuations in the interaction between a vorticity wave and a shock wave that serves as a building block flow for STI. A weakly nonlinear framework (WNLF) is introduced to analyze nonlinear effects in fluctuation kinetic energy (FKE) and identify the dominant physical mechanisms driving its amplification. The theoretical framework is validated through high-accuracy numerical simulations of shock–vorticity wave interactions. The simulation results are compared with the predictions derived from the WNLF for a range of intensities and inclinations of shock-upstream vorticity fluctuations at different Mach numbers, and WNLF is found to successfully scale the numerical data for FKE, thus confirming the validity and applicability of the framework. According to WNLF findings, at lower supersonic Mach numbers, the intermodal interaction between vorticity–vorticity modes is important, whereas the interaction between vorticity and acoustic modes becomes dominant at higher Mach numbers. Using the intermodal interactions, a model based on WNLF is proposed to predict TKE amplification in STI. In comparison to direct numerical simulation, the WNLF based model predicts the TKE amplification for moderate turbulent Mach number and lower supersonic flow Mach numbers. This is a significant improvement over the LIA results available in the literature. [ABSTRACT FROM AUTHOR]

This work presents a comparison of several high-order numerical methodologies for simulating shock/turbulence interactions based on the supersonic Taylor–Green vortex flow, considering a Reynolds number of 1600 and a Mach number of 1.25. The numerical schemes considered include high-order Finite Difference, Targeted Essentially Non-Oscillatory, Discontinuous Galerkin, and Spectral Difference schemes. The shock capturing methods include high-order filtering, localized artificial diffusivity, non-oscillatory numerical fluxes, and local low-order switching. The ability of the various high-order numerical methodologies to both capture shocks and represent accurately the development of turbulent vortices is assessed. [ABSTRACT FROM AUTHOR]

A ducted rotor system was used to produce turbulent jets with a Reynolds number up to 5.97 × 105 and Mach number of 0.222 based on mean streamwise velocity. Three rotors with a diameter of 11.8 cm were manufactured and tested inside a duct with a 1 mm tip clearance at a speed up to 30 000 revolutions per minute (rpm). All rotor blades contain the same aspect ratio of 2.2, a National Advisory Committee for Aeronautics 2410 airfoil, and ideal pitch distribution. However, three different blade planform shapes were used including a rectangular shape with constant chord, trapezoidal shape with a taper ratio of 0.5, and elliptical shape where the trailing edge of the blade is expressed with an elliptical function. The rotor thrust and electric power were measured, and the thrust coefficient and figure of merit was computed. The flow-field produced by the ducted rotors was measured in the near-field using laser Doppler velocimetry techniques. The inflow velocity approximately 3 mm upstream of the rotor blade leading edge was acquired and its significance on blade aerodynamics and performance is analyzed. Time-averaged contours of cross-stream vorticity reveal intense hub and blade tip vortex structures, which are impacted by the shape of the blade, particularly in the blade tip region. Tip vorticity as well as streamwise turbulence intensity and turbulent kinetic energy in this region were mitigated for the rotors with trapezoidal and elliptical blades. However, the turbulent structure of the jet produced by all three rotor blade shapes showed similarity at a mere 2.8 rotor diameters downstream of the rotor. This finding emphasizes the importance of blade design on the near-field dynamics of ducted rotor flows for aircraft propulsion. [ABSTRACT FROM AUTHOR]

The dual-mode combustor model of a kerosene-fueled dual-mode scramjet engine was investigated through a series of combustion experiments with different mode on the resistance heating direct-connected test system at Northwestern Polytechnical University. The experiment simulated the actual combustion process under the conditions of an inlet total pressure ranging from 600 to 900 kPa, an inlet total temperature of 810 K, and an inlet Mach number of 2.0. The results show that as the equivalence ratio of kerosene increases, the combustor goes through three modes sequentially: pure scramjet mode, dual-mode scramjet mode, and dual-mode ramjet mode. The peak pressure ratio of the combustor corresponding to the combustion mode transition boundary is 1.95 and 3.5, respectively, and the Mach number at the exit of the isolator is 2.26 and 1.0, respectively. The mode transition boundary remains unchanged regardless of the inlet total pressure. When the position of the peak pressure point is stably located near the upstream cavity, the combustor operates in dual-mode ramjet mode, while when the position of the peak pressure point is stably located near the downstream cavity, the combustor operates in dual-mode scramjet mode. As the position of the measurement point gets closer to the downstream, the relative coefficient of wall pressure change becomes higher in terms of wall pressure measurement. The greatest variation in the relative coefficient of wall pressure occurs at x = 370 mm, indicating the highest sensitivity. It is the most suitable parameter for controlling the fuel flow of the dual-mode scramjet. [ABSTRACT FROM AUTHOR]

This paper describes direct numerical simulations of a shock-wave/turbulent boundary-layer interaction (STBLI) process in a compression-ramp flow with a ramp angle of 24° and a free-stream Mach number of M a ∞ = 2.9. Spectral analysis, two-point cross correlation, convection velocity statistics, and individual vortex identification are used to elucidate the streamwise variation of multiscale turbulent structures in the STBLI process. Typical Lagrangian coherent structures in the turbulent boundary layer before the STBLI region are characterized as hairpin-like vortical structures, with heads that rise together with the separated mean flow in the STBLI region. In the downstream region, the reattached turbulent boundary layer has a two-layer structure. The outer layer is characterized as an intensification of large-scale velocity structures, which is attributed to the shock-wave-induced compression effect on vortical structures. A viscous-dominated layer develops independently in the vicinity of the wall, leading to a gradual restoration of the wall-shear effect that accumulates the inner-layer dynamics of small-to-moderate-scale turbulent motions. [ABSTRACT FROM AUTHOR]

Gas–solid coupling systems operating at low pressure or the micro/nanoscale generally exist in nature and industrial manufacture. Although the gas-scattering model has been widely used to study this problem on the dust surface, the consideration of gas physisorption was often neglected in previous applications of gas–surface scattering models. Therefore, this study aims to investigate the distribution of gas physisorption on the dust surface and assess its impact on the static force experienced by nonspherical dust in free-molecule flows. In this study, the prolate dust spinning around its minor axis is considered and the in-house direct simulation Monte Carlo code is used. Results show that gas physisorption on prolate dust is influenced by changes in gas number densities, Mach number, and dust shape. Furthermore, the gas physisorption enhances the gas–dust coupling for dust with a smooth surface at low gas pressure, attributed to the increasing ratio of Maxwell diffuse scattering of gas molecules on the gas-adsorbed part of the surface. Hence, gas physisorption was suggested as a potential factor for gas–dust coupling at low gas pressure. [ABSTRACT FROM AUTHOR]

This study explores the efficiency of employing a particle-spray cloud to mitigate shock wave propagation, which is essential in various industrial applications, especially in preventing potential hydrogen explosions within nuclear reactor containment buildings. Numerical simulations, primarily in one- and two-dimensional configurations, are utilized to examine the interaction between shock waves and a cloud of polydisperse particles, considering both air and hydrogen–air mixtures as carrier gases. A novel reduced-order theoretical model is developed to analyze the dispersion pattern of polydisperse particles, with validation conducted through direct numerical simulations. Results demonstrate that the polydispersion of cloud particles significantly reduces shock wave propagation compared to monodisperse particles. Notably, particles with smaller diameters and higher standard deviations (σ) show increased attenuation effects. Additionally, scenarios with higher particle volume fractions ( τ v , 0 ) contribute to enhanced shock wave attenuation. A critical incident Mach number is identified, indicating a significant change in shock wave transmission from supersonic to subsonic when M s < 2.8. [ABSTRACT FROM AUTHOR]

The drag force experienced by astronomical objects moving through gaseous media (gas dynamical friction) plays a crucial role in their orbital evolution. Ostriker derived a formula for gas dynamical friction by linear analysis, and its validity has been confirmed through subsequent numerical simulations. However, the effect of gas accretion onto the objects on the dynamical friction is yet to be understood. In this study, we investigate the Mach number dependence of dynamical friction considering gas accretion through three-dimensional nested-grid simulations. We find that the net frictional force, determined by the sum of the gravitational force exerted by surrounding gas and momentum flux transferred by accreting gas, is independent of the resolution of simulations. Only the gas outside the Bondi–Hoyle–Lyttleton radius contributes to dynamical friction, because the gas inside this radius is eventually absorbed by the central object and returns the momentum obtained through the gravitational interaction with it. In the subsonic case, the front–back asymmetry induced by gas accretion leads to larger dynamical friction than predicted by the linear theory. Conversely, in the slightly supersonic case with a Mach number between 1 and 1.5, the nonlinear effect leads to a modification of the density distribution in a way that reduces the dynamical friction, compared with the linear theory. At a higher Mach number, the modification becomes insignificant and the dynamical friction can be estimated with the linear theory. We also provide a fitting formula for dynamical friction based on our simulations, which can be used in a variety of applications. [ABSTRACT FROM AUTHOR]

During the flight of hypersonic vehicle, air will be decomposed and ionized due to the "friction" under ultra-high speed, thus forming a plasma layer. Because the plasma has the ability to absorb and reflect electromagnetic waves, the "black barrier" phenomenon is formed. In addition, when the hypersonic vehicle passes through the atmosphere, the surface temperature rises sharply due to aerodynamic heating, and the surface material undergoes a series of complex changes to form ablation. In this paper, the finite volume method and the laminar finite rate model are used to study the flow field velocity, pressure distribution, flow field temperature, and spatial distribution of each component of the aircraft at different Mach numbers, angles of attack, and heights. In the flow field of supersonic aircraft, N and O are mainly concentrated in the tail of the aircraft, NO is mainly concentrated in the head of the aircraft, and N2 and O2 are full of the whole space. Because of the accumulation of NO+ and O2+ in the tail of the aircraft, the charge accumulation is formed, which will further interfere with the electromagnetic wave signal. The mass fraction of N and O increases with the increase in Mach numbers, while the mass fraction of O2 decreases with the increase in Mach numbers. Different angles of attack will affect the asymmetry of the shock wave of the aircraft. In this paper, the evolution mechanism and distribution characteristics of aircraft plasma are revealed, which lay a theoretical foundation for solving the problem of black barrier and ablation. [ABSTRACT FROM AUTHOR]

Mixing characteristics of jet emerging from a subsonic nozzle exit has been experimented and the results are compared with uncontrolled jet and controlled jet configurations. The mixing enhancement was achieved using a passive method of jet control in which tandem tabs arrangement with rectangular cross section are fixed at the nozzle exit. Two Tab configurations, the Tandem tab (TT) and Stepped Tandem Tab (STT) are used to enhance the mixing characteristics of the jet, the aspect ratio (length/width) of the tabs was 1.67 offering a blockage ratio of 9.55% to the nozzle exit. The blockage ratio of TT and STT configurations are maintained to be equal so that the mixing characteristics can be compared. The axial and radial jet spread are compared for nozzle exit Mach numbers of 0.6, 0.8 and 1.0. The TT controlled jet offered a potential core reduction of 63%, 78% and 82% for Mach numbers 0.6, 0.8 and 1.0 respectively. The STT controlled jet offered a potential core reduction of 89%, 90% and 85% for Mach numbers 0.6, 0.8 and 1.0 respectively. The radial spread of uncontrolled jet, controlled jet with TT and STT are plotted at several X/D locations and found that the controlled jets have more jet spread in both radial directions. A simulation is conducted for jets with exit Mach number 0.8 and the results are validated with the experimental findings. Based on the preliminary experimentation and computation, the STT controlled jet achieved better jet mixing through more potential core reduction and radial spread characteristics as compared to the TT configuration and base nozzle. [ABSTRACT FROM AUTHOR]

Numerical investigations of high-speed rarefied gas outflow into a vacuum through channels with a forward- or backward-facing step have been conducted using the direct simulation Monte Carlo method. Calculations have been performed for various free-stream Mach numbers, covering transonic, supersonic, and hypersonic flow regimes, and over a wide range of gas rarefaction from free molecular to near hydrodynamic conditions. Mass flow rates through the channel and the gas flow field have been accurately calculated both inside the channel and in the regions upstream and downstream. It has been established that channel geometry, the free-stream velocity, and gas rarefaction strongly influence the gas flow. In the flow field, in front of the channel, a phenomenon known as a detached shock occurs, while inside the channel, a gas recirculation zone may form. [ABSTRACT FROM AUTHOR]