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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]

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]

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]

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]

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]

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 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 simultaneous flow past two rectangular tandem wings is considered. The heat loads on the main wing are studied as functions of the intensity of the vortex formations produced by a generator. The following flow regimes are considered; the freestream Mach number М∞ = 3 and the attack angles of the generator are 10° and 20°. It is shown that under certain conditions the surface temperature of the main wing can be considerably reduced due to its interaction with vortex structures. [ABSTRACT FROM AUTHOR]

All physics issues are expected to be solved with a concise expression, and the field of aero-optical effect is no exception. Due to the excessive influencing factors and coefficients that are difficult to determine, there are only a few available conclusions at present. Moreover, the validation of these conclusions is often inadequate due to their limited scope of applications. In this paper, the Reynolds averaged Navier–Stokes equations are used to simulate the flow field of a hemisphere-on-cylinder turret with various Mach numbers equal to 0.3, 0.4, and 0.5. The wavefront distortion of the beam is calculated using the ray-tracing method and Zernike fitting. The similarity law proposed by the predecessors is verified, and the variation trend of OPDsteady/[(ρ∞/ρSL)M2D] with window angle is studied for the first time. It is found that the functional relation between OPDsteady/[(ρ∞/ρSL)M2D] and the window angle is approximately quadratic. Based on the Prandtl–Glauert compressibility correction principle, the similarity law proposed by predecessors is improved. [ABSTRACT FROM AUTHOR]

Exodus 14:2 says that God parted the Red Sea with a strong east wind. If God complied with the natural laws, a compressible flow assessment can be applied. A notional scenario using stagnated flow was proposed and the penetration depth into the water column was calculated. Additionally, the temperature of the flow increases dramatically with increasing Mach number. For subsonic flow, the penetration depth is limited to less than 30 feet and the air temperature increase can be over 100°F. Hence, it seems the Red Sea was a bona fide miracle, and not merely a miracle of providential timing. [ABSTRACT FROM AUTHOR]

Pitot pressure is the most common measurement in high total enthalpy shock tunnels for test condition verification. Nozzle calculations using multi-temperature non-equilibrium thermochemistry are needed in conjunction with Pitot measurements to quantify freestream properties. Pitot pressure is typically matched by tuning the boundary layer transition location in these simulations. However, non-equilibrium thermochemistry effects on the Pitot probe are commonly ignored. A computational study was undertaken to estimate the effect of non-equilibrium thermochemistry on Pitot pressure and freestream conditions. The test flow was produced by a Mach 7 nozzle in a reflected shock tunnel for air at a relatively low total enthalpy of 2.67 MJ/kg. Three different thermochemical models (equilibrium, finite-rate chemistry and two-temperature thermochemistry) were employed to compute flow variables at the nozzle exit and Pitot probe. Pitot pressures from these simulations were compared against those obtained via experiments. The results show a departure from the commonly utilized C of 0.92 in the reduced Rayleigh-Pitot equation form for high Mach numbers. Additionally, calculations were done with a sweep of free-stream conditions and resulting in values for one- and two-temperature models to use in future shock tunnel studies. Overall, our results show that the influence of finite-rate thermochemistry should be taken into account, even at relatively low flow enthalpies. • Simulations show change from the mainly used C = 0.92 in the Rayleigh-Pitot equation. • Equilibrium simulations with C = 0.92 cause higher errors in end flow properties. • Influence of finite-rate thermochemistry should be always taken into account. • A sweep of free-stream conditions varying only enthalpy resulted in similar trends for two thermochemical models used. • Recommend C values estimated for both 1T and 2T models. [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]

To analyze the noise induced by moving rigid structures in low Mach number flows, acoustic governing equations based on the viscous/acoustic splitting method and the arbitrary Lagrangian–Eulerian method are rigorously derived. In order to resolve the numerical instability generated in a non-uniform mean flow, the modified viscous/acoustic method, based on the filtering method, is developed. The acoustic equations are transformed into the same form as the incompressible flow equations by introducing the acoustic co-velocity and solved based on a collocated grid finite volume method. An approach for solving acoustic equation based on the PIMPLE algorithm is presented and computed in open-source computational fluid dynamics software OpenFOAM, which brings down communication costs and speeds up computing efficiency. Furthermore, the source term decomposition is extended to study the noise generated by each source term in a motion grid. Several examples including stationary and moving meshes have been designed to prove the accuracy of this approach. Finally, the aerodynamic and acoustic properties for the flow past a transversely oscillating cylinder at Re = 200, Ma = 0.2 in lock-in and non-lock-in regions is present. [ABSTRACT FROM AUTHOR]

Flexible launch vehicle modes contribute to the deterioration of its stability performance. For this reason, their influence must be considered when designing and validating appropriate solutions for its automatic control system. The first three bending modes are the most relevant for control system design and analysis. Next, considerations regarding the elastic modes of the Vega actuator, used in the simulations performed throughout this study, are presented. The IMU (Inertial Measurement Unit) measurements providing attitude angles and angular rates are strongly corrupted by the local elastic deformations resulting from structural flexibility. These measurements are fed back to the TVC (Thrust Vector Control) system which controls the pitch and the yaw motion of the vehicle. The design methodology for this control system is based on an ℒ1 adaptive technique incorporating a Butterworth filter guarantees good stability performances for wide variations of the launcher’s dynamics parameters. [ABSTRACT FROM AUTHOR]

This paper investigates numerically the shock wave interaction with a V-shaped heavy/light interface. For numerical simulations, we choose six distinct vertex angles ( θ = 40 ∘ , 60 ∘ , 90 ∘ , 120 ∘ , 150 ∘ , and 170 ∘) , five distinct shock wave strengths ( M s = 1.12 , 1.22 , 1.30 , 1.60 , and 2.0 ), and three different Atwood numbers ( A t = − 0.32 , − 0.77 , and − 0.87 ). A two-dimensional space of compressible two-component Euler equations are solved using a third-order modal discontinuous Galerkin approach for the simulations. The present findings demonstrate that the vertex angle has a crucial influence on the shock wave interaction with the V-shaped heavy/light interface. The vertex angle significantly affects the flow field, interface deformation, wave patterns, spike generation, and vorticity production. As the vertex angle decreases, the vorticity production becomes more dominant. A thorough analysis of the vertex angle effect identifies the factors that propel the creation of vorticity during the interaction phase. Notably, smaller vertex angles lead to stronger vorticity generation due to a steeper density gradient, while larger angles result in weaker, more dispersed vorticity and a less complex interaction. Moreover, kinetic energy and enstrophy both dramatically rise with decreasing vortex angles. A detailed analysis is also carried out to analyze the vertex angle effects on the temporal variations of interface features. Finally, the impacts of different Mach and Atwood numbers on the V-shaped interface are briefly presented. [ABSTRACT FROM AUTHOR]

In this work, we study the bending vibrations of a thin wing for a profile with two axes of symmetry, taking into account the influence of torsional vibrations on them. The primary focus of the current study is on finding an analytic solution for the wing-bending equation through the Laplace transform, determining the frequency of freebending vibrations, and constructing the amplitude-frequency response and phase-frequency response of the wing. [ABSTRACT FROM AUTHOR]

The design of a supersonic combustor ramjet engine has gained significant attention for futuristic air-breathing engines. Achieving efficient fuel mixing and complete combustion within a short period poses a substantial challenge. This study has directed efforts towards improving mixing and minimizing total pressure losses across the supersonic combustor to enhance performance. Experiments were conducted to investigate the influence of fuel injection geometry in the supersonic combustor with an entry Mach number of 2. Two different fuel orifice geometries, inclined at 45°, were considered. The investigation covered four different momentum flux ratios: 0.8, 1.0, 1.2, and 1.4 respectively. Various measurements were conducted to observe flow phenomena inside the supersonic combustion, including wall static pressure measurement, Schlieren visualization, exit Mach number and total pressure loss measurement. Without injection cases exhibited weaker compression and expansion inside the combustor. During injection, the rise in wall pressure indicated that the bow shock formed in front of the arc injection was slightly weaker than that of the circular injection. The impinging bow shock on the opposite wall also exhibited higher strength, resulting in a static pressure rise. As a result, the lower total pressure ratio across the shock indicates a higher momentum exchange between the main flow and the orifice. Therefore, arc injection has proven more effective in exchanging momentum inside the supersonic combustor. Consequently, the Mach number at the exit of the combustor was higher for the arc injection. [ABSTRACT FROM AUTHOR]

The pressure ratio of a compressor increases when the engine performance is improved, which leads to a higher air velocity in the combustion chamber. This aggravates the pressure loss in a conventional diffuser, which is proportional to the square of the inlet velocity. It is, therefore, an urgent need for a new diffuser technology. Regardless of whether high-temperature, or low-pollution, combustion chambers are being used, the significant difference when compared with other chambers is the notable increase in the combustion air and its entry through the dome of the flame tube. However, the increased amount of air in the dome causes a mismatch between the diffuser outlet and the dome intake of the flame tube. Therefore, to solve the problem of an excessive total pressure loss and a mismatch of the intake air in the combustor, a novel air-bleeding aerodynamic diffuser has here been proposed. The effects of various parameters (including the number of air-bleed holes, the position of the first row of holes, and the position of the second row of holes) and their interactions on the performance of this diffuser have then been investigated by using the experimental design presented by Taguchi. The maximum static pressure recovery was achieved by using a genetic algorithm combined with the CFD method. Among the three parameters, the results revealed that the position of the first row of holes had the largest impact on the performance of the diffuser. Also, the optimal values of the three parameters varied for the inlet Mach numbers 0.10, 0.15, and 0.20. The relative difference between the predicted values and the values obtained by the numerical simulations were all below 3%, which showed the reliability of the predictions. As compared with the reference case, the optimized results for the three working conditions increased by 19.16%, 21.38%, and 40.62%, respectively. [ABSTRACT FROM AUTHOR]

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A numerical simulation of an unsteady gas flow containing inert solid particles in a shock tube is carried out using the interpenetrating continuum model. The gas and dispersed phases are characterized by governing equations that express the concepts of mass, momentum, and energy conservation as well as an equation that shows the change of the volume fraction of the dispersed phase. Using a Godunov-type approach, the hyperbolic governing equations are solved numerically with an increased order of accuracy. The working section of the shock tube containing air and solid particles of various sizes is considered. The shock wave structure is discussed and computational results provide the spatial and temporal dependencies of the particle concentration and other flow quantities. The numerical simulation results are compared with available experimental and computational data. [ABSTRACT FROM AUTHOR]

The study of Reynolds number (Re) effects is crucial for optimizing train aerodynamics, enhancing safety and reducing energy consumption of high-speed trains. In this paper, a wind tunnel test was carried out to investigate the influence of Re on the train aerodynamic performance considering the shift of the air compressibility, which has not yet been explored to date. The test was conducted in a low-speed and large-scale tunnel with a stationary floor and the vehicle model was based on a 1/8th scaled train with three units. The Re ranges from 0.75 × 106∼3.12 × 106 by accelerating the uniform wind speed from U = 27.8 m/s to U = 115 m/s with zero-yaw. The Mach number (Ma) of the maximum speed scenario has exceeded 0.3, indicating that the airflow can be considered as being in the compressible range. The results show that the aerodynamic characteristics of high-speed trains exhibit a self-similarity region of Re , which is dependent on the flow velocity. The aerodynamic loads change little when Re ≥ 1.51 × 106, which corresponds to U ≥ 55.6 m/s. Therefore, the compressibility of the airflow within the range up to U = 115 m/s has a negligible effect on aerodynamic loads, and can thus be disregarded. However, the surface pressure significantly decreases when the incoming flow surpasses 0.3 Ma and transitions into a compressible state. While the compressibility has a relatively minor impact on macroscopic aerodynamics, it cannot be overlooked when considering detailed flow field, such as surface pressure. [ABSTRACT FROM AUTHOR]

Forward-swept wings were researched and introduced to improve maneuverability, control, and fuel efficiency while reducing drag and they are often used alongside canards, to further enhance their characteristics. In this research, the effects of canard dihedral angles on the wing loading of a forward-swept wing in transonic flow conditions were studied, as the wing loading provides a measure of wing's efficiency (lift/drag). A generic aircraft model from literatures was selected, simulated, and compared to, using CFD software ANSYS/Fluent where the flow equations were solved to calculate the aerodynamic characteristics. The research was carried at two different Mach numbers, 0.6 and 0.9, for five different canard dihedral angles which traverses from below the wing plane to above it, at various flow angles of attack. It was concluded that as the dihedral angle increased, lift increases for the same angle of attack for Mach 0.6 and 0.9 which increases the efficiency of the wing. The wing span loading occur at 10° dihedral angle for both Mach speeds, while, at 10° anhedral, the lift was minimal due to leading-edge flow separation on the FSW's lower surface. Thus, it is concluded that the canard at positive dihedral angles ensures increased wing span loading efficiency. [ABSTRACT FROM AUTHOR]

A two-dimensional mathematical model of the multiphase flow field during top blowing was created. The model is based on the rotating top-blown furnace of a certain copper plant's lead anode slime. The accuracy of the model was validated by comparing the results with empirical formulas. The study investigated the multiphase flow behavior under production conditions, utilizing the standard k–ε turbulence model and the VOF multiphase flow model. The effects of top-blown operating pressure, nozzle Mach number, lance position, and melt depth on the jet flow have been analyzed. It was discovered that the top-blown operating pressure and lance position have the most significant impact on the refining effect. The optimal operating parameters were determined as follows: operating pressure of 0.6 MPa, nozzle Mach number of 1.5, lance position at 0.5 m, and melt depth of 1.5 m. These can provide guidance for optimizing the production conditions of the lead anode slime rotating top-blown furnace. [ABSTRACT FROM AUTHOR]

The dispersion of shock is universal in various media, and in plasmas, standing whistler waves represent the dispersion of collisionless shocks. These standing whistler waves are usually observed upstream of low Mach number planetary bow shocks and interplanetary shocks. However, our understanding of the plasma behavior and electric field properties within these waves remains limited. Using conjoint THEMIS and Magnetospheric Multiscale (MMS) observations, we report the first observation of standing whistler wave upstream of a fast shock in Earth's magnetosheath resulting from the interaction between a solar wind tangential discontinuity and the bow shock. High‐resolution MMS measurements provide unprecedented insights into these waves, characterizing their circular polarization, near‐parallel propagation to the shock normal, and fixed phase relative to the shock ramp. The observed field and particle characteristics are discussed in detail. Key Points: An interplanetary tangential discontinuity interacted with Earth's bow shock, generating a fast shock in the magnetosheathFirst observation of standing whistler waves upstream of the fast shock in magnetosheathThe field and particle characteristics are discussed [ABSTRACT FROM AUTHOR]

Magnetic reconnection changes the magnetic field topology and facilitates the energy and particle exchange at magnetospheric boundaries such as the Earth's magnetopause. The flow shear perpendicular to the reconnecting plane prevails at the flank magnetopause under southward interplanetary magnetic field conditions. However, the effect of the out‐of‐plane flow shear on asymmetric reconnection is an open question. In this study, we utilize kinetic simulations to investigate the impact of the out‐of‐plane flow shear on asymmetric reconnection. By systematically varying the flow shear strength, we analyze the flow shear effects on the reconnection rate, the diffusion region structure, and the energy conversion rate. We find that the reconnection rate increases with the upstream out‐of‐plane flow shear, and for the same upstream conditions, it is higher at the dusk side than at the dawn side. The diffusion region is squeezed in the outflow direction due to magnetic pressure which is proportional to the square of the Alfvén Mach number of the shear flow. The out‐of‐plane flow shear increases the energy conversion rate J·E′ $\boldsymbol{J}\cdot {\boldsymbol{E}}^{\prime }$, and for the same upstream conditions, the magnitude of J·E′ $\boldsymbol{J}\cdot {\boldsymbol{E}}^{\prime }$ is larger at the dusk side than at the dawn side. This study reveals that out‐of‐plane flow shear not only enhances the reconnection rate but also significantly boosts energy conversion, with more pronounced effects on the dusk‐side flank than on the dawn‐side flank. These insights pave the way for better understanding the solar wind‐magnetosphere interactions. Key Points: The reconnection rate increases with the out‐of‐plane flow shear and is higher at dusk than at dawn side for the same upstream conditionsThe diffusion region is squeezed in the outflow direction, causing its length to decrease as the out‐of‐plane shear flow increasesThe energy conversion rate increases with flow shear strength and is higher at dusk than at dawn side for the same upstream conditions [ABSTRACT FROM AUTHOR]

The purpose of this work is to shed light on the effect of the pivot position on the surface pressure distribution over a 3D wing in different flight conditions. The study is intended to support the design and development of aerospace vehicles where stability analysis, performance optimization, and aircraft design are of primary importance. The following parameters are considered: Mach numbers (M) of 1.3, 1.8, 2.3, 2.8, 3.3, and 3.8, angle of incidence (θ) in the range from 5° to 25°, pivot position from h = 0.2 to 1. The results of the CFD numerical simulations match available analytical data, thereby providing evidence for the reliability of the used approach. The findings provide valuable insights into the relationship between the surface pressure distribution, the Mach number and the angle of incidence. [ABSTRACT FROM AUTHOR]

The nanofluid/air interface is practically used in enhancing heat transfer efficiency in thermal management systems, such as in cooling electronics and improving the performance of solar collectors. Additionally, it finds applications in advanced manufacturing processes and biomedical devices, where precise temperature control is crucial. The study investigates the instability of the interface between a Newtonian nanofluid and air in a rectangular setup. This instability arises when the two fluids flow at different velocities, leading to Kelvin-Helmholtz instability at the interface. The air is treated as a viscous, incompressible fluid due to its low Mach number, positioned above the nanofluid. The stability of the interface is determined based on the relative velocity of the fluid layers. The study reveals that various flow parameters, including viscosity ratio, density ratio, volume fraction, and nanoparticle diameter, influence the stability or instability of the interface. Four types of nanofluids are considered, and a comparative analysis is conducted. Interestingly, the nanofluid/air system is found to be more stable compared to the viscous liquid/air system. Sensitivity analysis is performed to examine the impact of different physical variables and their interactions on the critical relative velocity. It is observed that the critical velocity consistently exhibits positive sensitivity to the density ratio. Moreover, the magnitude of critical velocity sensitivities for the density ratio remains constant across all cases. The critical velocity demonstrates the highest positive sensitivity with respect to the parameter of air thickness, with this maximum sensitivity occurring when the air thickness equals 1 and the densities of both fluids are identical. [ABSTRACT FROM AUTHOR]

A sudden increase in the area of a duct or at the blunt base of the projectile leads to flow separation and reattachment. In the flow separation process, the base pressure at the blunt base is sub-atmospheric, leading to significant drag, which can be around sixty to seventy percent. This study is undertaken to regulate the base pressure in the recirculation zone and the flow field of the duct. This paper focuses on the effectiveness of quarter ribs of various radii in the range from 1 mm to mm and nozzle pressure ratio ranging from 3 to 11 at Mach M = 1.48 for a duct diameter of 22 mm and its sizes ranging from 1D to 6D. Some oscillations are observed for the rib location of 11 mm from the exit of the nozzle. Due to the proximity to the nozzle exit, these oscillations are observed. With a progressive shift of passive control along the more significant length, a continued rise in the pressure in the base region for rib radii in the range from 2 mm to 4 mm and an extreme increase in the base pressure is achieved for 4 mm rib radii placed at 66 mm inside the duct. Nevertheless, despite the maximum enhancement in pressure for duct size L = 4D, a negligible reduction in base pressure and ambient pressure cannot affect the flow contained by the duct for a more considerable duct length. However, using a quarter rib radii of 1mm is inadequate, and base pressure values are identical with and without rib except for the nozzle pressure ratio (NPR) = 3, where the nozzle at NPR = 3 is over-expanded. Except at NPR = 3, the nozzles are under-expanded, and the control mechanism becomes efficient, resulting in a significant base pressure increase. One can make a final decision based on the mission requirements about the radius of the rib, rib location, and level of expansion to meet the user's requirements.