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

In the present study, we develop a theoretical model for predicting hydrogen jet ignition in a shock tube with a partially opened diaphragm for hydrogen as a driver gas and air as a driven gas. Effects of pressure losses associated with the gradual diaphragm opening process are taken into account by developing a new discharge coefficient model. The discharge coefficient for incompressible flows is first calculated and then converted into the discharge coefficient for compressible flows. Three regimes are identified for the discharge coefficient, which depend on whether the ruptured portion of the diaphragm, which remains attached to the orifice, is negligible or not. We extensively validate the new model against experimental results from the literature. We show that the critical initial pressure ratio across the diaphragm and the shock-wave strength required for ignition can be calculated in a conservative manner with a maximum relative error equal to 9% and 10%, respectively, for the ratio of diaphragm opening area to driven-section cross-sectional area from 0.125 to 1.0. Finally, we explore, via our new model, the influence of different parameters on the shock-wave Mach number and ignition limits. We show that all our model predictions can be generalized into two simple dimensionless correlations. [Display omitted] • Three regimes are identified for the discharge coefficient. • Model predicts the ignition limits within 9% of experimental values. • The ignition limit is constant for area ratios between 0.52 and 1.0. • Model predictions can be generalized into two correlations. [ABSTRACT FROM AUTHOR]

The Kelvin-Helmholtz Instability (KHI), arising from velocity shear across the magnetopause, plays a significant role in the viscous-like transfer of mass, momentum, and energy from the shocked solar wind into the magnetosphere. While the KHI leads to growth of surface waves and vortices, suitable detection methods for these applicable to magnetohydrodynamics (MHD) are currently lacking. A novel method is derived based on the well-established A-family of hydrodynamic vortex identification techniques, which define a vortex as a local minimum in an adapted pressure field. The J x B Lorentz force is incorporated into this method by using an effective total pressure in MHD, including both magnetic pressure and a pressure-like part of the magnetic tension derived from a Helmholtz decomposition. The AMHD method is shown to comprise of four physical effects: vortical momentum, density gradients, fluid compressibility, and the rotational part of the magnetic tension. A local three-dimensional MHD simulation representative of near-flank magnetopause conditions (plasma 's 0.5-5 and convective Mach numbers Mf ~ 0.4) under northward interplanetary magnetic field (IMF) is used to validate AMHD. Analysis shows it correlates well with hydrodynamic vortex definitions, though the level of correlation decreases with vortex evolution. Overall, vortical momentum dominates AMHD at all times. During the linear growth phase, density gradients act to oppose vortex formation. By the highly nonlinear stage, the formation of small-scale structures leads to a rising importance of the magnetic tension. Compressibility was found to be insignificant throughout. Finally, a demonstration of this method adapted to tetrahedral spacecraft observations is performed. [ABSTRACT FROM AUTHOR]

The present paper focuses on the optimization of large-flow coefficient centrifugal compressors, utilizing a mature centrifugal compressor impeller with a flow coefficient of 0.16 under design point condition in engineering as the research subject. Due to the more complex flow mechanism and more design parameters in the impeller with large flow coefficient, the traditional artificial optimization method is insufficient. In present paper, the impeller with a large flow coefficient is optimized using the concept of combining physical principles and artificial intelligence tools. Firstly, the impeller underwent a redesign based on the theory of maximum flow capacity, with the aim of reducing the Mach number at the impeller inlet to enhance the compressor's performance. And the efficiency of the impeller at the design point has been increased from 88.6% to 89.9%. In order to further improve the performance of the impeller, an optimization algorithm grounded in gradient variation was employed to facilitate the automatic compressor optimization, and the flow losses at the impeller's top under low-flow conditions has been mitigated. The results of three-dimensional numerical simulation showed that the operating range of the new impeller is 7% wider than that of the original impeller. [ABSTRACT FROM AUTHOR]

The transition from subAlfvénic to superAlfvénic flow in the solar atmosphere is examined by means of Parker Solar Probe (PSP) measurements during solar encounters 8 to 14. Around 220 subAlfvénic periods with a duration ≥10 min are identified. The distribution of their durations, heliocentric distances, and Alfvén Mach number are analysed and compared with a global magnetohydrodynamic model of the solar corona and wind which includes turbulence effects. The results are consistent with a patchy and fragmented morphology, and suggestive of a turbulent Alfvén zone within which the transition from subAlfvénic to superAlfvénic flow occurs over an extended range of helioradii. These results inform and establish context for detailed analyses of subAlfvénic coronal plasma that are expected to emerge from PSP's final mission phase, as well as for NASA's planned PUNCH mission. [ABSTRACT FROM AUTHOR]

The unsteady pressure field over an axisymmetric backward-facing step was investigated experimentally at transonic freestream Mach numbers of 1.05, 1.2, and 1.4. The study was aimed at examining the influence of transonic freestream Mach numbers on the spatio-temporal character of the unsteady pressure field and on the dominant modes/mechanisms driving it. Surface flow visualization, schlieren, and unsteady pressure measurements were carried out as part of the experimental investigation. From oil flow visualization and schlieren, the reattachment region was identified, and consequently, the mean reattachment length was estimated. The mean reattachment length shows an increase with the increase in freestream Mach number. The coefficient of mean pressure along the rearbody imitates a classical backward-facing step flow profile and can be divided into three distinct regions. The peak values of the coefficient of mean pressure and the coefficient of root mean square of the fluctuation are seen to decrease with an increase in the freestream Mach number. Conventional spectral analysis reveals that as the freestream Mach number increases, the dominant peak in the spectra shifts to lower frequencies. From the spectra, three dominant fluid dynamic mechanisms depending on the freestream Mach number have been identified. Proper Orthogonal Decomposition (POD) analysis shows that 79–84 % of the total energy contribution comes from the first six modes. The temporal dynamics of the POD modes indicate three prominent mechanisms are responsible for the unsteady pressure field. Spectral analysis of POD modes indicates that the spectra are primarily driven by the first three POD modes for freestream Mach number of 1.05 and the first two modes for freestream Mach numbers of 1.2 and 1.4. Moreover, it reveals the presence of three dominant modes, and the freestream Mach number strongly dictates the dominant mode that is driving the pressure field. [ABSTRACT FROM AUTHOR]

At present, there are two independent technologies and equipment for the test of the reciprocating compressor plate valve: the sealing test and blowing test. This method has single function and complicated test operation and may rely on theoretical calculations, which reduces the reliability of the test conclusion because the motion law of valve disk during opening and closing is extremely complicated. The existing test methods cannot meet the needs of more in-depth and accurate research, development and improvement of the plate valve. This study proposes a test device and method that can not only quantitatively test the performance of the plate valve but also qualitatively test the dynamic process of the suction valve plate, which can greatly reduce the testing time provide more intuitive and accurate valve plate motion tracks and valve paths. This paper takes the primary plate suction valve as the test object and adopts the automated multi-functional system to accurately obtain the valve's sealing property, effective flow area, thrust coefficient, flow coefficient, drag drop, valve gap Mach number, initial opening pressure difference, maximum opening, and residence time of the valve plate on the valve lift guard to provide necessary data for valve design and performance evaluation. The test results of valve opening time are verified by industrial reciprocating compressor, which proves the accuracy and practicability of the system. [ABSTRACT FROM AUTHOR]

In the context of hypersonic flows, shock-wave/turbulent boundary-layer interactions (SWTBLIs) can lead to substantial aerodynamic heating. The commonly used Reynolds-averaged Navier–Stokes (RANS) method in engineering applications faces challenges in accurately predicting heat transfer in these conditions, as the assumptions made by Morkovin's assumption in the RANS method are not applicable in hypersonic SWTBLI flows. This article addresses these challenges by introducing a variable turbulent Prandtl number (P r t ) model for non-adiabatic walls, established through field inversion and symbolic regression (SR) techniques. The methodology begins with field inversion for an oblique SWTBLI case at Mach 5. The corrected field, denoted as β (x) , obtained from this inversion, is integrated with selected local flow characteristics to derive a corrected expression using SR. Following the necessary adjustments, this expression is incorporated into the RANS solver. Various cases with different wall cooling ratios, Mach numbers, and Reynolds numbers are chosen to assess the generalization ability of the variable P r t model. The outcomes demonstrate a substantial improvement in the model's ability to predict heat transfer in hypersonic SWTBLI flows without compromising the baseline model's performance in predicting the undisturbed boundary layer. The correction effect remains notably enhanced even in complex three-dimensional cases. • Corrected heat flux and skin friction by adjusting Reynolds stress and turbulent Prandtl number. • Developed a variable Prandtl number model using symbolic regression. • Analyzed physical mechanisms of the modified SST model for separated flow. analyzed. • Nine hypersonic SWTBLI flows showed significant improvements in calculations. [ABSTRACT FROM AUTHOR]

The compressor transition duct transfers the air from the low-pressure compressor to the high-pressure compressor. An aircraft always experiences off-design conditions except for level flight conditions during aircraft range. So, to throw some light on the performance of transition ducts under off-design conditions, this paper presents its performance assessment numerically for uniform and swirl flow cases under varying Mach numbers (M). This analysis reveals that a higher Mach number provokes an adverse pressure gradient and non-uniformity. In the uniform flow, the non-uniformity attains a maximum value of 0.04366 near the inner wall, which indicates the presence of a highly distorted flow near the inner wall; however, its maximum value is 0.03661 near to the outer wall in the swirl flow. Also, the pressure recovery reduces with an increased Mach number. Notably, as the Mach number increasesfrom the design point, the pressure coefficient along the convex curvature significantly drops to the lowest point of −0.65015 and −0.79324 for uniform and swirl flow, respectively. Consequently, the inner wall experiences a large adverse pressure gradient (−0.90822) compared to the decreasing Mach number from the design point for the uniform flow, whereas it is experienced by the outer wall in the case of swirl flow (−0.94729). [ABSTRACT FROM AUTHOR]

This study investigates the potential of an optimized combination of active nose droop and active Gurney flap (CADAG) in a new flow control strategy to manage dynamic stall over a pitching blade section under variable Mach number flow. The optimization method employs the genetic algorithm coupled with a computational fluid dynamic (CFD) solver and artificial neural network. The base blade section belongs to a section positioned at r/R=0.865 of the rotor blade of the UH-60A helicopter in forward flight condition. A high relative angle of attack on the retreating side makes the flow susceptible to dynamic stall. A nose droop is employed to control the dynamic stall of the blade section, and a Gurney flap is used to maintain the balance of the generated lift of the blade during 360° of rotation. A comprehensive investigation is performed to determine the most significant parameters affecting the performance of the present combined active flow control. The ratio of the total generated lift to the drag is chosen as the objective function of the optimization. Results show that this ratio and the total generated lift in one rotation cycle increase by 193% and 13%, respectively, at the optimum condition of the present combined active flow control, while the ratio of the generated lift over the advancing side to the retreating side is equal to that of the base blade section. In addition, the dynamic stall hysteresis loop reduces significantly, and the maximum value of the drag coefficient and the negative aerodynamic damping decrease up to 87% and 83% compared to the base blade section, respectively. In general, the proposed innovative combined active flow control is an adjustable method to alleviate dynamic stall and improve the aerodynamic performance of rotary wings in different operation conditions. Practical Applications: The rotary blades are extensively used in rotorcraft, turbo engines, and wind turbines. Despite their massive use, they suffer from some essential issues, of which the most important one is the so-called dynamic stall. Dynamic stall is a complex phenomenon that limits the performance of the rotary blades. Due to the physics governing a rotary wing, such as a helicopter rotor blade, dynamic stall and flow separation are very common. Understanding dynamic stall physics and providing solutions to prevent it is still one of the main challenges of aerodynamic scientists. The present study introduces a novel adjustable method for practically alleviating the dynamic stall on helicopter blade section in forward flight conditions to improve its aerodynamic performance in different operational conditions. A comprehensive investigation is carried out to determine the key parameters affecting the proposed method. These findings can serve as a valuable tool for other researchers to develop various active flow control strategies. This article applies an optimization process using artificial neural networks, genetic algorithms, and CFD tools, which forms a comprehensive framework that can be easily extended to other applications. [ABSTRACT FROM AUTHOR]

The elliptic shock wave has potential applications in the design of hypersonic inlets and waveriders. In this study, an optimization design process and a mathematical model for an elliptic shock wave are proposed. An elliptic cone flow field is employed to generate an elliptic shock wave whose shape is predetermined using the value of aspect ratio. The design of the shock wave is transformed into an optimization problem, and an optimization design process with the gas viscosity considered is developed. The elliptic cone is parameterized, and the parameters are adjusted by an optimization algorithm to meet the requirements. The results show that the optimization process can realize the accurate design of an elliptic shock wave. Based on the analysis of the elliptic cone flow field, the shock wave can be well approximated as an elliptic cone, and the aspect ratio of the shock wave is positively correlated with three factors: the aspect ratio and slenderness ratio of the elliptic cone and the incoming Mach number. The relationship between the aspect ratio of the shock wave and these three factors is modeled by a feed-forward neural network with accuracy verified, which is beneficial for the rapid design of elliptic shock waves. [ABSTRACT FROM AUTHOR]

Rocket sled test avoids boundary effects in wind tunnel test and inconvenience of flight test, which is one of the feasible options for future single-stage-to-orbit. To analyze potential safety issues during payload separation and optimize the arrangement of testing sensors, different structural layouts and operating speeds of the rocket sled are conducted based on computational fluid dynamics and computational aeroacoustics. The hypersonic winged standard model is utilized as the load for these simulations. The analysis encompasses the evolution of shock waves, the forces exerted on the payload and noise propagation. Variations in flow field and aeroacoustic characteristics of rocket sled are analysed, revealing underlying physical mechanisms. It is observed that placing the payload in front of the thruster can reduce the head pressure, excessive wingspan may have an impact on the structural safety of the wingtip, and the regions of high sound pressure levels mainly exists in the middle and rear sections of the rocket sled. Moreover, as the Mach number increases, the characteristic frequency initially rises and then declines. These researches can serve as a valuable reference for the development of ground test systems and single-stage-to-orbit. [ABSTRACT FROM AUTHOR]

This paper introduces a novel platform that integrates three preconditioning matrices based on conservative variables: the Turkel, Choi–Merkle and Onur matrices. The platform aims to compare these matrices in terms of accuracy and robustness by investigating their performance in solving three distinct and challenging high-gradient laminar flow problems: (i) Bi-Plane NACA0012 airfoil, (ii) lid-driven flow in a square cavity and (iii) flow in a planar T-junction. These problems serve as new and challenging test cases to accurately determine the abilities of the preconditioning matrices. The preconditioning matrices are applied to evaluate the numerical solutions, and their performance in complex flow fields is assessed in terms of accuracy and efficiency. By solving these flow problems, the effectiveness of the preconditioning matrices is thoroughly analyzed. By integrating these preconditioning matrices into a single platform, this paper significantly contributes to the field. The approach enables a direct and meaningful comparison of the performance of the Turkel, Choi–Merkle and Onur matrices in solving these new and challenging laminar flow problems. While all three matrices demonstrate comparable accuracy in predicting flow characteristics like pressure coefficients, Turkel's method shows superior convergence acceleration across the almost test cases due to alpha parameter and modifications of the momentum equations. In the external flow case, Turkel converges 22–53% faster than the other matrices by adjusting momentum terms with an alpha parameter. For the internal lid-driven cavity case, Turkel again accelerates convergence up to 38% over Choi–Merkle as Reynolds and Mach numbers increase. However, Onur's method stalls at high Reynolds/Mach numbers. At low values of Reynolds and Mach numbers, Onur reduces computational cost by 17%. Finally, for the T-junction case, Turkel and Choi–Merkle perform almost identically, decreasing CPU time by 55% vs Onur, which lacks convergence robustness. While the preconditioning matrices have similar accuracy, Turkel offers the best convergence improvement by accounting for momentum effects. Onur works well at low Reynolds/Mach numbers but shows limitations at higher values. The selection of the optimal preconditioning matrix should be based on whether momentum or viscosity dominates in the flow field, as well as the intensity of the viscosity gradient rate. [ABSTRACT FROM AUTHOR]

High-order accurate Large-eddy simulations (LES) have been carried out for compressible, turbulent flow through two bell-shaped convergent-divergent nozzles- one with a short divergent portion and the other with a longer one. An explicit filtering approach based on approximate deconvolution method is used for the LES. The nozzles have isothermal walls and circular cross-sections. The incoming flow is a turbulent, subsonic fully developed pipe flow at centreline Mach number of around 0.4 and friction Reynolds number 216. An adaptive filter is used to capture the shocks appearing in the divergent portion of the nozzles. The complicated behaviour of the mean flow, Reynolds stresses, production and pressure-strain terms of the axial Reynolds stress budget in the shock train region is shown. The broadband nature of the shock train oscillations having multiple dominant frequencies in these nozzles is observed in the present LES. [ABSTRACT FROM AUTHOR]

A stable upwind finite-difference method for unsteady gas-liquid two-phase flows is proposed and applied to shock tube flows. The artificial dissipation terms in the flux difference splitting upwinding scheme are derived using a preconditioned matrix to enhance the stability and convergence of the numerical calculation of mixed compressible and incompressible flows with arbitrary void fractions. A homogeneous gas-liquid two-phase flow model is used. A stable four-stage Runge-Kutta method and the flux difference splitting upwind scheme combined with a third-order MUSCL TVD scheme are employed. Using the proposed method, we compute gas-liquid mixture shock tube problems and compare their results with the exact solution to check the reliability of the proposed method. Shock and expansion wave propagations through the gasliquid two-phase media are observed in detail. The effect of the preconditioned artificial dissipation on the numerical stability and convergence rate are investigated. We confirm that the proposed method is stable and effective for computations of unsteady two-phase complex flows with arbitrary Mach numbers. [ABSTRACT FROM AUTHOR]

A computational fluid dynamics investigation of a turbulent supersonic jet impacting a solid flat plate is conducted utilizing the OpenFOAM software. The research focuses on simulating the three-dimensional mean compressible flow for jet impingement on an inclined plate by analyzing the complex flow field and the surface distribution of pressure. Various simulations are carried out at a jet Mach number of 2.2 maintaining a constant nozzle-to-plate distance while varying the angle of inclination of the plate. In contrast to earlier numerical studies, this work employs a modern turbulence modeling technique known as detached eddy simulation (DES), along with a traditional unsteady Reynolds-averaged Navier–Stokes model. Making a comparison with experimental findings, the current analysis reveals that both turbulence modeling techniques effectively predict the mean pressure distribution on the plate. However, the DES approach offers deeper insights into the turbulent flow field, showing notable consistency with the experiments. The complex compressible flow patterns are simulated with higher accuracy compared to the traditional approach. Enhanced turbulence resolution is attained by utilizing the same computational grid with a limited increase in computational complexity. [ABSTRACT FROM AUTHOR]

When better fuel-air mixing in the combustion chamber or a reduction in base drag are required in vehicles, rockets, and aeroplanes, the base pressure control is activated. Controlling the base pressure and drag is necessary in both scenarios. In this work, semi-circular ribs with varying diameters (2, 4, and 6 mm) positioned at six distinct positions (0.5D, 1D, 1.5D, 2D, 3D, and 4D) inside a square duct with a side of 15 mm are proposed as an efficient way to apply the passive control technique. In-depth research is done on optimising rib size for various rib sites. According to this study, the base pressure rises as rib height increases. Furthermore, the optimal location for the semi-circular ribs with a diameter of 2 mm is at 0.5D. The 1D location appears to be optimal for the 4 mm size as well. For the 6 mm size, however, the 4D position fills this function. Graphic Abstract [ABSTRACT FROM AUTHOR]

The unsteady mechanism of unstart flow for an inlet with rectangular-to-elliptical shape transition (REST) under the off-design condition at a Mach of 4 is investigated using the delay detached eddy simulation method. With the help of numerical simulations, the unsteady dynamics, especially the low-frequency characteristics of the REST inlet unstart flow, as well as the self-sustaining mechanism, is investigated. The instantaneous flow illustrates the unsteady phenomena of the REST unstart flow, including the interaction between the cowl-closure leading edge (CLE) shock and the shear layer, breathing of the separation bubble, flapping of the separation shock, instability of the shear layer and vortex shedding along the shear layer. The spectral analysis reveals that the lower frequency dynamics is associated with the breathing of the separation bubble and the flapping motion of the separation shock wave, while the higher frequency is related to the instability of the shear layer affected by cowl-closure leading edge shock and the formation of shedding vortices. Further, coherence analysis shows that the contribution of these flow structures dominating the low-frequency dynamics couple with each other. Based on the dynamic mode decomposition results, the characteristics that contribute to the unsteady behaviour of unstart flow are summarized. The streamwise vortices downstream of the separation and the shedding vortices are believed to be the main driving force of the global low-frequency unsteadiness of the REST inlet unstart flow under the off-design condition. Moreover, the CLE shock plays an important role in the process during the dominant flow structure conversion from the backflow within the separation bubble into elongated streamwise structures. [ABSTRACT FROM AUTHOR]

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

The efficacy of steady large-amplitude blowing/suction on instability and transition control for a hypersonic flat plate boundary layer with Mach number 5.86 is investigated systematically. The influence of the blowing/suction flux and amplitude on instability is examined through direct numerical simulation and resolvent analysis. When a relatively small flux is used, the two-dimensional instability critical frequency that distinguishes the promotion/suppression mode effect closely aligns with the synchronisation frequency. For the oblique wave, as the spanwise wavenumber increases, the suppression effects would become weaker and the mode suppression bandwidth diminishes/increases in general in the blowing/suction control. Increasing the blowing/suction flux can effectively broaden the frequency bandwidth of disturbance suppression. The influence of amplitude on disturbance suppression is weak in a scenario of constant flux. To gain a deeper insight into disturbance suppression mechanism, momentum potential theory (MPT) and kinetic energy budget analysis are further employed in analysing disturbance evolution with and without control. When the disturbance is suppressed, the blowing induces the transport of certain acoustic components along the compression wave out of the boundary layer, whereas the suction does not. The velocity fluctuations are derived from the momentum fluctuations of the MPT. Compared with the momentum fluctuations, the evolutions indicated by each component's velocity fluctuations greatly facilitate the investigations of the acoustic nature of the second mode. The rapid variation of disturbance amplitude near the blowing is caused by the oscillations of the acoustic component and phase speed differences between vortical and thermal components. Kinetic energy budget analysis is performed to address the non-parallel effect of the boundary layer introduced by blowing/suction, which tends to suppress disturbances near the blowing. Moreover, viscous effects leading to energy dissipation are identified to be stronger in regions where the boundary layer is rapidly thickening. Finally, it is demonstrated that a flat plate boundary layer transition triggered by a random disturbance can be delayed by a blowing/suction combination control. The resolvent analysis further demonstrates that disturbances with frequencies that dominate the early transition stage are dampened in the controlled base flow. [ABSTRACT FROM AUTHOR]

Wall temperature has a significant effect on shock wave/turbulent boundary layer interactions (STBLIs) and has become a non-negligible factor in the design process of hypersonic vehicles. In this paper, direct numerical simulations are conducted to investigate the wall temperature effects on STBLIs over a 34° compression ramp at Mach number 6. Three values of the wall-to-recovery-temperature ratio (0.50, 0.75 and 1.0) are considered in the simulations. The results show that the size of the separation bubble declines significantly as the wall temperature decreases. This is because the momentum profile of the boundary layer becomes fuller with wall cooling, which means the near-wall fluid has a greater momentum to suppress flow separation. An equation based on the free-interaction theory is proposed to predict the distributions of the wall pressure upstream of the corner at different wall temperatures. The prediction results are generally consistent with the simulation results (Reynolds number Reτ ranges from 160 to 675). In addition, the low-frequency unsteadiness is studied through the weighted power spectral density of the wall pressure and the correlation between the upstream and downstream. The results indicate that the low-frequency motion of the separation shock is mainly driven by the downstream mechanism and that wall cooling can significantly suppress the low-frequency unsteadiness, including the strength and streamwise range of the low-frequency motions. [ABSTRACT FROM AUTHOR]

In the present study, we perform direct numerical simulations of compressible turbulent boundary layers at free stream Mach numbers $2\unicode{x2013}6$ laden with dilute phase of spherical particles to investigate the Mach number effects on particle transport and dynamics. Most of the phenomena observed and well-recognized for inertia particles in incompressible wall-bounded turbulent flows – such as near-wall preferential accumulation and clustering beneath low-speed streaks, flatter mean velocity profiles, and trend variation of the particle velocity fluctuations – are identified in the compressible turbulent boundary layer as well. However, we find that the compressibility effects are significant for large inertia particles. As the Mach number increases, the near-wall accumulation and the small-scale clustering are alleviated, which is probably caused by the variations of the fluid density and viscosity that are crucial to particle dynamics. This can be affected by the fact that the forces acting on the particles with viscous Stokes number greater than 500 are modulated by the comparatively high particle Mach numbers in the near-wall region. This is also the reason for the abatement of the streamwise particle velocity fluctuation intensities with the Mach numbers. [ABSTRACT FROM AUTHOR]

Starting from a microscopic particle model whose hydrodynamic limit under hyperbolic space-time scaling is a 1D conservation law, we derive the large deviation rate function encoding the probability to observe a density profile which is a non entropic shock, and compare this large deviation rate function with the classical Jensen-Varadhan functional, valid for the totally asymmetric exclusion process and the weakly asymmetric exclusion process in the strong asymmetry limit. We find that these two functionals have no reason to coincide, and in this sense Jensen-Varadhan functional is not universal. However, they do coincide in a small Mach number limit, suggesting that universality is restored in this regime. We then compute the leading order correction to the Jensen-Varadhan functional. [ABSTRACT FROM AUTHOR]

This paper introduces a novel numerical method for investigating the dynamic stability of a flutter panel exposed to a supersonic gas flow and a fluctuating axial excitation force. Initially, the system equation of motion was derived using Lagrange’s equation, where the first two modes are coupled Mathieu–Hill equations with damping, constituting a system of linear second-order differential equations with periodically variable coefficients. Subsequently, a new numerical method was proposed to analyze the dynamic stability of coupled Mathieu–Hill equations with damping. This method involves breaking down an arbitrary parametric load into discrete segments to approximate the variable excitation function using a step function. The system responses of each segment are then accumulated in matrix form. The proposed numerical method proves particularly effective for dynamic systems whose parameters cannot be treated as small. In practical application, the method allows the construction of instability regions corresponding to natural frequencies, subharmonics, and combination frequencies. Dynamic stability diagrams were generated based on dynamic pressure ratio, air/panel density ratio, Mach number, panel thickness—length ratio, and excitation frequency. The results demonstrated general agreement with those obtained through Hsu’s perturbation method, however, our numerical results have proven more accurate. The paper concludes by offering suggestions for suppressing panel flutter through appropriate parameter combinations. [ABSTRACT FROM AUTHOR]

The effects of inlet Mach number on the unsteadiness of shock-boundary layer interactions (SBLIs) over curved surfaces are investigated for a supersonic turbine cascade using wall-resolved large eddy simulations. Three inlet Mach numbers, 1.85, 2.00, and 2.15 are considered at a chord-based Reynolds number 395,000. The curved walls of the airfoils impact the SBLIs due to the state of the incoming boundary layers and local pressure gradients. On the suction side, due to the convex wall, the boundary layer entering the SBLI evolves under a favorable pressure gradient and bulk dilatation. On the other hand, the concave wall on the pressure side imposes an adverse pressure gradient and bulk compression. Variations in the inlet Mach number induce different shock impingement locations, enhancing these effects. A detailed characterization of the suction side boundary layers indicates that a higher Mach number leads to larger shape factors, favoring separation and larger bubbles, while the reverse holds for the pressure side. A time-frequency analysis reveals the presence of intermittent events in the separated flow occurring predominantly at low-frequencies on the suction side and at mid-frequencies on the pressure side. Increasing the inlet Mach number leads to an increase in the time scales of the intermittent events on the suction side, which are associated with instants when high-speed streaks penetrate the bubble, causing local flow reattachment and bubble contractions. Instantaneous flow visualizations show the presence of streamwise vortices developing on the turbulent boundary layers on both airfoil sides and along the bubbles. These vortices influence the formation of the large-scale longitudinal structures in the boundary layers, affecting the mass imbalance inside the separation bubbles. [ABSTRACT FROM AUTHOR]

To address airflow interference and inertial altitude dispersion during high-speed aircraft flight, a novel algorithm integrating cubature Kalman filter (CKF), air data observation system, and inertial navigation system (INS) is proposed for real-time aircraft air data estimation through sensor data fusion. The algorithm utilizes nonlinear equations and models interactions among the inertial system, satellite positioning system, and air data system to integrate sensor data for computing aircraft velocity and altitude, followed by estimation of parameters such as angle of attack and sideslip. Experimental results demonstrate the high accuracy and stability of the proposed method in estimating airflow angle and Mach number. [ABSTRACT FROM AUTHOR]

A shock-tube facility capable of generating a planar shock with the Mach number higher than 3.0 is developed for studying Richtmyer–Meshkov instability induced by a strong shock wave (referred to as strong-shock RMI). Shock enhancement is realized through the convergence of shock within a channel with the profile determined by using shock dynamics theory. The facility is designed considering the repeatability of shock generation, transition of shock profile, and effects of viscosity and flow choking. By measuring the dynamic pressure of the tube flow using pressure sensors and capturing the shock movement through the high-speed shadowing technique, the reliability and repeatability of the shock tube for generating a strong planar shock are first verified. Particular emphasis is then placed on the ability of the facility to study strong-shock RMI, for which a thin polyester film is adopted to form the initial interface separating gases of different densities. The results indicate that the shock tube is reliable for conducting strong-shock RMI experiments. [ABSTRACT FROM AUTHOR]

A free argon jet flow accompanied by small clusters formation is studied with the direct simulation Monte Carlo method. Some near-continuum flow regimes characterized by Knudsen numbers in the 2 × 10 − 4 − 2 × 10 − 3 range are considered. A model for the argon clusters' growth/decay is proposed, taking into account the phase state of the clusters. The model consists of a chain of reactions leading to the clusters' formation, including the clusters' growth via triple/pair collisions of particles, and the clusters decay according to the collisional/unimolecular mechanism. The cluster size distributions in the jet far field are obtained. The results are compared with two experimental datasets. Good agreement is shown for most of the considered range of parameters. The triple particle collisions' influence on the argon clusters growth process is studied, and their important role in small cluster formation is demonstrated. It has been established that the cluster formation process is limited to an enough small spatial zone near the source outlet, of the order of several exit orifice diameters. The simulation shows a significant influence of cluster formation on the temperature and Mach number distributions, and a weak influence on the flow velocity. The formed clusters' translational temperatures and their velocities are close to the argon atoms' corresponding parameters. A non-equilibrium state, featured by a significant difference between the clusters' internal temperatures and the flow temperature, develops with distance from the source outlet. [ABSTRACT FROM AUTHOR]

Efficient initiation of oblique detonation waves (ODWs) is crucial for optimizing the performance of oblique detonation wave engines. A novel approach is proposed for enhancing ODW initiation through expansion waves in this research. Validation of the expansion wave-reinforced initiation method is conducted via numerical simulations employing multi-species reactive Euler equations and a pressure-dependent reaction mechanism. Results demonstrate a significant reduction in the initiation length of ODWs with the addition of an expansion wave ahead of the wedge, contrasting with the absence of detonation wave initiation on a wedge lacking an expansion wave. A theoretical model, based on expansion wave and shock wave relations, along with constant volume combustion theory, elucidates the underlying mechanism of reinforcement. The model reveals that crossing the expansion wave elevates the fluid's Mach number and locally enlarges the flow deflection angle on the wedge surface, without altering the wedge's structure. Furthermore, post-shock temperature increases and pressure decreases compared to the wedge not encountering an expansion wave. The heightened temperature predominantly triggers ODW initiation, thus reinforcing the process. Theoretical analyses indicate the reinforcement's greater efficacy at lower inflow temperatures and lower inflow Mach numbers, suggesting the expansion wave's suitability for initiation in the early flight stages of an aircraft equipped with oblique detonation wave engines. [ABSTRACT FROM AUTHOR]

The present study investigates the cause of low-frequency unsteadiness in shock wave/turbulent boundary layer (TBL) interactions. A supersonic turbulent flow over a compression ramp is studied using wall-resolved large eddy simulation (LES) with a freestream Mach number of 2.95 and a Reynolds number (based on δ 0 : the thickness of the incoming TBL) of 63 560. From the view of stability analysis, the effect of intrinsic instability on such low-frequency unsteadiness is excluded from the flow system by designing a ramp angle of 15 ° , and our attention is paid to the convective instability contributed by the incoming TBL. The LES results are analyzed by linear and nonlinear disambiguation optimization (LANDO), spectral proper orthogonal decomposition (SPOD), and resolvent analysis. The LANDO results reveal a streamwise scale-frequency relation of coherent structures in a very long (around 60 δ 0 ) TBL, which indicates that the dynamics of very large-scale motions (VLSMs) in the TBL are featured by a low frequency. The SPOD results reveal that the most energetic SPOD mode features a low frequency that is identical to the dominant low frequency of the wall-pressure spectrum. Additionally, coherent structures of the mode resemble the VLSMs in the incoming TBL. These consistencies imply that the dynamics of VLSMs contribute to the low-frequency unsteadiness of the present flow. A resolvent analysis then further suggests that the origins of low-frequency dynamics of the present flow are from the VLSMs, which can be optimally amplified by the forcing in the turbulent flow. [ABSTRACT FROM AUTHOR]

This paper presents the findings of a wind tunnel experiment aimed at investigating the mode transition process of a two-dimensional over-under turbine-based combined cycle inlet at an incoming flow Mach number of 2.9. The study utilized high-speed schlieren and dynamic pressure acquisition systems to examine the evolution process of the shock-dominated flow structure of the high-speed duct during the mode transition process. Additionally, the impact of mode transition speed on the unstart/restart characteristics of the high-speed duct was analyzed. The results indicate that, during the forward mode transition process, the increasing captured airflow of the high-speed duct leads to a higher number of shock reflections and the shock train moves forward in the duct, ultimately resulting in unstart. The unstarting flow field exhibits a small oscillation characteristic dominated by the separation bubble located at the entrance. However, evident hysteresis characteristics were observed in the restart process during the reverse transition. Furthermore, a higher mode transition speed delays the unstart and restart of the high-speed duct, consequently increasing the hysteresis interval. Theoretical analysis suggests that a larger mode transition speed leads to lower mass accumulation efficiency in the high-speed duct, thereby slowing the pressure response and causing the shock train to lag forward, resulting in delayed unstart. The delay in the restart process is attributed to the relative slip motion of the separation bubble with the upper surface of the splitter, in addition to its forced motion. [ABSTRACT FROM AUTHOR]

Research into the flow mechanisms and heat transfer characteristics of two-dimensional supersonic cooling films has been extensive, yet studies on supersonic conical cooling films remain scarce. This study examined the flow characteristics and mechanisms of supersonic conical cooling film under varying ratios of static pressure (RSPs) using nanoparticle planar laser scattering and particle image velocimetry techniques. The mainstream Mach number was M a ∞ = 3.8 , and the supersonic conical cooling film was tangentially injected through a precisely calibrated annular nozzle of M a j = 2.8. Analysis of the instantaneous flow structure characteristics along and normal to the flow direction segmented the flow downstream of the slot into three distinct zones: the potential core region, the large-scale coherent structure region, and the turbulent boundary layer region. The reattachment location was assessed both qualitatively and quantitatively based on time-averaged and statistical velocity field data. The results indicated that the reattachment location extended with higher RSP. Moreover, the reattachment point of the supersonic conical cooling film was further forward than that of a two-dimensional supersonic cooling film. It was found that the supersonic conical cooling film was more significantly affected by the compression effect of the conical stream tube compared to the impact of RSP. [ABSTRACT FROM AUTHOR]

A V-shaped blunt leading edge(VSBLE) at the lips of three-dimensional inward-turning inlets often generates complex shock interactions and severe aerothermal loads. To date, few heat flux reduction schemes have been devised based on the generation principle of the heat flux peaks on the VSBLE. Employing a synergistic approach of theoretical analysis and numerical simulation, it is determined that the maximum pressure jump ( p max / p ∞ ) provided by the near-wall secondary shock interaction structure significantly influences the outermost heat flux peak ( q peak 1 ) and exhibits a strong correlation with the deflection angle ( θ EF ) of the upstream flow. Consequently, we design a heat flux reduction model incorporating an expansion corner (EC) to validate this finding. The simulation results demonstrated a 54.23% reduction in q peak 1 under the influence of the expansion fan generated by the EC at Mach 6, although the remaining heat flux peaks, q peak 2 and q peak 3 , experience some deterioration. After conducting a comparative analysis and modeling the causes of deterioration, we propose an improved model by designing the crotch centerline. This improved model consistently demonstrated excellent performance in reducing the maximum heat flux peak and preventing the deterioration of the other heat flux peaks at Mach 6-12. Additionally, correlations between q peak 1 and p max / p ∞ , as well as between q peak 1 and θ EF are established. This paper presents an investigation into the key factors affecting the heat flux peaks on the VSBLE and proposes a model capable of ensuring a stable reduction in the maximum heat flux of over 40% under a wide range of Mach numbers. [ABSTRACT FROM AUTHOR]

Thermochemical nonequilibrium phenomena are key factors affecting the numerical simulation results of flow fields of Mars rover. In order to make clear the flow field properties of the Mars Pathfinder, a multi-physics numerical model including the Navier–Stokes equations, 11 species and 59 chemical reactions of CO2 plasma, and a four-temperature thermodynamic nonequilibrium model is established in this study. Numerical simulations on the thermal-chemical nonequilibrium flow fields of the Mars Pathfinder spacecraft are performed. The distributions of flow-field parameters such as temperature, pressure, Mach number, molar fractions of CO and CO2 around the Mars Pathfinder at different altitudes are obtained. It is found that there is a subsonic backflow at the rear of the Mars Pathfinder, resulting in slightly higher temperatures than other areas; after the shoulder, the pressure decreases rapidly; as the altitude decreases, the intensity of chemical reactions decreases, and the thermal-chemical nonequilibrium effects gradually weaken. Based on the chemical kinetic model of CO2 we proposed, the numerical simulation of the nonequilibrium flow field of the Mars Pathfinder spacecraft showed good predictive accuracy and application potential, providing reference value for subsequent studies on the thermal-chemical nonequilibrium effects of the Mars Pathfinder spacecraft. [ABSTRACT FROM AUTHOR]

The RBCC (rocket-based combined-cycle) engine integrates a rocket engine into the flow passage of the ramjet engine, thereby significantly broadening the operating range and becoming one of the potential solutions for the reusable space transportations. The embedded rocket, as one of the core components of an RBCC engine, is strongly coupled to other components, such as the inlet and combustor, and can induce significant impacts on the inlet start behaviors. For the stable operation of the RBCC inlet, the start/unstart hysteresis characteristics and the reliable start boundaries driven by embedded rocket are experimentally and numerically studied. The results show that due to the function of embedded rocket jet, the inlet start/unstart hysteresis range of Mach number rises from Mach 1.79–1.87 to 1.85–1.90, and the back pressure hysteresis range changes from Pc/P∞ of 3.9–5.8 to 4.7–5.8. Meanwhile, during the increasing and decreasing process of the embedded rocket pressure, the inlet undergoes a transition between start and unstart, accompanied by hysteresis phenomena as well. Furthermore, three different control strategies are proposed for the start of RBCC inlet. Similar to the traditional ramjet inlets, the RBCC inlet can self-/re-start by reducing the back pressure. Particularly, owing to the synergistic effect of embedded rocket jet and back pressure, the RBCC inlet can self-/re-start by reducing the embedded rocket pressure from Procket/P∞ = 58 to 39 in the "embedded rocket dominated" cases, while the inlet self-/re-start can be achieved by increasing the rocket pressure from Procket/P∞ = 20 to 39 in the "back pressure dominated" cases. [ABSTRACT FROM AUTHOR]

Large-eddy simulation (LES) of the Navier–Stokes equations is carried out to investigate the supersonic turbulent boundary layer (STBL) developing over a flat plate at free-stream Mach number M ∞ =2.0 and Reynolds number R e δ ≈ 13 500. The seventh-order weighted essentially non-oscillatory (WENO-7) scheme is adopted to ensure calculation accuracy. To demonstrate the credibility of the proposed improvements, a posteriori test is conducted. Other LES-like methods are also adopted for comparison. The validation results show that WENO-LES can provide credible predictions when grid resolution is sufficient. Thus, current LES results can be regarded as a reliable database for further analysis. The distributions of flow topologies and turbulent statistics are obtained after time average operation. A preliminary conclusion can be drawn that the position y + ≈ 14 in buffer layer may be the appropriate interface for Reynolds-averaged Navier–Stokes (RANS) and LES computations of LES-like methods in STBL. Diverse length scales in LES-like methods are calculated based on high-fidelity flow fields. On this basis, a blending function is added to the original von Karman length formula. A dissipation-adaptive length scale is proposed for the scale-adaptive simulation (SAS) method. A comparison of the original SAS and delayed-DES (DDES) methods reveals that the improved method exhibits grid-independent characteristics of RANS/LES interface. Furthermore, the coherent vortical structures and flow visualization of the a posteriori test indicate that the improved method has a satisfactory ability to promote the generation of small-scale structures and to capture turbulent fluctuations. [ABSTRACT FROM AUTHOR]

Stable combustion is desirable for efficient operation of scramjet engines at high flight Mach numbers, and being able to reconstruct the flow-field wave patterns in stable combustion facilitates proactive evaluation of engine operating conditions. Proposed here is a multi-level connected shifted-window transformer (MCSwinT) model for reconstructing the flow-field wave patterns of stable combustion in a supersonic combustor. A combustion feature conversion block is used to convert high-dimensional and low-dimensional combustion features; a deep pressure feature extraction block is used to extract the flow-field wave patterns, and MCSwinT blocks enable multi-level fusion, thereby extracting the high-dimensional combustion features of the flow-field wave system. A dynamic loss function unifies spatial content loss and feature space loss, leading to enhanced reconstruction results. Separately, data on the stable combustion process of a hydrogen-fueled scramjet engine were collected in a direct-connect supersonic pulse combustion wind tunnel, and these data are used to validate the robustness and generalization capability of MCSwinT. The experimental results show that the flow-field wave patterns of stable combustion are reconstructed successfully using MCSwinT of different scales. Compared to other models, MCSwinT exhibits lower model complexity while achieving performance improvements of 7% and 17% in peak signal-to-noise ratio and structural similarity index, respectively. Additionally, the high generalization ability of the proposed model is validated in a sparsity experiment. This model effectively reconstructs the flow-field wave patterns of stable combustion, providing a crucial foundation for further research on scramjet engines. [ABSTRACT FROM AUTHOR]

We develop an asymptotic theory of compressible turbulent boundary layers on a flat plate, in which the mean velocity and temperature profiles can be obtained as exact asymptotic solutions of the boundary-layer and energy equations, which are closed using functional relations of a general form connecting the turbulent shear stress and turbulent enthalpy flux to mean velocity and enthalpy gradients. The laws of the wall for velocity and temperature are constructed in the form of expansions in a small parameter that is proportional to the Mach number formed with the friction velocity and the speed of sound on the wall. The leading term of the expansion for velocity coincides with the Van Driest formula; however, the law of the wall also contains a term of order one, the presence of which explains the discrepancy between the Van Driest formula and experimental and calculated data. The formulation of the law of the wall for temperature takes into account the fact that in the case of a cooled wall, the temperature varies non-monotonically across the boundary layer and has a local maximum in the logarithmic sublayer. Along with the constants known for incompressible flow, the theory contains three new universal constants, which are determined from a comparison with direct numerical simulation data for velocity and temperature. [ABSTRACT FROM AUTHOR]

The compressibility effect of flow over a circular cylinder has been investigated using wall-resolved large eddy simulation. The Reynolds number in terms of the freestream quantities and the cylinder diameter is fixed at 3900 while varying the freestream Mach number from 0.01 to 0.5. Mesh quality, numerical scheme, and subgrid-scale model are carefully verified prior to the detailed flow study. Results such as Mach number effects on the drag coefficient, the shape of the mean streamwise velocity profile in the near-wake, the length of the recirculation zone, and shear layer instability are provided. Although compressibility suppresses the Kelvin–Helmholtz instability and mixing process within the shear layer, it increases the velocity fluctuation in the boundary layer and the pressure difference between the freestream and recirculation zone, which intensifies the wake oscillation, shedding amplitude, thus shortens the recirculation zone significantly with the Mach number up to 0.5. This phenomenon is different from that found in lower Reynolds number cases where the length of the recirculation zone elongates as the Mach number increases. The deficit of the velocity profile shape in the near-wake depends on the length of the recirculation zone. Furthermore, inside the wake zone and near the cylinder back wall, two pairs of recirculation bubbles emerge as the Mach number increases. [ABSTRACT FROM AUTHOR]

Among the design requirements of hypersonic vehicles, reducing aerodynamic heating and drag force simultaneously is the main challenge. This paper proposes a novel spike aerodome-channel combination concept to realize the flow field reconstruction around the hypersonic blunt body. The novel configuration is investigated in the axisymmetric flow at a Mach number of 6 at zero angle of attack. The two-dimensional Reynold-averaged Navier–Stokes equations are numerically solved, and the shear-stress transport k–ω model is the turbulence model implemented in this study. Parameters such as spike length and lateral jet location are investigated to explore the drag and heat reduction performance and the flow control features. The obtained results show that the application of the novel spike aerodome-channel concept alters the flow field by eliminating or replacing the strong bow shock wave, and the design of hypersonic vehicles can benefit from the application of the proposed concept. The blunt body coupled with a frustum of cone-tipped spike-channel configuration provides a remarkable drag reduction effect of 20.71% with respect to the case without channel. Considering the effect of lateral jet location, the drag reduction performance of the case with LR = 0.75 is superior to that of the root jet case at the same spike length, and a considerable drag reduction of 28.93% is obtained with L/D = 2.4. In addition, longer spike length is beneficial for improving drag reduction performance, while excellent efficiency of heat protection is obtained in a certain spike length range. For the case of L/D = 1.6 with root jet, the peak Stanton number is significantly decreased by 33.51%. [ABSTRACT FROM AUTHOR]

Direct numerical simulations of compressible nonisothermal turbulence in a periodic box with up to 4096 3 grid points were conducted by varying Reynolds numbers and resolution levels. The results were compared with those of compressible isothermal turbulence by Sakurai and Ishihara ["Direct numerical simulations of compressible isothermal turbulence in a periodic box: Reynolds number and resolution-level dependence," Phys. Rev. Fluids 8, 084606 (2023)] to study the effect of isothermal assumption on turbulence statistics. The turbulent Mach number and ratio of the dilatational to solenoidal root mean square velocities were fixed at approximately 0.3 and 0.4, respectively. A comparison under approximately equal flow conditions showed that the dilatational component of the energy spectra for the nonisothermal case exhibited approximately equal k − 3 scaling at k η > 1 as observed for the isothermal case and was consistently smaller in the wavenumber range 0.05 < k η < 0.6 than that for the isothermal case, where η is the Kolmogorov length. The dilatational energy is mainly dissipated around k η ≈ 0.3 , the same wavenumbers as the solenoidal energy dissipation irrespective of the isothermal assumption. As the Reynolds number increased, the dilatational energy dissipation caused by shocklets around k η ≈ 2 became larger, especially in the nonisothermal case. It was found that the isothermal assumption weakened the intermittency of the velocity divergence. No significant differences were observed in the normalized mean energy dissipation rates and pressure statistics. The local flow topology was also marginally affected by the isothermal assumption; however, the difference was significantly less than the changes owing to the different values of parameters such as the Reynolds number and Mach number. [ABSTRACT FROM AUTHOR]

The canonical shock/isotropic turbulence interaction (SITI) at high shock Mach numbers (Ms) is studied by conducting direct numerical simulation (DNS) for thermally perfect gas (TPG) and calorically perfect gas (CPG). Combining DNS with linear interaction analysis (LIA), the amplification of vorticity variance across the shock wave is studied. It is found that the changes in vortical velocity fluctuation amplitude and turbulent length scales under vibrational excitation have a competitive effect on vorticity amplification. The latter is dominant and leads to the transverse vorticity amplification increasing by 32.2% at Ms = 6.0. Based on the LIA theory, a vorticity amplification model for SITI considering vibrational excitation is established. Furthermore, the impact of vibrational excitation on the downstream vorticity transportation is examined through an analysis of the transport equation. The vibrational excitation strengthens both the vortex stretching and viscous dissipation of streamwise vorticity but only alters the viscous dissipation of transverse vorticity. Then, the vorticity transportations of different turbulent structures for CPG and TPG are compared. The comparison indicates that the increment of vortex stretching for streamwise vorticity variance is sustained by the enhanced turbulent structures corresponding to the stable-node/saddle/saddle, and the rapid decay of transverse vorticity variance for TPG is associated with the enhanced viscous dissipation of the nonfocal turbulent structure. [ABSTRACT FROM AUTHOR]

Wall pressure fluctuations caused by turbulent boundary layers have a significant impact on aircraft structural vibration and cabin noise. This study aims to investigate the mechanism of turbulence-induced pressure fluctuations by focusing on the randomness of wall pressure fluctuations, analyzed in both the time–frequency and spatial-wavenumber domains using measured data obtained from a phase array in a wind tunnel. Three roughness elements were designed and installed upstream of the plate to manipulate the turbulent boundary layer at a specific Mach number. The results of the investigation demonstrate that the disturbance strength induced by the roughness element influences the randomness of wall pressure fluctuations, in addition to the parameters utilized for data analysis. Generally, stronger turbulence fluctuations tend to decrease the randomness of pressure fluctuations. Moreover, wall pressure fluctuations also exhibit certain statistical principles that cannot be precisely calculated using mathematical expressions, highlighting their inherent randomness. Further investigation into randomness in the spatial-wavenumber domain revealed the hydrodynamic modes of turbulence fluctuations with varying convection velocity analyzed through wavenumber maps computed using the beamforming algorithm. These modes with variable convective speed significantly contribute to the generation of randomness in wall pressure fluctuations. Both the time–frequency domain and the spatial-wavenumber domain affect the randomness characteristics of wall pressure fluctuations. However, such effects are not easily discernible through a rudimentary analysis of the space–time correlation of turbulence fluctuations. [ABSTRACT FROM AUTHOR]

The linear spatial stability of a parallel two-dimensional (2D) compressible boundary layer on an adiabatic plate is investigated by considering both 2D and three-dimensional (3D) disturbances. The compound matrix method is employed here, for the first time, for compressible flows, which, unlike other conventional techniques, can efficiently eliminate the stiffness of the equations governing the spectral amplitudes. The method is first validated with published results in the literature corresponding to spatial and temporal instability of flows ranging from low subsonic to high supersonic Mach numbers (M), which shows a good match depending upon the proper choice of free-stream temperature and the wall dispersion relation. Subsequently, flow compressibility effects and the spanwise variation of disturbances are also investigated for M ranging from low subsonic to high supersonic cases (from M = 0.1 to 6). Mack (AGARD Report No. 709, 1984) reported the existence of two unstable modes for M > 3 from viscous calculations (the so-called "second mode") that subsequently fuse to create only one unstable zone when M increases. Our calculations show a series of higher-order unstable modes for M > 3 in addition to the findings of Ma and Zhong ["Receptivity of a supersonic boundary layer over a flat plate. Part 1. Wave structures and interactions," J. Fluid Mech. 488, 31–78 (2003)], where such higher-order modes for supersonic boundary layers are all noted to be spatially stable. The number and the frequency extent of the corresponding unstable zones increase with an increase in M beyond 3 while propagating downstream at a higher speed than those corresponding to incompressible, subsonic, and low supersonic (M < 2) cases. [ABSTRACT FROM AUTHOR]

In order to investigate the effects of the throat heights of the air annular slot on the rotating detonation waves, experiments have been carried out while varying the throat height of the air annular slot (i.e., 0.5, 1.5, 2.5, and 4.0 mm), the mass flow rate of oxidizer (35–300 g/s), and the equivalence ratio (0.15–1.8). Air and ethylene have been utilized as oxidizer and fuel, respectively. The experimental results indicate that rotating detonation waves can be obtained under different inflow conditions, and two operating phenomena were observed, i.e., the detonation cases (the single-wave mode and the oscillation mode) and the deflagration cases (the deflagration mode and the chaotic mode). The effects of the throat height on the operating range of stable rotating detonations are related to the minimum mass flow rate of oxidizer. As the throat height increases, the minimum mass flow rate of oxidizer also increases. Furthermore, the relationship between the minimum Mach number to obtain rotating detonations at each throat height and the throat height has been obtained. Moreover, the effects of the eccentricity ratio of the air annular slot on the propagation modes have also been demonstrated. [ABSTRACT FROM AUTHOR]

Wedge tail airfoils refer to the addition of a wedge section to the tail of an airfoil. This modification has demonstrated greater efficiency in increasing the lift forces compared to the base airfoil design. Such airfoils are particularly important in scenarios where increased lift forces are required, such as in the fast and efficient maneuvering of marine and aerospace vehicles. This study aimed to predict the flow fields and aerodynamic coefficients of wedge tail airfoils using a multi-head perceptron (MHP) network and classical machine learning (ML) algorithms, including k-nearest neighbors, decision tree, and random forest. These predictions were based on airfoil sections, x–y grid coordinates, Mach number, and angle of attack, eliminating the need to solve the Navier–Stokes (NS) equations. The database required for training the MHP network and the ML models were generated using the Ansys solver, which solved the NS equations. The results of the models were compared, and it was observed that the MHP network consistently outperformed the others in terms of prediction accuracy. Additionally, the prediction process demonstrated a significant speed improvement of 125 times compared to conventional computational fluid dynamics techniques. Once the training was completed, the flow fields and aerodynamic coefficients for a wedge tail airfoil could be obtained within seconds. [ABSTRACT FROM AUTHOR]

This paper investigates the aerodynamic performance of a system of multiple vehicles (three) propelled by an electric fan/compressor system in a closed low-pressure environment. The study was conducted on a 2-D axisymmetric model of multiple vehicles traveling at the same cruising speed inside a tube. The flow field was computed using the unsteady RANS model. The results demonstrate a significant improvement in aerodynamic performance when using a multiple-vehicle system equipped with an air-breathing electrical engine. The combined use of multiple-vehicle and the air-breathing propulsion system eliminated the piston effect at low Mach numbers and reduced transonic range effects at higher Mach numbers. Consequently, a considerable enhancement in aerodynamic performance over the single-vehicle model was attained. A minimum of 11% and a maximum of 57% reduction of further in average system drag was achieved for M = 0.3 and 0.7, respectively. An increase in thrust was also observed in the multiple-vehicle system over the single-vehicle system. The results of the study demonstrated the advantage of using an air-breathing multiple-vehicle system as an efficient and environmentally friendly high-speed transportation system. [ABSTRACT FROM AUTHOR]