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

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

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

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

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

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

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

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

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

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

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

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

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

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]

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]

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]

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 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]

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]

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 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]

Upstream propagating waves are observed to correlate with the reflected ions in the foot of a high Mach number quasi‐perpendicular collisionless shock. The respective wavefronts show time varying amplitude matching the variation of the reflected ions and the shock reformation cycles. We interpret this process in terms of the fast two‐stream instability and investigate the results of the PIC simulation as compared with the prediction of the linear analysis. Evidence for upstream propagating whistler waves supported by two‐stream instability is continuously growing, see for example, recent surveys based on MMS in‐situ observations of the Earth bow‐shock. The present study suggests that upstream propagating waves are consistent with and in particular supported by the mechanism of fast two‐stream instability developing in the upstream region of the shock foot. Plain Language Summary: By using particle‐in‐cell simulations, a highly non‐stationary quasi‐perpendicular collisionless shock exhibits reformation cycles characterized by extensions and compressions of the foot, ramp, and overshoot. Waves with upstream propagation are crossing the foot, with large variation of their phase velocities with respect to the rest frame of the shock, repeated periodically. Additionally, strong downstream propagating waves driven by incoming ions ‐ incoming electrons two‐stream instability (or slow two‐stream instability) are observed close to the end of each shock cycle. In a system attached to the electron flow (regarded as proxy for the system used in the study of two‐stream instabilities), the variations of the phase velocities appear to be significantly smaller. The linear theory based on dispersion analysis supports, in our case, the use of a 1‐D code for exploring upstream whistler propagation driven by the fast two‐stream instability, associated by former studies with 2‐D configuration. At the same time, the analytical examination of the real part of the dispersion relation shows good agreement with the simulation results, while the imaginary part indicates a significant local growth rate of the fast two‐stream instability, that can presumably compensate the otherwise strong damping of the oblique whistler waves. Key Points: Fast two‐stream instability excited in the foot of a collisionless shock by using 1‐D PIC simulationsCyclic emissions of whistler waves in the foot of the shocks are supported by fast and slow two‐stream instabilityConvergence of three independent approaches: simulations, analytical, and observations (MMS) [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]

This study presents a self-consistent one-temperature approach for modeling shock waves in single-component methane. The rigorous mathematical model takes into account the complex structure of CH4 molecules with multiple vibrational modes and incorporates exact kinetic theory-based transport coefficients, including bulk viscosity. The effects of the bulk viscosity on gas-dynamic variables and transport terms are investigated in detail under varying degree of gas rarefaction. It is demonstrated that neglecting bulk viscosity significantly alters the shock front width and peak values of normal stress and heat flux, with the effect being more evident in denser gases. The study also evaluates limitations in the use of a constant specific heat ratio, revealing that this approach fails to accurately predict post-shock parameters in polyatomic gases, even at moderate Mach numbers. To enhance computational efficiency, a simplified approach based on a reduced vibrational spectrum is assessed. The results indicate that considering only the ground state leads to substantial errors in the fluid-dynamic variables across the shock front. Another approach explored involves the application of machine learning techniques to calculate vibrational energy and specific heat. Among the methods tested, the Feedforward Neural Network (FNN) proves to be the most effective, offering significant acceleration in calculations and providing one of the lowest errors. When integrated into the fluid-dynamic solver, the FNN approach yields nearly a three-fold increase in speed in numerical simulations of the shock wave structure. [ABSTRACT FROM AUTHOR]

We investigate and characterize the effect of compressibility and rarefaction on vortex structures in the benchmark lid-driven cavity flow. Direct numerical simulations are performed, employing the unified gas kinetic scheme to examine the changes in vortex generation mechanisms and the resulting flow structures at different Mach and Knudsen numbers. At high degrees of rarefaction, where inter-molecular interactions are minimal, the molecules mainly collide with the walls. Consequently, the dominant flow structure is a single vortex in the shape of the cavity. It is shown that increasing compressibility or decreasing rarefaction lead to higher molecular density in the cavity corners, due to more frequent inter-molecular collisions. This results in lower flow velocities, creating conditions conducive to the development of secondary and corner vortices. The physical processes underlying vortex formations at different Knudsen numbers, Mach numbers, and cavity shapes are explicated. A parametric map that classifies different regimes of vortex structures as a function of compressibility, rarefaction, and cavity shape is developed. [ABSTRACT FROM AUTHOR]

The influence of outer large-scale motions (LSMs) on near-wall structures in compressible turbulent channel flows is investigated. To separate the compressibility effects, velocity fluctuations are decomposed into solenoidal and dilatational components using the Helmholtz decomposition method. Solenoidal velocity fluctuations manifest as near-wall streaks and outer large-scale structures. The spanwise drifting of near-wall solenoidal streaks is found to be driven by the outer LSMs, while LSMs have a trivial influence on the spanwise density of solenoidal streaks, consistent with the outer LSM impacts found in incompressible flows (Zhou et al. , J. Fluid Mech. , vol. 940, 2022, p. A23). Dilatational motions are characterized by the near-wall small-scale travelling-wave packets and the large-scale parts in the outer region. The streamwise advection velocity of the near-wall structures remains at $16 \sim 18u_{\tau }$ , hardly influenced by Mach numbers, Reynolds numbers and wall temperatures. The spanwise drifting of near-wall dilatational structures, quantified by the particle image velocimetry method, follows a mechanism distinct from solenoidal streaks. This drifting velocity is notably larger than those of the solenoidal streaks, and the influence of outer LSMs is not the primary trigger for this drifting. [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 scaling relations mapping the turbulence statistics in compressible turbulent boundary layers (TBLs) onto their incompressible counterparts are of fundamental significance for turbulence modelling, such as the Morkovin scaling for velocity fields, while for pressure fluctuation fields, a corresponding scaling relation is currently absent. In this work, the underlying scaling relations of pressure fluctuations about Mach number ($M$) contained in their generation mechanisms are explored by analysing a series of direct numerical simulation data of compressible TBLs over a wide Mach number range $(0.5\leq M \leq 8.0)$. Based on the governing equation of pressure fluctuations, they are decomposed into components according to the properties of source terms. It is notable that the intensity of the compressible component, predominantly originating from the acoustic mode, obeys a monotonic distribution about the Mach number and wall distance; further, the intensity of the rest of the pressure components, which are mainly generated by the vorticity mode, demonstrates a uniform distribution consistent with its incompressible counterpart. Moreover, the coupling between the two components is negligibly weak. Based on the scaling relations, semiempirical models for the fluctuation intensity of both pressure and its components are constructed. Hence, a mapping relation is obtained that the profiles of pressure fluctuation intensities in compressible TBLs can be mapped onto their incompressible counterparts by removing the contribution from the acoustic mode, which can be provided by the model. The intrinsic scaling relation can provide some basic insight for pressure fluctuation modelling. [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 burning in a two-section channel with an abrupt enlargement of its cross section in the process of double-belt supply of different-type gaseous fuels into it at a Mach number of the fuel flow M = 2.2 and at its stagnation temperature of 1700 K was numerically investigated. For the ignition of a hydrocarbon fuel fed into the channel along its axis (the first supply belt) and for the formation of a transonic regime of burning of the fuel, it was subjected to a gasdynamic pulse action, and, after this action was terminated, for the maintenance of the burning regime, the second belt of supply of a fuel (hydrogen or a hydrocarbon fuel) into the channel and the side supply of an air jet into it were organized. The Reynolds-averaged Navier–Stokes equations were solved with the use of the k–ε model of turbulence. The burning of a fuel in the channel was simulated by gross reactions. It was established that the use of hydrogen as a fuel in the second supply belt makes it possible to maintain the transonic burning regime in the channel. In this case, the side air jet breaks the shock-wave formation in the second section of the channel and, in so doing, provides a high-velocity homogeneous fuel flow in it. [ABSTRACT FROM AUTHOR]

Shocks are ubiquitous in astrophysical environments, and particle acceleration at such astrophysical shocks is related to high-energy phenomena. In particular, the acceleration mechanism and the time evolution of the particle distribution function have been extensively examined. This paper describes a simple analytic method using the one-dimensional Fokker-Planck equation in the test-particle regime. We aim to investigate the evolution of the particle distribution function in the shock upstream, which could be streaming toward Earth along the open magnetic field geometry. The behavior of the analytical solution is examined over a wide range of parameters representing shock structure, such as the shock Mach number, plasma beta, injection fraction into diffusive shock acceleration, and the scale of the upstream magnetic field. The behavior is associated with upstream turbulence for diffusive shock acceleration, as expected. Additionally, pre-accelerated particles could affect the time evolution of the particle distribution only when the radiative or advection losses are small enough for the pre-accelerated distribution to have a flatter power-law slope than the power-law slope based on shock acceleration theory. We also provide a formula for a spherically expanding shock and its relevant application to calculate high-energy emission due to hadronic interactions. We suggest that the simple analytic method could be applied to examine astrophysical shocks with a wide range of plasma parameters. [ABSTRACT FROM AUTHOR]

This study experimentally determines high-frequency aeroacoustic noises of the high-speed dual-impinging jet, focusing on the generation and evolution mechanism between impingement tones modes and far-field acoustic spectra. The variables of the high-speed impinging jet are Mach number (Ma = 0.9 and 1.1), fixed diameter (D) of nozzle, nozzle spacing (3.5D), and impingement distance (4D). A novel fast-responding pressure-sensitive paint (fast-PSP) with a significantly extended frequency response ability was designed to develop an accurate phase-resolving propagation process of aeroacoustic noises and handle high impingement momentum challenges posed by the high-speed impinging jet. The PSP raw data were enhanced by calibration, image restoration, and proper orthogonal decomposition filtering. Two distinct far-field spectral characteristics were identified based on synchronized acoustic measurements. The existence of the stagnation region affected by the fountain effect in the dual-impinging jet was determined using spatial–temporal cross-correlation analysis. Subsequently, the concurrent axisymmetric dual annulus mode and its coupling behavior at Ma = 0.9 and the non-axisymmetric helical mode and its periodic fragmentation-reconstruction patterns at Ma = 1.1 were identified by spectral proper orthogonal decomposition. Finally, the spatial–temporal evolution of the phase-locked first-order mode was extracted, and the transient variations of coherent structures and their mechanisms for discrete tone noise generation were quantitatively investigated. The expansion and coupling of the dual annulus modes promoted the dominance of single-tone peaks across the entire acoustic spectrum. The helical mode, exhibiting both rotational and expansion behaviors, enhanced the coherent vorticity at the periphery of the coherent structures, resulting in intense impingement tones sound sources. [ABSTRACT FROM AUTHOR]

The present study applies a framework of the spatiotemporal superresolution measurement based on the total-least-squares dynamic mode decomposition, the Kalman filter and the Rauch-Tung-Striebel smoother to an axisymmetric underexpanded supersonic jet of a jet Mach number of 1.35. Dual planar particle image velocimetry was utilized, and paired velocity fields of the flow with a short time interval were obtained at a temporal resolution of 5000 Hz. High-frequency acoustic data of 200,000 Hz were simultaneously obtained. Then, the time-resolved velocity fields of the supersonic jet were reconstructed at a temporal resolution of 200,000 Hz. Also, time coefficients of dynamic modes in high temporal resolution were calculated. The correlation between time coefficients implies that the mixing promotion by screech tone causes the lift-up of the high-velocity fluid from the jet center and accelerates at the downstream side. [ABSTRACT FROM AUTHOR]

In this study, the acoustic damping performances of the dual Helmholtz resonators were numerically evaluated using a 3D model. The grazing flow passes tangentially through the resonator neck, with a Mach number range of 0 ≤ Ma ≤ 0.1. The numerical model operates by solving the linearized Navier–Stokes equations. The current model is validated through a comparison with experimental data. The model is then utilized to explore the effects of the dual Helmholtz resonators on acoustic transmission loss performance in the presence of a grazing flow. Three key parameters are examined: 1) different implementation configurations of the dual Helmholtz resonators (including Models (b), (c), and (d)), 2) the mean temperature of the grazing flow, and 3) the axial distance between the dual Helmholtz resonators. For comparison, the acoustic damping performance of these dual Helmholtz resonators is compared to the single Helmholtz resonator case (Model (a)). The maximum transmission loss of Model (c) is significantly higher, recording values of 91%, 89.4%, and 92.5% than those observed for Model (a) at Ma = 0, Ma = 0.05, and Ma = 0.1, respectively. It is observed that the dual Helmholtz resonators dramatically increase the transmission loss. Model (c) is demonstrated to be associated with the most significant damping on the acoustic plane waves in comparison with that of Model (a). Additionally, the maximum transmission loss of Model (c) is 23.23 dB, 30.32 dB, and 34.58 dB at mean temperatures of 300 K, 600 K, and 900 K, respectively. Therefore, increasing the mean temperature is shown to be beneficial to enhance transmission losses in the presence of the grazing flow. Furthermore, under Ma = 0.1, the resonant frequency of Model (c) is 127 Hz, 152 Hz, and 172 Hz, corresponding to mean temperatures of 300 K, 600 K, and 900 K. It can be concluded that increasing the temperature has the effect of broadening the resonant frequency, especially at a high grazing flow Mach number. However, increasing the mean temperature results in a reduction of transmission loss in the absence of the grazing flow. In the case of Model (c), a 32 cm axial distance results in a 5.6% larger transmission loss at Ma = 0 and a 26.4% larger loss at Ma = 0.1 compared to a 16 cm axial distance. This indicates that increasing the axial distance between the dual Helmholtz resonators improves transmission loss. [ABSTRACT FROM AUTHOR]

Gurney flap is a simple and effective high-lift device that has the potential to reduce the takeoff-and-landing distance for transport aircraft. It is also a good option for enhancing the airfoil cruising characteristics. However, its effectiveness in non-continuum regimes has rarely been studied. The Unified Gas Kinetic Scheme algorithm, which is valid for all flow regimes, is further developed and used to simulate the flow field around a cambered airfoil and a high-speed quadrilateral airfoil with and without the Gurney flap in the supersonic non-continuum flow regimes. The distribution function shapes at different spatial locations are provided and analyzed. The effects of angle of attack, free-stream Knudsen number, free-stream Mach number, Gurney flap height, and Gurney flap mounting angle on the aerodynamic characteristics of the two airfoils are studied. The results show that the lift and drag of the airfoil with the attached Gurney flap are increased. The pressure rise on the lower surface of the airfoil is the primary factor contributing to the increase in lift. The overall increase in drag mainly comes from the Gurney flap, which also generates a small amount of negative lift. When the angle of attack is less than the critical angle of attack, the Gurney flap can increase the lift-to-drag ratio of the airfoil and improve its aerodynamic performance. The critical angle of attack decreases with the increase in the free-stream Knudsen number and the free-stream Mach number. It decreases with increasing Gurney flap height and increases with increasing Gurney flap mounting angle. [ABSTRACT FROM AUTHOR]

With advances in research on hypersonic vehicles, the precise simulation of the effects of thermochemical non-equilibrium has become increasingly important in their design. In light of this, this study explores the influence of high-temperature thermochemical non-equilibrium on the characteristics of the flow field around the hypersonic wings of an aircraft. We initially conducted a numerical simulation by using the model of flow through a cylinder to validate the accuracy and reliability of an 11-species gas model in representing high-enthalpy flow fields. Subsequently, a systematic analysis was conducted on the impact of thermochemical nonequilibrium effects on the temperature, pressure, and enthalpy distribution in the flow field around a symmetric diamond wing under different Mach numbers and angles of attack. The research results indicated that the deeper reason behind the differing thermochemical nonequilibrium effects in the flow field at various Mach numbers lies in the distinct distribution of enthalpy of the air components at different locations, which provided a new perspective for understanding flow field variations from the standpoint of enthalpy. It is disclosed that the thermochemical non-equilibrium significantly altered the characteristics of the flow field, particularly at high Mach numbers and angles of attack, with a significant impact on the aerodynamic parameters of both the windward and the leeward sides of the wing. Through explaining the mechanisms of thermochemical non-equilibrium under flow fields with different structures, this study provides a theoretical foundations for and a fresh perspective on the design of hypersonic vehicles. [ABSTRACT FROM AUTHOR]

As Mach number increases, thermochemical nonequilibrium is recognized as potentially affecting the flow field structure, as well as mixing and combustion characteristics, where shock-induced thermochemical nonequilibrium is a common and crucial phenomenon in compressible flow fields. A numerical study of shock-induced thermochemical nonequilibrium effects within a high Mach flow field of the electre vehicle is conducted by employing a two-temperature model-based solver hy2foam. The validation through experimental and simulation data confirms that hy2foam coupled with Park's two-temperature model and Park's five-species mechanism correctly predicts the flow structure and nonequilibrium characteristics. Four regime cases of thermochemical equilibrium, thermal nonequilibrium, chemical nonequilibrium, and thermochemical nonequilibrium are designed for comparison. First, the mechanism of shock-induced nonequilibrium is revealed. The shock induces the thermal nonequilibrium to occur instantly, and then the equilibrium is reestablished by undergoing the relaxation process. However, chemical nonequilibrium works delayed after the shock, and the high temperature induced by the shock motivates deviation from the chemical equilibrium by turning on chemical reactions. Further comparison of the four cases reveals that thermodynamic nonequilibrium significantly affects both shock position and intensity. In contrast, chemical nonequilibrium only significantly affects the distance to the shock detachment. Furthermore, it is found that thermodynamic and chemical nonequilibria behave in a complex coupling relationship after the shock. [ABSTRACT FROM AUTHOR]

Data-driven modal analysis methods provide a powerful way to decompose data into a sum of modes. The spatiotemporal Koopman decomposition (STKD) enables the computation of modes defined by global frequencies and growth rates in various spatial dimensions and time. The method is an extension of the dynamic mode decomposition (DMD) and higher-order dynamic mode decomposition (HODMD) that represents the data as a sum of standing and traveling, possibly growing or decaying, waves. In this paper, the STKD with HODMD is applied to schlieren video highlighting second mode instability waves traveling down the length of a 3-degree half-angle cone and a 7-degree half-angle cone, both at a freestream Mach number of 6. The HODMD is able to compute dominant modes and frequencies that align with those from associated experimental measurements of unsteady pressure fluctuations, and whose mode shapes clearly show the intensifying wavepacket structure of the waves. The STKD algorithm is used to compute streamwise wavenumbers, spatial growth rates, and wave speeds. The spatial growth rates from the STKD and the magnitudes of the HODMD mode shapes are used to compute the N-factor for waves of several frequencies. Overall, the STKD with HODMD is shown to be a useful tool for extracting spatiotemporal disturbance growth from a schlieren video. [ABSTRACT FROM AUTHOR]

The axisymmetric conical and curved shock waves are studied theoretically via the streamline transformation method (STM), which integrates the differential geometries of streamlines into the conservation laws and solves the flow field geometrically. Under the condition of a straight shock wave, the governing equation is significantly simplified, from which the assumptions of the conical flow and geometric self-similarity are mathematically proved. The simplified equation is also demonstrated as equivalent to the classical Taylor–MacColl equation. The flow mechanism after a conical shock wave is discussed from a geometric perspective, where the key factor is the dimensionless lateral distance between symmetric planes. The corresponding geometric and flow parameters, e.g., the distance between streamlines, the streamline curvature, and the Mach number, are computed for the demonstration of this mechanism. In the flow field over a circular cone, the isentropic compression is modeled geometrically. Due to the isentropic compression, the conical shock wave remains approximately a Mach wave at small semi-angles and the detachment is suspended at large semi-angles. For the axisymmetric curved shock wave, the flow field is computed directly by the streamline transformation method, where both the concave and convex shock waves are considered. Along the wall streamline, an explicit proportional relation between the shock curvature and the orthogonal derivatives of flow parameters is also provided and verified numerically. These results are applicable in the theoretical analysis and inverse design of axisymmetric shock waves. [ABSTRACT FROM AUTHOR]

During atmospheric reentry, the vehicle surface is exposed to highly non-equilibrium flow. The vehicle surface can experience heterogeneous recombination of reactive atoms, which contributes to its aerothermodynamic heating. This process is followed by chemical energy accommodation (CEA), where the released energy is either transferred to the surface or the internal energy modes of the recombined molecule. Heterogeneous recombination can be categorized into Eley–Rideal (ER) and Langmuir–Hinshelwood mechanisms, which differ in their methods of molecule formation and degrees of CEA. The complete CEA assumption may not consider the dependency of CEA on the mechanisms of heterogeneous recombination. This study aims to consider the mechanism-specific CEA for a more accurate prediction of surface heat flux. The authors implement mechanism-specific CEA within the direct simulation Monte Carlo framework using the finite-rate surface chemistry model, resolving elementary surface reactions and assigning a CEA coefficient, β, to each mechanism. The model is verified through comparisons with analytical solutions of surface coverage and validated against benchmark references. A parametric investigation of rarefied hypersonic flow over a two-dimensional cylinder is conducted under different freestream Mach and Knudsen numbers. Results show a reduction in total heat flux of up to 14.44% using mechanism-specific CEA compared to the complete CEA assumption. The reduction is attributed to the relative contribution of the ER mechanism, which can be a function of atomic partial pressure at the boundary layer. [ABSTRACT FROM AUTHOR]

The acceleration self-starting performance of hypersonic inlets is of critical importance for the stable operation of scramjet engines. The occurrence of soft unstart during the transition from hard unstart to start is an important flow state that has yet to be fully elucidated. The stability mechanism and corresponding self-starting characteristics of soft unstart remain poorly understood, and there is a pressing need for detailed modeling research in this area. This paper presents a rapid prediction model for the self-starting Mach number of two-dimensional hypersonic inlets with soft critical unstart mode, fully considering the influence of various geometric parameters and Reynolds number in the internal contraction section, and achieving a quantitative analysis of the two-dimensional soft unstart critical flow field. Given the incoming flow conditions and the inlet geometry, the prediction model is capable of accurately representing the actual viscous unstart flow field. It can fully map the unstart separation bubble and its surrounding critical wave structures, and calculate the minimum pressure rise required to maintain the current scale of the main separation bubble and the pressure rise exerted on the unstart separation bubble by the current actual flow field structure. Comparing the relative magnitude of these two pressures determines whether the inlet can transition from soft unstart to start. The proposed prediction model was validated using results from unsteady numerical simulations. The predicted results align well with the simulation results and are significantly better than previous prediction methods. [ABSTRACT FROM AUTHOR]

The interstage turbine burner can improve the compactness of the engine structure and increase the thrust–weight ratio of the engine without using afterburner. Achieving this goal requires stable combustion within a very short distance after the high-pressure turbine. It is necessary to study the flow field structure in the interstage turbine burner. In this study, numerical simulation was used to investigate the flow field structure and flow characteristics in the combustor. The influence of the inlet residual swirl and the radial distribution of the inlet flow velocity on the flow field structure, the proportion of the mainstream and secondary flows, and the flow loss of the combustor under different inflow conditions was obtained. The analysis of the numerical simulation results showed that the influence of inlet residual swirl on the total pressure loss coefficient showed a positive correlation trend, but within a range of 4500–6000 revolutions per minute (rpm), swirl would minimize the total pressure loss coefficient. Residual swirl has a significant impact on the strength and circumferential spatial distribution of primary and secondary vortices within the combustor, while the radial velocity distribution has a lower impact on the vortex structure within the cavity, but it has a certain influence on the total pressure loss coefficient and the proportion of the mainstream and secondary flows. At 0.315 Mach number (Ma), the total pressure loss of the faster inside radial distribution of the velocity is reduced by 1% compared to a uniform inlet. [ABSTRACT FROM AUTHOR]

This study numerically investigates the mechanisms for generating and suppressing the leading edge vortex (LEV) in a vaned diffuser of a centrifugal compressor, aiming to enhance the compressor's stable operating range. Detailed analyses focus on the mechanisms of the rotating stall and the initiation process of the LEV. Notably, the end wall contouring method is applied to vaned diffuser to suppress the LEV and improve diffuser stability. The suppression mechanisms of the LEV are examined to deepen the understanding of stability enhancement mechanisms. Results demonstrate that, under stall conditions, the vaned diffuser experiences two-cell rotating stalls moving opposite to the impeller rotation direction, each rotating at approximately 12% of impeller rotation speed. The formation of diffuser stall is attributed to a longitudinal vortex developing from the LEV and causing the passage blockage. The increasing spanwise distortion of leading edge (LE) incidence angles and circumferential distortion of static pressures at the diffuser inlet promote the evolution of the longitudinal vortex from LEV. The end wall contouring of vaned diffuser effectively enhances the compressor's stable operating range by 15.10% and 8.90% toward lower flow rate conditions for machine Mach numbers 1.3 and 1.2, respectively. At near-stall points, the end wall contoured diffuser reduces spanwise distortion of LE incidence angles and circumferential distortion of static pressures at the diffuser inlet, mitigates the LE flow separation and corner separation, suppresses the generation and development of the LEV, delays the onset of rotating stall in vaned diffuser, and thereby improves the stable operating range of centrifugal compressor stage. [ABSTRACT FROM AUTHOR]

Wall pressure fluctuations beneath turbulent boundary layers are a fundamental source of aerodynamic noise by exciting the wall structure, with their space-time characteristics serving as the basic ingredient for predicting the wall structural response. To this end, direct numerical simulations of fully developed compressible turbulent boundary layers at Mach numbers of 0.5, 1.2, and 2.0 are conducted to investigate wall pressure fluctuations comprehensively. The effects of Mach number on the single-point statistics of wall pressure fluctuations, such as the root mean square, skewness and flatness factors, probability density function, and frequency spectrum, are assessed to be very weak. Regarding the space-time characteristics, the convection velocity Uc determined by the space-time correlation of wall pressure fluctuations increases slightly with the Mach number, which only reflects the convective behavior of turbulent vortices. On the wavenumber–frequency spectrum, characteristic peaks of both the acoustic wave and convective vortices are identified. At Mach 0.5, the peaks of the fast (U c + c) and slow (U c − c) acoustic waves are unattached to others with c denoting acoustic speed, while only the peak of the fast acoustic wave is distinguishable from the convective peak at Mach 1.2 and 2.0. Due to the aerodynamic heating at supersonic conditions, the thermal effect on acoustic speed should be taken into account in determining the acoustic wavenumber. By introducing a convective Prandtl–Glauert parameter, a refined relation is proposed to provide a more accurate depiction of the acoustic domain in the wavenumber–frequency spectrum. [ABSTRACT FROM AUTHOR]

Semi-local scales have been widely used in compressible wall-bounded turbulence, but it is still unclear whether they are applicable to the scaling of coherent structures, especially under conditions of high Mach number and cold wall temperature. By scrutinizing the direct numerical simulation dataset at different Mach numbers and wall temperatures, this paper demonstrates that the coherent structures normalized by semi-local scales are universal in size. In addition to this, we find that the ratios of Kolmogorov scales to semi-local scales are independent of Mach number and wall temperature. Thus, Kolmogorov scales can achieve the same scaling effect as the semi-local scales. The velocity spectra are also compared to verify the current scaling method quantitatively. A method to determine the threshold for the vortex identification criterion is proposed, allowing the same threshold for different cases to obtain vortices of similar size. The scaling of other statistics including turbulent kinetic energy, streamwise Reynolds normal stress, and root mean square of fluctuating vorticity is also investigated. A new velocity scale is proposed based on the total-stress-based transformation for mean streamwise velocity, which can collapse the profiles of these statistics more accurately than the semi-local velocity scale. The present paper demonstrates that through appropriate normalization, the structures and statistics of compressible turbulence become universal, reaffirming the validity of Morkovin's hypothesis even for the present high Mach number and cold wall cases. [ABSTRACT FROM AUTHOR]

Advancements have been achieved in the optimization of waverider designs with the aid of machine learning to expedite the design process. However, these approaches are hampered by the need for extensive sample sizes and susceptibility to becoming ensnared in local optima. This study undertakes a parametric design based on the wedge-derived, power-law-shaped waverider, increasing configuration diversity and creating a dataset with limited samples by calculating waverider geometry and aerodynamic parameters. At a Mach number of 10, a multi-objective optimization design is implemented using the Young's double-slit experiment-least squares support vector regression (YDSE-LSSVR) surrogate model in conjunction with improved congestion distance multi-objective particle swarm optimization algorithm, focusing on maximizing the lift-to-drag ratio and volumetric efficiency as much as possible. The results indicated that, under conditions of limited samples, the YDSE-LSSVR model outperforms standard models such as support vector regression, LSSVR, Kriging, and Polynomial Chaos Expansions-Kriging regarding prediction accuracy. The Pareto solutions for both concave and convex waveriders, obtained through multi-objective optimization, improve the lift-to-drag ratio by 17.36% and 21.70%, respectively, and increase the volumetric efficiency by 88.89% and 105.56%, in comparison to baseline configurations. In addition, the research examines the impact of various design parameters on the Pareto solutions. Finally, the study applies the K-means method to conduct a cluster analysis of the Pareto solutions, generating three-dimensional waverider configurations based on distinguished solutions from different clusters. [ABSTRACT FROM AUTHOR]

Shock wave propagation in gases through turbulent flow has wide-reaching implications for both theoretical research and practical applications, including aerospace engineering, propulsion systems, and industrial gas processes. The study of normal shock propagation in turbulent flow over non-ideal gas investigates the changes in pressure, density, and flow velocity across the shock wave. The Mach number is derived for the system and explored across various gas molecule quantities and turbulence intensities. This study analytically investigated the normal shock wave propagation in turbulent flow of adiabatic gases with modified Rankine–Hugoniot conditions. Artificial neural network (ANN) techniques are used to estimate the solutions for shock strength and Mach number training validation phases of back-propagated neural networks with the Levenberg–Marquardt algorithm. The results reveal that pressure ratio with density ratio increase for higher values of increase in the turbulence level as well as intermolecular forces. A reverse trend is observed in velocity coefficient after shock in the presence of adiabatic gas. The regression coefficient values obtained using the network model ranged from 0.999 99 to 1, indicating an almost perfect correlation. These findings demonstrate that the ANN can predict the Mach number with high accuracy. [ABSTRACT FROM AUTHOR]

Computation of impulsively started flows presents difficulties associated with the presence of a singularity at time equal to zero. When using the lattice Boltzmann method, the standard practice is to start the computation from a potential flow field that is not part of the solution. A different approach to the problem is presented in this article where three new criteria for the selection of computational parameters in highly unsteady flow environments are presented. These criteria, which do not overrule the conventional one that sets limits to the computational Mach number, are based on fluid physics considerations. They represent additional constrains related to (a) the distance traveled by sound waves at early times, (b) the importance of viscous length during the onset of impulsive motion, and (c) the presence of spurious reflected pressure waves at the beginning of computations. The proposed methodology was tested in the case of an impulsively started cylinder, and the results were compared to those of analytical, numerical, and experimental nature published in specialized literature. It is intended that this study facilitates the computation of highly unsteady flows for researchers who use the lattice Boltzmann method. [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]