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We propose the results of a theoretical and mathematical model of the growing bubbles at low Mach number in a generalized Newtonian fluid. The system occurs under the effect of magnetic field and s...

The reattached boundary layer in the interaction of an oblique shock wave with a flat-plate turbulent boundary layer at Mach number 2.25 is studied by means of Direct Numerical Simulation (DNS). The numerical results are carefully compared with available experimental and DNS data in terms of turbulence statistics, wall pressure and skin friction. The coherent vortex structures are significantly enhanced due to the shock interaction, and the reattached boundary layer is characterized by large-scale structures in the outer region. The space-time correlation of fluctuating wall shear stress and streamwise velocity fluctuation reveals that the structural inclination angle exhibits a gradual decrease during the recovery process. The scale interactions are analyzed by using a two-point amplitude modulation correlation. A possible mechanism is proposed to account for the strong amplitude modulation in the downstream region. Moreover, the mean skin-friction is decomposed to understand the physically informed contributions. Unlike the upstream Turbulent Boundary Layer (TBL), the contribution associated with the Turbulence Kinetic Energy (TKE) production is greatly amplified, while the spatial growth contribution induced by the pressure gradient largely inhibits skin-friction generation. Based on bidimensional empirical mode decomposition, the turbulence kinetic energy production contribution is further split into different terms with specific spanwise length scales.

The flow of a compressible fluid with slip through a cylinder with an asymmetric local constriction has been considered both numerically, as well as analytically. For the numerical work, a particle-based method whose dynamics is governed by the multiparticle collision (MPC) rule has been used together with a generalized boundary condition that allows for slip at the wall. Since it is well known that an MPC system corresponds to an ideal gas and behaves like a compressible, viscous flow on average, an approximate analytical solution has been derived from the compressible Navier–Stokes equations of motion coupled to an ideal gas equation of state using the Karman–Pohlhausen method. The constriction is assumed to have a polynomial form, and the location of maximum constriction is varied throughout the constricted portion of the cylinder. Results for centerline densities and centerline velocities have been compared for various Reynolds numbers, Mach numbers, wall slip values and flow geometries.

Gas atomization is now an important production technique for Fe-based amorphous alloy powders used in additive manufacturing, particularly selective laser melting, fabricating large-sized Fe-based bulk metallic glasses. Using the realizable k-e model and discrete phase model theory, the flow dynamics of the gas phase and gas-melt two-phase flow fields in the close-wake condition were investigated to establish the correlation between high gas pressure and powder particle characteristics. The locations of the recirculation zones and the shapes of Mach disks were analyzed in detail for the type of discrete-jet closed-coupled gas atomization nozzle. In the gas-phase flow field, the vortexes, closed to the Mach disk, are found to be a new deceleration method. In the two-phase flow field, the shape of Mach disk changes from “S”-shape to “Z”-shape under the impact of the droplet flow. As predicted by the wave model, with the elevation of gas pressure, the size of the particle is found to gradually decrease and its distribution becomes more concentrated. Simulation results were compliant with the Fe-based amorphous alloy powder preparation tests. This study deepens the understanding of the gas pressure impacting particle features via gas atomization, and contributes to technological applications.

Inverse design has long been an efficient and powerful design tool in the aircraft industry. In this paper, a novel inverse design method for supercritical airfoils is proposed based on generative models in deep learning. A Conditional Variational AutoEncoder (CVAE) and an integrated generative network CVAE-GAN that combines the CVAE with the Wasserstein Generative Adversarial Networks (WGAN), are conducted as generative models. They are used to generate target wall Mach distributions for the inverse design that matches specified features, such as locations of suction peak, shock and aft loading. Qualitative and quantitative results show that both adopted generative models can generate diverse and realistic wall Mach number distributions satisfying the given features. The CVAE-GAN model outperforms the CVAE model and achieves better reconstruction accuracies for all the samples in the dataset. Furthermore, a deep neural network for nonlinear mapping is adopted to obtain the airfoil shape corresponding to the target wall Mach number distribution. The performances of the designed deep neural network are fully demonstrated and a smoothness measurement is proposed to quantify small oscillations in the airfoil surface, proving the authenticity and accuracy of the generated airfoil shapes.

Implicit large-eddy simulations of reflecting shock-wave turbulent boundary-layer interactions at a freestream Mach number of 2.05 were carried out for sweep angles of 0, 20, and 40 deg. The simula...

In this paper a new hybrid semi-implicit finite volume / finite element (FV/FE) scheme is presented for the numerical solution of the compressible Euler and Navier-Stokes equations at all Mach numbers on unstructured staggered meshes in two and three space dimensions. The chosen grid arrangement consists of a primal simplex mesh composed of triangles or tetrahedra, and an edge-based / face-based staggered dual mesh. The governing equations are discretized in conservation form. The nonlinear convective terms of the equations, as well as the viscous stress tensor and the heat flux, are discretized on the dual mesh at the aid of an explicit local ADER finite volume scheme, while the implicit pressure terms are discretized at the aid of a continuous $\mathbb{P}^{1}$ finite element method on the nodes of the primal mesh. In the zero Mach number limit, the new scheme automatically reduces to the hybrid FV/FE approach forwarded in \cite{BFTVC17} for the incompressible Navier-Stokes equations. As such, the method is asymptotically consistent with the incompressible limit of the governing equations and can therefore be applied to flows at all Mach numbers. Due to the chosen semi-implicit discretization, the CFL restriction on the time step is only based on the magnitude of the flow velocity and not on the sound speed, hence the method is computationally efficient at low Mach numbers. In the chosen discretization, the only unknown is the scalar pressure field at the new time step. Furthermore, the resulting pressure system is symmetric and positive definite and can therefore be very efficiently solved with a matrix-free conjugate gradient method. In order to assess the capabilities of the new scheme, we show computational results for a large set of benchmark problems that range from the quasi incompressible low Mach number regime to compressible flows with shock waves.

The interaction between an incident shock wave and a wavy end-wall in a three-dimensional geometry is numerically simulated by using a high-order finite-difference solver with a ghost-cell immersed boundary method. The aim is to discover the differences of the unsteady propagation characteristics of triple bifurcation points and transverse waves at different incident shock Mach numbers ( $$M_{1})$$ and wavy wall amplitudes ( $$A_{\mathrm{ww}}$$ ). For a benchmark case with $$M_{1}=1.5$$ and $$A_{\mathrm{ww}}=1$$ mm, the simulated results are in a good agreement with other studies, indicating the reliability of the current simulation technique. The numerical results at $$M_{1}=1.5$$ , 1.9, 2.5, and 3.5 show that the Mach numbers of the transverse shock waves issued from the triple bifurcation points decrease with time according to a power function. It indicates that the deformation of the shock wave attenuates with time and its flat shape is gradually recovered. The stronger the incident shock wave is, the faster the deformation decays. In the central region, a petal-like vortex structure is observed near the wavy wall and its advancing speed with periodic fluctuation correlates with the cycle of the transverse motion of the triple bifurcation point. With the increase of $$M_{1}$$ , the petal-like vortex gradually grows up in size, and a faster rate can be observed in the normal direction. By comparing the propagation characteristics of transverse waves at different wavy wall amplitudes, it is discovered that the cellular pattern becomes more diverse as the wall amplitude increases. This is due to the multiple collisions of the transverse waves on the wavy wall, which leads to the multi-modal waves system in the shocked gas.

We present regime maps for subsonic flow in dense gas-particle systems, which demarcate regions of compressible and (effectively) incompressible flow. These maps should aid researchers and industrialists in selecting the appropriate modeling approach, as well as in verifying numerical solvers. Demonstrating compressibility at Mach numbers lower than 0.3, we show that this commonly used criterion is insufficient for flows in porous media. For M

The overset grid assembly (OGA) method is an effective computational fluid dynamics technique for the simulation of flows with complex geometries or relative moving parts. Using OGA, the effort for...

Strong laminar/transitional shock boundary layer interactions (SBLIs) have been investigated in a Mach 4 vacuum-driven wind tunnel with supporting computational fluid dynamics analysis. Such flows ...

The flow physics associated with shock-wave/boundary-layer interactions (SBLIs) on an axisymmetric body, simulating a projectile near an aircraft, is of significant interest to the aerodynamic comm...

This paper investigates aerodynamic heating on a flared cone in a Mach 6 wind tunnel. Two data processing methods to calculate the Stanton number and recovery factor simultaneously are presented, u...

This manuscript designed an axisymmetric variable geometry scramjet combustor equipped with a strut flame holder to help the combustor work in a wide flight Mach numbers and conducted a series of n...

The computational approaches of CFD are more powerful than the analytical solutions for high-speed compressible flows over a flat plate. A small number of expensive experimental data can be generated to aid high-speed vehicle design. However, CFD can be used to simulate a large variety of flows with different freestream and wall conditions in a cost-effective manner. The current work aims to numerically calculate the turbulent boundary layer flows over a flat plate at different Mach numbers in the range of 2–8 at different wall conditions and unit Reynolds numbers. The Reynolds-averaged Navier–Stokes method with k − ω turbulence model is applied to resolve the flow over the flat plate at zero angle of attack. The computed skin friction coefficient and Stanton number are compared with the available experimental data in the literature. The calculated results indicate some agreement with the experimental data. An increase in the Mach number and wall temperature decreases the skin friction and the Stanton number. A polynomial curve fit data estimation is proposed for skin friction under adiabatic wall conditions for Mach numbers in the range of 2–8. Numerical simulations over compression corner flows are also presented in the present work. The freestream Mach number influences the thickness of the subsonic layer in the undisturbed boundary layer on the flat plate in compression corner flows. The higher freestream Mach number results in a lower subsonic thickness near the wall, leading to a small separation bubble and lower peak heat transfer in shock/boundary-layer interaction flows.

Abstaining from a corrector step of SIMPLE-like algorithms, a coupled pressure-based algorithm is designed using a cell-centered finite-volume Δ -approximation on a collocated grid to predict all-speed fluid flows. A consistently formulated cell-face dissipation scheme is invoked to unravel the occurrence of pressure–velocity decoupling. Modified Rusanov, Roe and Lax–Wendroff flux-difference schemes accompanied by a newly devised differentiable slope limiter function are used to compute inviscid fluxes over the entire spectrum of subsonic to hypersonic flow. Numerical experiments in reference to incompressible and compressible flows demonstrate that the proposed algorithm maintains an excellent consistency with analytic solutions and available literature data. The method is anticipated to be stable for an extensive range of Mach numbers with an avoidance of under-relaxation (UR) factors.

Mach 4.5 tests in a conventional trans-supersonic wind tunnel are often accompanied by the air liquefaction phenomenon, resulting in the low reliability of test data. The Mach 4.5 nozzle developed in a hypersonic wind tunnel is able to heat airflow and provide more accurate test data. At present, China does not have the capability to test the Mach 4.5 nozzle in the 0.5-meter hypersonic wind tunnel. This gap may be filled by developing the Mach 4.5 nozzle in the hypersonic wind tunnel. The axisymmetric nozzle profile was calculated by the inviscid flow calculation method, and the boundary layer was modified by the Sivells-Payne method. Then, the numerical simulation was carried out, and the simulation results prove that the nozzle profile thus calculated meets the design requirements of the Mach number. For its structural design, a three-section design method is adopted to ensure the continuity and smoothness of the inner surface so as to better calibrate the flow field. Standard model tests were also carried out. The test results show that the velocity field of the Mach 4.5 nozzle we developed meets technical requirements. The standard model test data provide data reliable support for the development of aircraft.

Stable and efficient mass exchange is crucial to the combustion stability of a dual-mode scramjet. In subsonic mode, the progression of mass exchange between the cavity fluid and the mainstream, which is controlled by the cavity shear layer, is different from the scramjet mode due to the weak compressibility effects. However, mass exchange mode studies in the subsonic mode are rare. In this study, experimental research has been conducted to obtain the flow structure of Mach 0.3 and 0.5 inflow over a rectangular cavity and reveal the properties of mass exchange using particle image velocimetry technology. The results reveal the weak compressibility effects and geometry effects on the mass exchange by analyzing the growth rate and turbulence characteristics of the cavity shear layer. The mass exchange between the cavity fluid and mainstream is determined based on the vertical velocity along the cavity lip line. When the flow Mach number increased from 0.3 to 0.5, the growth rate of the cavity shear layer decreases due to the weak compressibility effects. Also, the growth rate of the cavity shear layer is smaller than that of the compressible free shear layer due to the effects of recirculation within the cavity. For length-to-depth (L/D) ratios ranging from 0.8 to 1.2, the cavity shear-layer growth rate decreases with increasing L/D. However, at an L/D of 1.5, the shear-layer growth rate increases, which changes the mass exchange mode. Finally, the effects of increasing the Mach number (Mach = 0.3 versus 0.5) show that the L/D ratio at the “crossover” point of the mass exchange mode transition is reduced (L/D = 1.5 versus 1.2) due to the weak compressibility at a higher Mach number.

This paper reports an active method for aerodynamic force control of a hypersonic spacecraft via shock wave mitigation with plasma energy deposition. The energy deposition was generated by quasi-direct current discharge at two positions in the streamwise direction on a flat plate upstream a compression surface. Wind tunnel experiments of visualizing a Mach 6 flow over a ramp with high-speed schlieren imaging were performed to verify the method's effect on diluting a shock wave. Meanwhile, the Reynolds averaged Navier-Stokes equations, with a source term of simulating plasma heating due to discharge, were solved to evaluate the method's control abilities. From the experimental results, low-density plasma layers can be formed by quasi-DC discharge, which were continuously weakening the shock wave induced by the ramp. The numerical results predicted flow topology under control well with the experimental schlieren images, with the corresponding force reduction rate as high as 58%, which mainly depends on the amount of energy injection, instead of arrangement of heating zones in this study.

Isolated rotating tail rudder technology provides a low-cost and miniaturized solution for the correction and guidance of a man-portable rocket. The results of three turbulence models (the Spalart–Allmaras model, standard k–e model, and shear stress transport k–ω model) were compared with wind-tunnel model test data, and the best turbulence model was selected. An aerodynamic model of the rotating tail rudder was developed by identifying its turbulent region, and the influences of the Mach number, angle of attack, and tail rudder speed on the projectile aerodynamics were revealed. The aerodynamic parameters were fitted using a least-squares method, and the vector variation characteristics of the period-averaged control force were analyzed. The results from the shear stress transport k–ω model were closest to the results of the wind-tunnel tests. The aerodynamic model was able to fit the simulation results well. The average control force of the tail rudder over a rotation cycle is not zero, and it increases with increasing angle of attack, Mach number, and tail rudder rotation speed. This study provides a basis for aerodynamic research examining the same type of projectile, and it has guiding significance for the control design of isolated rotating tail rockets.

The influence of the flat-plate leading-edge shape on laminar–turbulent transition is investigated. Experiments were carried out at a Mach number of M∞=5; a unit Reynolds number Re1∞ from 1.5×107 t...

The purpose of this research work is to perform accurate numerical computations of supersonic flow in a converging nozzle and specifically to study Mach-disks. The latter process has been widely studied over the last years. In the present study numerical simulations are performed for transient supersonic flow, tracing the transition from a Mach reflection to a regular one. This has been done by enforcing the walls of a converging nozzle to come closer together, changing the deflection angle with time. Viscosity was taken into account and the full Navier- Stokes have been solved. The results obtained clearly show the gradual extinction of the Mach disk and the eventual wave intersection to a single point

A single-stage axial fan having a pressure ratio of 1.01 is designed in the current study. The design pressure ratio is chosen based on the power available from the existing motor (2.2 kW). The design space for the axial flow fan was generated by varying specific flow and geometrical parameters in suitable steps, using a program written in MATLAB. The varied flow parameters are mass flow rate, inlet Mach number, inlet flow angle, and rotor speed. The geometrical parameters that were varied are hub to tip ratio, aspect ratio, and blade solidity. Using these as the input variables and applying free vortex theory for 3-dimensional blade design, the aerodynamic design of the axial flow fan was carried out. Performance parameters like flow coefficient, stage loading coefficient, degree of reaction, diffusion factor, De Haller’s number, and blade angles were calculated at the blade’s hub, mean, and tip. Total design space of 92160 data points was obtained from the combination of input parameters. Several constraints were applied to optimise the design space based on the available power from the existing motor and in-house manufacturing limitations. The initial design space was reduced to 82 data points using these constraints. To further reduce the number of points in the design space, off-design performance was evaluated for each of these data points. Following this, one design point was selected based on the optimum performance range in off-design operation, while considering manufacturing limitations. Using Mellor charts, a suitable blade profile was chosen based on the inlet and exit blade angles. NACA 65-410 airfoil was selected with a stagger of 55 degrees and an incidence of 6 degrees for optimum performance.

The results of a parametric, numerical study of aftbody exhaust flows on a conical, airbreathing, hypersonic vehicle are presented. Three-dimensional Euler simulations were performed at Mach 9 and ...

Turbulent jet flows at the Mach number equal to 0.25 from an industrial nozzle are studied using the large eddy simulation (LES). The feasibility of controlling the jet flow and noise by using chevrons is explored. A comparative study is carried out between the baseline and chevrons nozzles with the focus paid to the influences of the chevrons on the flow mixing and noise. The results show that the chevrons can significantly change the flow field. Especially the jet potential core length is reduced. In addition, the contours of velocity in the cross sections of the jet change from circular into daisy shaped structures by the chevrons. Flow visualization using the vorticity and the Q criterion shows that large vortical structures in the shear layer are fragmented into small-scale eddies. Meanwhile, the sound pressure level of the jet noises is weakened by chevrons.

The effects of static surface deformations on a spatially developing supersonic boundary layer flow at Mach number M = 4 and Reynolds number R e δ i n ≈ 49300 , based on inflow boundary layer thickness ( δ i n ), are analyzed by performing large eddy simulations. Two low-order structural modes of a rectangular clamped surface panel of dimensions ≈ 33 δ i n × 48 δ i n are prescribed with modal amplitudes of δ i n . The effects of these surface deformations are examined on the boundary layer, including changes in the mean properties, thermal and compressibility effects and turbulence structure. The results are analyzed in the context of deviations from concepts typically derived and employed for equilibrium turbulence. The surface deflections, to some degree, modify the correlations that govern both Morkovin’s hypothesis and strong Reynolds analogy away from the wall, whereas in the near-wall region both the hypotheses breakdown. Modifications to the turbulence structure due to the surface deformations are elucidated by means of the wall pressure two-point correlations and anisotropy invariant maps. In addition to the amplification of turbulence, such surface deformations lead to local flow separation, instigating low-frequency unsteadiness. One consequence of significance to practical design is the presence of low frequency unsteadiness similar to that encountered in impinging or ramp shock boundary layer interactions.

Abstract— The linear and nonlinear stages of disturbance development in the supersonic boundary layer over a 10° expansion corner is investigated numerically within the framework of Navier—Stokes equations for Mach number 3. The effect of sudden flow expansion on the disturbance evolution is analyzed. The flow stabilization effect observable in the aerodynamic experiment is also discussed.

In this paper a novel design method of aerodynamic configuration is proposed to integrate forebody, strut and inlet for strutjet engine, and a model at design point of Mach number 6 is generated to investigate the aerodynamic performance by both simulations and experiments. The basic flow field employed by proposed method is a combined flow named IBB, which is combined by Internal Conical Flow A (ICFA), truncated Busemann flow I (BI) for external section, and truncated Busemann flow II (BII) for internal section. The model configuration is generated by streamline tracing method from basic flow field, in which the forebody section is traced from ICFA and BI flows, and the inlet as well as strut section is traced from BII flow. The simulations in Mach number 4, 5, and 6 demonstrate uniform starting flow fields with relatively high total pressure recovery, which agree well with experiments in wind tunnel. Additionally, in low Mach number cases, this inlet could start at Mach number 3 while it is unstarted at Mach number 2.7; in high Mach number cases, a uniform flow could still exist in Mach number 6.5 while a relatively strong shock wave boundary layer interaction is found in cowl area of Mach number 7 case, indicating the inlet designed by proposed method works in a relatively wide Mach number range.

It was experimentally observed that the front of a propagating flame of a well-mixed diluted methane-oxygen mixture at 298 K and 100–300 Torr does not form von Karman vortex shedding behind the obstacle of cylindrical shape of 30–50 mm in diameter, including a perforated cylinder; however, the instability under the same conditions occurs in the flow of hot products. In the perforated cylinder, the occurrence of local primary ignition centers on its inner surface is observed. In the mathematical modeling, the main observed features of the flame front propagation were taken into account: the chain branched mechanism of gaseous combustion and the absence of vortex shedding behind the obstacle at flame propagation. It was shown that a qualitative model of compressible dimensionless non-reactive/reactive Navier–Stokes equations in low Mach number approximation yields both the mode of the emergence of von Karman instability in chemically inert gas and the absence of the instability in the mode of flame propagation in a reacting flow. The model computations confirmed the occurrence of local primary ignition centers on the inner surface of the obstacle.

The accurate prediction of aerothermal surface loading is of paramount importance for the design of high speed flight vehicles. In this work, we consider the numerical solution of hypersonic flow over a double-finned geometry, representative of the inlet of an air-breathing flight vehicle, characterized by three-dimensional intersecting shock-wave/turbulent boundary-layer interaction at Mach 8.3. High Reynolds numbers ($Re_L \approx 11.6 \times 10^6$ based on free-stream conditions) and the presence of cold walls ($T_w/T_o \approx 0.27$) leading to large near-wall temperature gradients necessitate the use of wall-modeled large-eddy simulation (WMLES) in order to make calculations computationally tractable. The comparison of the WMLES results with experimental measurements shows good agreement in the time-averaged surface heat flux and wall pressure distributions, and the WMLES predictions show reduced errors with respect to the experimental measurements than prior RANS calculations. The favorable comparisons are obtained using an LES wall model based on equilibrium boundary layer approximations despite the presence of numerous non-equilibrium conditions including three dimensionality, shock-boundary layer interactions, and flow separation. Lastly, it is also demonstrated that the use of semi-local eddy viscosity scaling (in lieu of the commonly used van Driest scaling) in the LES wall model is necessary to accurately predict the surface pressure loading and heat fluxes., Comment: 27 pages, 27 figures

Ballistic range experiments were performed for a hemisphere-flare-cylinder model at supersonic Mach numbers in the transitional Reynolds number range at Nagoya University. The free-flight portion was modelled as axisymmetric in ANSYS Fluent® V.19.0. Projectile deceleration was included in the simulation as a function of the drag force over an approximate flight Mach number range 2.0–1.90. The projectile deceleration magnitude averaged approximately 700g and 1550g (g = 9.81 m s $$^{-2}$$ ) for two experiment cases. The Reynolds number (Re $$_\mathrm{d}$$ ) for each case based on the initial flight Mach number was Re $$_\mathrm{d}$$ = 90,000 and 177,000 respectively. The flow field, separation shock angle, averaged deceleration magnitude and averaged drag coefficient were compared between experiment and simulation. Agreement of these parameters was consistent for the Re $$_\mathrm{d}$$ = 177,000 case. This result contributed towards validation of the numerical acceleration technique. Differences for the Re $$_\mathrm{d}$$ = 90,000 case are explained with reference to experiment and simulation data.

An experimental study on the traveling crossflow instability over a 60 $$^\circ $$ swept flat plate was conducted. The Mach number is 6, the angle of attack of the model is 5 $$^\circ $$ . The traveling crossflow waves and the secondary instability of the traveling crossflow waves were visualized by nano-tracer-based planar laser scattering (NPLS) technique. In the spanwise NPLS images, the traveling crossflow waves appeared as regular strikes, and the secondary instability appeared as small eddies attached to strikes. The wavelet transform was used to study the wavelength contents of the traveling crossflow waves. The most amplified wavelength is stable before the secondary instability happening, which is around 12 mm at $$Re_\infty $$ = 3.45 $$\times $$ 10 $$^6$$ m $$^{-1}$$ . Besides, the Reynolds number effects on the boundary layer transition and traveling crossflow instability were discussed.

The energy transfer and conversion of exhaust gas flowing across junctions of exhaust manifold plays an important role in determining performance of turbocharging system. With the increase of engine boost pressure, exhaust gas velocity increases significantly, which increased compressibility of exhaust gas flow simultaneously. The existed exhaust T-junction models that do not take gas compressibility into consideration are not capable to simulate the exhaust gas flow with high boost pressure good enough. In order to predict pressure loss coefficients of high-pressure gas flow with higher accuracy, this paper developed a new T-junction models in which compressibility of gas flow was taken into consideration. A particle image velocimetry (PIV) test rig was established to provide data for investigating influences of flow parameters on the flow state and thermodynamic parameters in the T-junction. The results show that there are obvious streamline contractions in the internal flow field of the junction, and the formed boundary streamline divides the junction into two regions, as the flow ratio increases, the absolute value of the vorticity increases, and the vorticity in most areas of the branch joint is zero. In addition, the flow ratio of the branch pipe and the main pipe and the Mach number of the outflow affect the flow of the junction. Based on the results of the PIV test, a new pressure loss coefficient T-junction model for compressible flow is proposed. The new model has good prediction accuracy of the pressure loss coefficient and the prediction accuracy of the exhaust pressure wave increases by 4.43% when the pressure coefficient model is used in one-dimensional simulation program.

The distribution and evaporation characteristics of a transverse vaporized kerosene jet in supersonic flow were experimentally investigated in a ground direct-connected experimental system. Smear-free and high spatiotemporal resolution images of the jet plume were obtained using the pulsed laser sheet imaging method. The incoming flow had a total pressure of 1.0 MPa, a total temperature of 980 K, and an inlet Mach number of 2.0. The boundary and length of the jet plume were identified using the maximum interclass variance method (Otsu). The experimental results show that the kerosene evaporation process is closely related to the fuel temperature and local flow velocity. The two factors competitively influence the phase distribution in the jet plume, resulting in different evaporation distances in the direction perpendicular to the wall. Under subcritical conditions, the jet penetration depths remain basically constant, while the surface wave structures are distinct on the windward side. The jet of critical-state kerosene exhibits a quasi-gaseous flow structure in the near field and undergoes a complex phase transition of secondary condensation and secondary evaporation. According to the different phase states and evaporation degrees, the jet plume is divided into three evaporation regions, providing a complete description of the evaporation characteristics of kerosene at different temperatures in supersonic flow.

Experiments were carried out to assess the effect of rear face geometrical modifications of a wall mounted rectangular cavity on pressure fluctuations at supersonic speed. Initially, tests were conducted on a base cavity of length to depth ratio of 3.6, in a 50 × 100 mm at a Mach number of 2, and a Reynolds number of 2 × 106 based on cavity length. Geometrical modifications to the rear face of the base cavity in the form of single ramp, double ramp and partial ramp circle were done to study their acoustic emission characteristics. Oil flow images revealed the complex nature of the flow inside the cavities. Time resolved flow visualization revealed attenuation of receptive interaction at the cavity leading edge, and minimized mass flux into the cavity by adopting these modifications. Unsteady measurements indicated reduction in the peak tones, broadband levels, and coherence. Among the modified geometries, double ramp configuration significantly reduced the overall sound pressure levels and the fluctuating pressures of the order of 10 dB and 50% respectively.

The ramjet/scramjet engines require the control-oriented model to predict the inlet flow field in less than a few seconds. However, it is challenging for these kinds of inlets which utilize curved shock waves to compress the air flow. In this paper, a reduced-order model based on the computational fluid dynamics, the proper orthogonal decomposition theory, and the radial basis function interpolation method is developed. After that, the curved shock waves dominated flow fields of a ramjet inlet under different angles of attack and free stream Mach numbers are predicted with this reduced-order model and compared to the full order computational fluid dynamics solutions. The results show that this reduced-order model can successfully predict the curved shock waves, the curved shock wave/boundary layer interactions, and the shock trains caused by a back pressure with high accuracies. The consumed time is only 0.11 s. The performance parameters are also predicted with the relative errors no more than 2%.

We consider the homogenization limit of the compressible barotropic Navier–Stokes equations in a three-dimensional domain perforated by periodically distributed identical particles. We study the regime of particle sizes and distances such that the volume fraction of particles tends to zero but their resistance density tends to infinity. Assuming that the Mach number is decreasing with a certain rate, the rescaled velocity and pressure of the microscopic system converges to the solution of an effective equation which is given by Darcy’s law. The range of sizes of particles we consider is exactly the same which leads to Darcy’s law in the homogenization limit of incompressible fluids. Unlike previous results for the Darcy regime we estimate the deficit related to the pressure approximation via the Bogovskiĭ operator. This allows for more flexible estimates of the pressure in Lebesgue and Sobolev spaces and allows to proof convergence results for all barotropic exponents [Formula: see text].

Effect of eccentricity on jets delivered from a primary nozzle surrounded by secondary flow is studied experimentally. The eccentricity is achieved by offsetting the annular spacing of the co-flow passage, surrounding the center nozzle generating the primary jet. This is achieved by making a base plate with its center which is offset by 2 mm in the downward direction with the center of inner nozzle. The studies have been carried out in possible flow regimes of jet exiting from a convergent nozzle like subsonic and correctly and underexpanded sonic conditions. The experiments were carried out for primary nozzle exit Mach numbers of 0.3, 0.5, 0.8 and 1.0 at correct expansion. The studies in underexpanded regime were conducted for nozzle pressure ratios (NPR) of 3 and 4. The eccentricity in the co-flow jet leads to 38% reduction in the supersonic core length at NPR 3 and 60% at NPR 4 in comparison with concentric co-flow configuration. The eccentricity also leads to early mixing of co-flowing jet.

In this paper, the effects of streamwise Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators on Shock Wave/Boundary Layer Interaction (SWBLI) are investigated in a Mach 2.5 supersonic flow. In this regard, the numerical investigation of NS-DBD plasma actuator effects on unsteady supersonic flow passing a 14° shock wave generator is performed using simulation of Navier-Stokes equations for 3D-flow, unsteady, compressible, and k ‐ ω SST turbulent model. In order to evaluate plasma discharge capabilities, the effects of plasma discharge length on the flow behavior are studied by investigating the flow friction factor, the region of separation bubble formation, velocity, and temperature distribution fields in the SWBLI region. The numerical results showed that plasma discharge increased the temperature of the discharge region and boundary layer temperature in the vicinity of flow separation and consequently reduced the Mach number in the plasma discharge region. Plasma excitation to the separation bubbles shifted the separation region to the upstream around 6 mm, increased SWBLI height, and increased the angle of the separation shock wave. Besides, the investigations on the variations of pressure recovery coefficient illustrated that plasma discharge to the separation bubbles had no impressive effect and decreased pressure recovery coefficient. The numerical results showed that although the NS-DBD plasma actuator was not effective in reducing the separation area in SWBLI, they were capable of shifting the separation shock position upstream. This feature can be used to modify the structure of the shock wave in supersonic intakes in off-design conditions.

We conduct numerical experiments to determine the density probability distribution function (PDF) produced in supersonic, isothermal, self-gravitating turbulence of the sort that is ubiquitous in star-forming molecular clouds. Our experiments cover a wide range of turbulent Mach number and virial parameter, allowing us for the first time to determine how the PDF responds as these parameters vary, and we introduce a new diagnostic, the dimensionless star formation efficiency versus density ($\epsilon_{\rm ff}(s)$) curve, which provides a sensitive diagnostic of the PDF shape and dynamics. We show that the PDF follows a universal functional form consisting of a log-normal at low density with two distinct power law tails at higher density; the first of these represents the onset of self-gravitation, and the second reflects the onset of rotational support. Once the star formation efficiency reaches a few percent, the PDF becomes statistically steady, with no evidence for secular time-evolution at star formation efficiencies from about five to 20 percent. We show that both the Mach number and the virial parameter influence the characteristic densities at which the log-normal gives way to the first power-law, and the first to the second, and we extend (for the former) and develop (for the latter) simple theoretical models for the relationship between these density thresholds and the global properties of the turbulent medium., Comment: 18 pages, 10 figures, accepted in MNRAS, revised version

Streamwise vortices and the associated streaks evolve in boundary layers over flat or concave surfaces due to disturbances initiated upstream or triggered by the wall surface. Following the transient growth phase, the fully-developed vortex structures become susceptible to inviscid secondary instabilities resulting in early transition to turbulence via `bursting' processes. In high-speed boundary layers, more complications arise due to compressibility and thermal effects, which become more significant for higher Mach numbers. In this paper, we study G\"{o}rtler vortices developing in high-speed boundary layers using the boundary region equations (BRE) formalism, which we solve using an efficient numerical algorithm. Streaks are excited using a small transpiration velocity at the wall. Our BRE-based algorithm is found to be superior to direct numerical simulation (DNS) and ad-hoc nonlinear parabolized stability equation (PSE) models. BRE solutions are less computationally costly than a full DNS and have a more rigorous theoretical foundation than PSE-based models. For example, the full development of a G\"{o}rtler vortex system in high-speed boundary layers can be predicted in a matter of minutes using a single processor via the BRE approach. This substantial reduction in calculation time is one of the major achievements of this work. We show, among other things, that it allows investigation into feedback control in reasonable total computational times. We investigate the development of the G\"{o}rtler vortex system via the BRE solution with feedback control parametrically at various freestream Mach numbers $M_\infty$ and spanwise separations $\lambda$ of the inflow disturbances., Comment: 22 pages

This work focuses on the heat flux characteristics of a hypersonic blunt body with a tangential Mach 3 supersonic cooling film. Experiments were carried out with main flow’s Mach numbers of 6, 7, a...

Experiments were conducted on a cylinder-induced shock-wave/boundary-layer interaction (SBLI) at Mach 6. The state of the interaction was brought through transition using a combination of boundary-...

Results of a numerical study of mixing, ignition, and combustion of a cold hydrogen jet propagating along the lower wall of a channel parallel to a supersonic (M = 2) flow of an inert gas mixture/humid hot air are reported. The computations are performed with the use of the ANSYS CFD Fluent commercial software by means of solving transient Favre-averaged Navier–Stokes equations supplemented with the $$k$$ – $$\omega$$ SST turbulence model and several kinetic schemes of hydrogen combustion. Two single-step schemes and three detailed kinetic schemes including 16, 38, and 37 forward and backward reactions are considered. The goal of the study is to choose a computation method and kinetic mechanism that ensure good agreement with experimental data on supersonic combustion of a coflowing hydrogen jet. In the case of a non-reacting flow, it is demonstrated that the computational algorithm can accurately predict the parameters of mixing of the hydrogen jet and external flow. In the case of a reacting flow, the flow characteristics are significantly affected by large vortex structures developing at the interface between the combustion layer and the external flow. If the flow unsteadiness is taken into account and a detailed kinetic scheme with 37 reactions is used, good agreement of the mean characteristics of the flow with experimental data on the distributions of pressure, temperature, Mach number, and species concentrations at the combustor exit is provided.

Conjugate heat transfer analysis of supersonic turbulent flow over a finite thickness flat plate is studied numerically. The flow field is modeled by employing the Favre averaged Navier–Stokes (FANS) equations. Turbulence is modeled with the Menter's shear-stress-transport (SST) k−ω model. The comparison of NCHT (non-conjugate heat transfer) and CHT (conjugate heat transfer) analysis shows that the interface temperature and fluid temperature variation for the CHT analysis are lower compared to the NCHT analysis. The results indicate that the interface temperature is significantly increased for the low thermal conductivity plate material, due to the reduced rate of heat transfer. The effect of plate thickness on the flow field temperature variation is negligible. However, it is observed that the maximum interface temperature Ti,max increases with the increase in plate thickness. The values of maximum interface temperature difference between the 5 mm and 2 mm thick Steel-1006 plates at Mach numbers 2, 4, and 5 in non-dimensional (non-dimensionalized with free-stream temperature) form are 0.05594, 0.129, and 0.1816 respectively for the isothermal boundary condition. Results indicate that for Mach 2, the heated wall thermal boundary conditions acts like heat source, leading to the higher interface temperature, whereas for Mach 4, and 5 the heated wall thermal boundary conditions acts like heat sink, thus lowering the interface temperature. The results show that for NCHT analysis the leading edge shock strength, and mean pressure along the wall increases with an increase in Mach number. The non-conjugate results (flow and heat transfer characteristics) of the present numerical study is validated with the direct numerical simulation (DNS) results available in the literature.

The objective of the present paper is to reveal the intrinsic mechanisms by which two-dimensional surface indentations influence the laminar–turbulent transition of natural route in hypersonic boun...

An initial (linear) stage in the development of rotating transverse detonation waves in a flat-radial annular combustion chamber is determined and simulated. The problem of linear modal stability of the cylindrical front of Chapman–Jouguet deflagration combustion in a radially diverging subsonic flow with a small Mach number in the presence of perturbation waves of the flow entropy is solved. The steady flame front is described by discontinuity of the gas-dynamic parameters provided that the combustion products are in chemical equilibrium. It is revealed that the flame front is unstable for some types of small perturbations of the main flow of the combustible mixture and the flame front. Instability is determined under the condition of a constant flow rate in the mixture injection system. The spatial forms of oscillations and perturbation waves of the combustion front in the annular combustion chamber are obtained by numerical and analytical methods.

Computational fluid dynamics was used to study the flow through a scaled, mixed-compression, high-speed inlet with a rotating cowl at Mach 4.0 conditions. First, steady Reynolds-averaged Navier–Sto...

This contribution describes experimental and numerical research of an unsteady behaviour of a flow in an end-wall region of a linear nozzle cascade. Effects of compressibility ($$M_\mathrm {2,is}$$ M 2 , is ) and inlet flow angle ($$\alpha _1$$ α 1 ) were investigated. Reynolds number ($$Re_\mathrm {2,is}$$ R e 2 , is $$=8.5\times 10^5$$ = 8.5 × 10 5 ) was held constant for all tested cases. Unsteady pressure measurement was performed at the blade mid-span in the identical position $${\mathfrak {s}}$$ s to obtain reference data. Surface flow visualizations were performed as well as the steady pressure measurement to support conclusions obtained from the unsteady measurements. Comparison of the surface Mach number distributions obtained from the experiments and from the numerical simulations are presented. Flow visualizations are then compared with calculated limiting streamlines on the blade suction surface. It was shown, that the flow structures in the end-wall region were not affected by the primary flow at the blade mid-span, even when the shock wave formed. This conclusion was made from the experimental, numerical, steady as well as unsteady points of view. Three significant frequencies in the power spectra suggested that there was a periodical interaction between the vortex structures in the end-wall region. Based on the data analyses, anisotropic turbulence was observed in the cascade.

The paper describes the impact of aerodynamic coefficients on the ballistic target (BT) velocity and proposes a method of approximation of the dependence of ballistic target drag coefficient Cxa on the Mach number and angle of attack. The paper proves that the proposed approach allows to substantially reduce errors in drag coefficient simulation, but requires a more complicated calculation process.