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Direct numerical simulation of interaction of a laminar boundary layer with a shock wave at a Mach number of 1.45 is performed. The influence of a periodic thermal source on the nature of this interaction is described.

The steady laminar boundary layer flow and heat transfer from a warm, laminar liquid flow to a melting surface moving parallel to a constant free stream is studied in this paper. The continuity, momentum and energy equations, which are coupled nonlinear partial differential equations are reduced to a set of two nonlinear ordinary differential equations, before being solved numerically using the Runge–Kutta–Fehlberg method. Results for the skin friction coefficient, local Nusselt number, velocity profiles as well as temperature profiles are presented for different values of the governing parameters. Effects of the melting parameter, moving parameter and Prandtl number on the flow and heat transfer characteristics are thoroughly examined. It is found that the problem admits dual solutions.

At hypersonic speeds, the external wall temperatures of an aerospace vehicle vary significantly. As a result, there is a considerable heat transfer variation between the boundary layer and the wall of the hypersonic vehicle. In this article, numerical computations are performed to investigate the effect of wall temperature on the separation bubble length in laminar hypersonic shock-wave/boundary-layer interaction flows over double-cone configuration at the Mach number of 12.2. The flow field is described in detail in terms of different shocks, expansion fans, shear layer and separation bubble. The variation of the Prandtl number has a negligible effect on the flow field and wall data. A specific heat ratio of less than 1.4 results in the better prediction of wall pressure and heat flux in the shock/boundary-layer interaction region. It is observed that as the wall temperature is increased, the separation bubble size and hence the separation shock length increases. The high firmness of the laminar boundary-layer at a high Mach number shows that the wall temperature in the shock/boundary-layer interaction region has little effect. The peak wall pressure and heat flux decrease with an increase in wall temperature. An estimation is developed between separation bubble length and wall temperature based on the computed results.

The work is devoted to modeling the disturbance propagation in viscous incompressible laminar boundary layers, using linearized equations for disturbance amplitudes. Along with the numerical model based on original linearized equations, the article considers three models based on equations derived from the original ones by neglecting the streamwise pressure gradient, or the streamwise viscous terms, or both. The models are compared numerically by the example of generation and propagation of disturbances in the boundary layer over a slightly concave plate. Conclusions are drawn about the feasibility of the same simplified models to adequately simulate both Tollmien–Schlichting waves and Görtler vortices in a range of practically important parameters.

Motivated by recent experimental observations, we consider a steady-state boundary layer flow with polymers in forced convection above a heated plate and study how the heat transport might be affected by the polymers. We discuss how a set of equations can be derived for the problem and how these equations can be solved numerically by an iterative scheme. By carrying out such a scheme, we find that the effect of the polymers is equivalent to producing a space-dependent effective viscosity that first increases from the zero-shear value at the plate then decreases rapidly back to the zero-shear value far from the plate. We further show that such an effective viscosity leads to a decrease in the streamwise velocity near the plate, which in turn leads to a reduction in heat transport.

Similarity solution of the laminar boundary layer equations corresponding to an unsteady stretching surface have been studied. The governing time-dependent boundary layer are transformed to ordinary differential equations containg Prandtl number and unsteadiness parameter. The effect of various govern-ing parameters such as Prandtl number and unsteadiness param-eter which determine the velocity and temperature profiles and heat transfer coefficient are studied.

A wall model for large-eddy simulation (LES) is proposed by devising the flow as a combination of building blocks. The core assumption of the model is that a finite set of simple canonical flows contains the essential physics to predict the wall shear stress in more complex scenarios. The model is constructed to predict zero/favourable/adverse mean pressure gradient wall turbulence, separation, statistically unsteady turbulence with mean flow three-dimensionality, and laminar flow. The approach is implemented using two types of artificial neural networks: a classifier, which identifies the contribution of each building block in the flow, and a predictor, which estimates the wall shear stress via a combination of the building-block flows. The training data are obtained directly from wall-modelled LES (WMLES) optimised to reproduce the correct mean quantities. This approach guarantees the consistency of the training data with the numerical discretisation and the gridding strategy of the flow solver. The output of the model is accompanied by a confidence score in the prediction that aids the detection of regions where the model underperforms. The model is validated in canonical flows (e.g. laminar/turbulent boundary layers, turbulent channels, turbulent Poiseuille–Couette flow, turbulent pipe) and two realistic aircraft configurations: the NASA Common Research Model High-lift and NASA Juncture Flow experiment. It is shown that the building-block-flow wall model outperforms (or matches) the predictions by an equilibrium wall model. It is also concluded that further improvements in WMLES should incorporate advances in subgrid-scale modelling to minimise error propagation to the wall model.

At moderate Reynolds numbers and angles of attack, the Laminar Boundary Layer (LBL) becomes unstable on the surface of airfoils, and causes periodic vortex shedding, which means undesired tonal peaks in the spectrum of the emitted aeroacoustic noise along with increased vibration and decreased aerodynamic performance. In the past, numerous research campaigns focused on the LBL vortex shedding, including measurements and numerical simulations as well. The results of these investigation showed that, the formation of the LBL instability related to the presence of the laminar separation bubble. It was also shown that, the spectrum of the emitted noise has a multitonal behavior, and the scaling of the mean frequency with the free stream velocity has a ladder structure. Based on these results, the LBL instability is a complex phenomenon; however, in the preliminary design of axial flow turbomachines the prediction of the frequency of the vortex shedding is essential, therefore the use of semi-empirical formulas is usual to achieve this goal. The previous researches mostly focused on separated airfoils, however, in case of turbomachines, the blades form a cascade, which can significantly affect the aerodynamic of the airfoils, i.e. it can affect the behavior of the LBL instability as well. According to this, in the present paper the LBL instability of NACA 0012 cascades are investigated, using 2D computational fluid mechanics and aeroacoustics simulations. The investigation involves the variation of the angle of attack, the chord based Reynolds-number and the spacing. The results are compared to the semi-empirical Brooks-Pope-Marcolini model.

Abstract—The results of numerical simulation of the interaction between a shock and the laminar boundary layer on a flat plate in motion in supersonic perfect-gas flow at Mach number М∞ = 3 are considered. The shock is preassigned using the Rankine—Hugoniot boundary conditions, which corresponds to a shock wave produced by a wedge with a given semi-vertex angle in an inviscid gas flow. The simulation is based on the numerical solution of the time-dependent, two-dimensional Navier—Stokes equations by time marching to steady state. The numerical results are verified by means of comparing the results for separation flow past a flat plate at rest with experimental data. The numerical data are used to investigate the effect of the plate velocity of the separation flow structure and the basic laws governing the problem. It is shown that the motion of the plate downstream diminishes the separation zone length, whereas the opposite motion leads to its increase.

We employ global input–output analysis to quantify amplification of exogenous disturbances in compressible boundary layer flows. Using the spatial structure of the dominant response to time-periodic inputs, we explain the origin of steady reattachment streaks in a hypersonic flow over a compression ramp. Our analysis of the laminar shock–boundary layer interaction reveals that the streaks arise from a preferential amplification of upstream counter-rotating vortical perturbations with a specific spanwise wavelength. These streaks are associated with heat-flux striations at the wall near flow reattachment and they can trigger transition to turbulence. The streak wavelength predicted by our analysis compares favourably with observations from two different hypersonic compression ramp experiments. Furthermore, our analysis of inviscid transport equations demonstrates that base-flow deceleration contributes to the amplification of streamwise velocity and that the baroclinic effects are responsible for the production of streamwise vorticity. Finally, the appearance of the temperature streaks near reattachment is triggered by the growth of streamwise velocity and streamwise vorticity perturbations as well as by the amplification of upstream temperature perturbations by the reattachment shock.