Introduction

Give a brief overview of the test case. Describe the main characteristics of the flow. In particular, what are the underlying flow physics which must be captured by the computations ? Give reasons for this choice (e.g. poorly understood flow physics, difficulty to predict the flow with standard turbulence models, ...). Detail any case-specific data that needs to be generated.

This test case features a NACA0020 profile mounted on a flat plate, which is representative of the wing-body junction flow problems encountered in applications of aeronautical interest. The flow features the interaction between the incipient turbulent boundary layer and the mounted airfoil and the main physical phenomenon of interest is the horseshoe vortex developping at the junction and the corner separation. This flow is also highly 3D and anisotropic regarding the turbulent stresses. Establishing a DNS database of this flow is of crucial interest since it has been shown that RANS models (both Boussinesq and Reynolds stresses-based models) display strong difficulties in recovering data from the available experiments. Such a database allows for a more thorough availability of the flow field with respect to the experiments and gives the possibility of using Machine Learning or data-assimilation techniques to improve standard RANS models.

Review of previous studies

Provide a brief review of related past studies, either experimental or computational. Identify the configuration chosen for the present study and position it with respect to previous studies. If the test case is geared on a certain experiment, explain what simplifications ( e.g. concern- ing geometry, boundary conditions) have been introduced with respect to the experiment in the computational setup to make the computations feasible and avoid uncertainty or ambiguity.

A thorough listing of existing experimental and numerical studies regarding wing-body junction flows can be found in Gand et al. (2010). The present DNS is based on the configuration considered in the simulations by Apsley & Leschziner (2001), who were based themselves on the experimental studies by Devenport and Simpson (1990) and Fleming et al. (1995). The Reynolds number based on the airfoil thickness is 115,000 and the flow is almost incompressible with a Mach number based on the freestream velocity of 0.078. The DNS setup is aimed at reproducing the experimental conditions but with half the experimental impacting boundary layer thickness.

Description of the test case

A detailed self-contained description should be provided. It can be kept fairly short if a link can be made to an external data base where details are given. Then only the differences should be clearly indicated.

Geometry and flow parameters

Describe the general set up of the test case and provide a sketch of the geometry, clearly identifying location and type of boundaries. Specify the non-dimensional flow parameters which define the flow regime (e.g. Reynolds number, Rayleigh number, angle of incidence etc), including the scales on which they are based. Provide a detailed geometrical description, by preference in form of a CAD, or alternatively as lists of points and a description of the interpolation.

The following sketch displays a view of the computational domain and flow configuration geometry:

The reference length scale is the wing chord T, with a corresponding Reynolds number $Re_{T}=115,000$ . The computational domain size in the streamwise direction is $75T,8T$ in the spanwise direction and $3T$ in the wall-normal direction. The coordinates origin is located at the root leading edge of the airfoil. The is no flow incidence relatively to the wing, corresponding to an angle of attack of 0 degrees.

Boundary conditions

Specify the prescribed boundary conditions, as well as the means to verify the initial flow development. In particular describe the procedure for determining the in flow conditions comprising the instantaneous (mean and fluctuating) velocity components and other quantities. Provide reference profiles for the mean flow and fluctuations at in flow - these quantities must be supplied separately as part of the statistical data as they are essential as input for turbulence-model calculations. For checking purposes, these profiles should ideally also be given downstream where transients have disappeared; the location and nature of these cuts should be specified, as well as the reference result.

For this particular flow configuration, we aim at simulating a transitional boundary layer in order to replicate the flow conditions of the experiment. To do so, a Blasius velocity profile is imposed at the inlet corresponding to $Re_{x}=900,000$ with a uniform temperature. A laminar boundary layer is then established, and a flow perturbation is introduced at $Re_{x}=950,000$ with amplitude $A_{T}={\frac {\rho _{ref}U_{ref}^{2}}{T}}$ to trigger transition to turbulence. The total length of the boundary layer was determined such that the turbulent boundary layer thickness upwind the airfoil reaches half of the experimental value. Symmetry conditions are imposed at the lateral boundary conditions ( $z=\pm 4T$ ). The bottom boundary is a no-slip adiabatic wall type for $xz$ planes at $y=0$ between $x=-12.75T$ and $x=17T$ , and symmetry type between $x=17T$ and $x=67T$ . The outlet is located away from the profile at $x=67T$ , and the zone between $x=17T$ and $x=67T$ acts as a sponge layer, featuring increasingly coarse elements such that a constant flow field is recovered when reaching the outlet.

Contributed by: Alessandro Colombo (UNIBG), Francesco Carlo Massa (UNIBG), Michael Leschziner (ICL/ERCOFTAC), Jean-Baptiste Chapelier (ONERA) — University of Bergamo (UNIBG), ICL (Imperial College London), ONERA

Front Page

Description

Computational Details

Quantification of Resolution

Statistical Data

Instantaneous Data

Storage Format