Grey area in embedded wall-modelled LES on a transonic nacelle-aircraft configuration

Abstract

A scale-resolving hybrid RANS–LES technique is applied to an aircraft-nacelle configuration under transonic flow conditions using the unstructured, compressible TAU solver. In this regard, a wall-modelled LES methodology is locally applied to the nacelle lower surface to examine shock-induced separation. To circumvent the grey-area issue of delayed turbulence onset, a Synthetic Turbulence Generator (STG) is used at the RANS–LES interface. Prior to the actual examinations, fundamental features of the simulation technique are validated by simulations of decaying isotropic turbulence as well as flat plate flows. For the aircraft-nacelle configuration at a Reynolds number of 3.3 million, a sophisticated mesh with 420 million points was designed which refines 32 % of the outer casing surface of the nacelle. The results show a development of a well-resolved turbulent boundary layer with a broad spectrum of turbulent scales which demonstrates the applicability of the mesh and method for aircraft configurations. Furthermore, the necessity of a low-dissipation low-dispersion scheme is demonstrated. However, a noticeable drop of the surface skin friction downstream of the STG motivates further research on the impact of the interface modelling on the shock–boundary layer interaction.