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The friction component is responsible for more than 40% of typical civil aircraft drag. As a consequence, the issue of laminar flow has been of prime importance in aeronautics for many years now. This article is focused on TollmienSchlichting-induced transition and drag predictions of two-dimensional laminar airfoils obtained with experimental and numerical methods. In 2012, a test campaign in the ONERA-S2MA wind tunnel, including infrared acquisitions, pressure sensors, and wake analyses, allowed substantial data to be obtained on such airfoils in transonic conditions. To complete this study, two-dimensional fluid dynamics computations have been performed, either with a Reynoldsaveraged Navier-Stokes solver using transition criteria or with a boundary-layer code combined with direct stability analysis. Furthermore, experimental (wake survey) and numerical (far-field theory) techniques allowing airfoil drag breakdown have been employed. Wind-tunnel and computational fluid dynamics transition predictions have been compared. Good agreement has been observed but the transition criteria may show some limitations in particular situations, such as long separation bubble development. The gains in lift and drag due to laminar flow have been quantified (natural vs triggered transition). Concerning drag reduction, the importance of the viscous pressure component has been highlighted. Finally, the effects of parameters such as angle of attack, Mach number, and Reynolds number on transition location and drag have been investigated.subscript for local state value f = frequency of boundary-layer instability M = freestream Mach number N T = value of factor N at transition onset P i = stagnation pressure Re = freestream Reynolds number based on the airfoil chord T i = stagnation temperature Tu = turbulence level U, V, W = x, y, z velocity components x, y, z = x, y, z coordinates Y = normalized first cell height Λ 2 = mean Pohlhausen parameter ρ = density Subscript 0 = freestream state value

A multi-fidelity optimization technique is applied to the design of a helicopter rotor blade to improve its performance in forward flight. This optimization technique is based on surrogate model that replace the high-fidelity CFD/CSD simulations necessary to capture the three-dimensional unsteady effects generated in the flow field of a complex blade geometry. The single low-fidelity model based on Kriging methodology and generated by lifting-line simulations, leads to a power benefit of 2.5%, which is not reproducible by an a posteriori high-fidelity CSD/CFD computation. The optimization procedure using Co-Kriging surrogate models based on two levels of fidelity (lifting line and CSD/CFD simulations) leads to a realistic blade planform, for which the power benefit is estimated at 2.2%. This optimized solution, obtained after a factor 6 reduction in CPU time, shows the advantages of using a Co-Kriging surrogate model (rather than a single-fidelity Kriging model) for aerodynamic optimizations.