Various aspects of the numerical modeling of a scale-model turbofan blade are investigated for the purpose of future noise assessment. The fan geometry considered is the baseline configuration of the "Fan Noise Source Diagnostic Test" (SDT) experimental setup. Both Reynolds Averaged Navier-Stokes (RANS) simulations and Large Eddy Simulations (LES) are performed for a single rotor blade without the stator vanes and compared against the full configuration from previous works. With RANS simulations the mesh convergence is systematically studied. In the LES, two wall models and two numerical schemes are considered to evaluate the sensitivity of the boundary-layer transition to turbulence on the blade suction side. The different LES results are compared with the RANS solutions obtained using fully turbulent and transitional turbulence models. All RANS results show similar performance, whereas in the LES the no-slip wall condition gives better performance than the log-law slip wall condition. The prism layers on the hub and the shroud change the boundary layer profile upstream of the blade but do not affect the performance. On the blade, the RANS simulations show a laminar recirculation bubble whereas the LES exhibit a leading-edge vortex spiralling radially. The transitional RANS turbulence model gives the best agreement with the LES on the pressure side. The LES results are more affected by the mesh resolution than by the wall model or the numerical scheme, especially in the tip. In the wake all results show a good agreement with the experiment. Nomenclature Variables y + Dimensionless wall distance τ w Wall shear stress C p = P−P ∞ 0.5ρ ∞ U 2 ∞ Mean pressure coefficient C f = τ w 0.5ρU ∞ Mean skin friction coefficient U Mean velocity P Mean pressure Indices w Wall quantity ∞ Free-flow quantity t Total (stagnation) quantity This work is licensed under a Creative Commons Attribution-NonCommercial-NoDerivatives 4.0 International License CC-BY-NC-ND 4.0

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A classic approach to computational fluid dynamics is to perform simulations with a fixed set of variables in order to account for parameters and boundary conditions. However, experiments and real-life performance are subject to variability in their conditions. In recent years, the interest of performing simulations under uncertainty is increasing, but this is not yet a common rule, and simulations with lack of information are still taking place. This procedure could be missing details such as whether sources of uncertainty affect dramatic parts in the simulation of the flow. One of the reasons of avoiding to quantify uncertainties is that they usually require to run an unaffordable number of CFD simulations to develop the study. To face this problem, Non-Intrusive Uncertainty Quantification (UQ) has been applied to 3D Reynolds-Averaged Navier-Stokes simulations of an under-expanded jet from an aircraft exhaust with the Spalart-Allmaras turbulent model, in order to assess the impact of inaccuracies and quality in the simulation. To save a large number of computations, sparse grids are used to compute the integrals and built surrogates for UQ. Results show that some regions of the jet plume can be more sensitive than others to variance in both physical and turbulence model parameters. The Spalart-Allmaras turbulent model is demonstrated to have an accurate performance with respect to other turbulent models in RANS, LES and experimental data, and the contribution of a large variance in its parameter is analysed. This investigation explicitly outlines, exhibits and proves the details of the relationship between diverse sources of input uncertainty, the sensitivity of different quantities of interest to said uncertainties and the spatial distribution arising due to their propagation in the simulation of the high-speed jet flow. This analysis represents first numerical study that provides evidence for this heuristic observation.

A wall-modeled Large Eddy Simulation (LES) of the turbulent flow in the NASA Source Diagnostic Test turbofan is successfully performed for the first time. A good agreement with aerodynamic measurements is observed for both Reynolds Averaged Navier-Stokes and LES results, although the LES provides be er results in the tip regions where large coherent structures appear and no flow separation on the stator vanes is observed. In the LES the boundary layer naturally transition to turbulence on the blade suction side but remains quasi laminar over most of its pressure side. e rotor-wake turbulence yielding the stage broadband noise is then seen to be quasi isotropic. Transition on the downstream stator vanes is not triggered by the wake impingement but rather occurs at mid-chord. Finally, acoustics are investigated using both Ffowcs Williams & Hawkings' and Goldstein's analogies from the recorded LES noise source on the stator vanes. e la er analogy provides levels much closer to the measurements especially at high frequencies, although the results are most likely still influenced by too coherent rotor tip secondary flow at low frequencies.