The effect of discrete roughness elements on the development and breakdown of stationary crossflow instability on a swept wing is explored. Receptivity to various element heights and chordwise locations is explored using a combination of experimental and theoretical tools. Forcing configurations, determined based on linear stability predictions, are manufactured and applied on the wing in a low turbulence facility. Measurements are performed using infrared thermography, quantifying the transition front location, and planar particle image velocimetry, providing a reconstruction of stationary crossflow instabilities and their associated growth. Measurements are corroborated with simulations based on nonlinear parabolised stability equations. Results confirm the efficacy of discrete roughness elements in introducing and conditioning stationary crossflow instabilities. Primary instability amplitudes and resulting laminar-turbulent transition location are found to strongly depend on both roughness amplitude and chordwise location. The Reynolds number based on element height is found to satisfactorily approximate the initial forcing amplitude, revealing the importance of local velocity effects in non-zero-pressure gradient flows. Direct estimation of initial perturbation amplitudes from nonlinear simulations suggests the existence of pertinent flow mechanisms in the element vicinity, active in conditioning the onset of modal instabilities. Dedicated velocimetry planes, elucidate the development of a momentum deficit wake which rapidly decays downstream of the element followed by mild growth, representing the first experimental evidence of transient behaviour in swept wing boundary layers. The outcome of this work identifies a strong scalability of the transition dynamics to roughness amplitude and location, warranting the upscaling of roughness elements to more accessible, measurable and spatially resolved configurations in future experiments.

The present work is dedicated to the investigation of the effect of an isolated roughness element on a swept wing boundary layer. In particular, the flow modifications incurred by a single cylindrical element applied on a swept wing model are measured, towards describing the nature of the perturbations introduced in the flow field, their development in the near and far wake region, as well as their eventual breakdown. The measurements are performed using infrared thermography, to achieve a general overview of the element wake origin and spatial spreading. Local quantitative characterization of the stationary and unsteady disturbances evolving in the flow are instead acquired through hot wire anemometry. When present in an undisturbed laminar boundary layer, isolated roughness elements are found to introduce flow disturbances which lead to the formation of a turbulent wedge. As it develops downstream, the wedge undergoes rapid spanwise expansion, affecting the adjacent laminar flow regions. The wedge origin and development is mostly associated with the instabilities introduced by the shedding process initiated in the roughness element wake, comparably to the dominant flow features characterizing the transition of two-dimensional boundary layers conditioned by an isolated roughness element. Nonetheless, the presence of the crossflow velocity component in the boundary layer baseflow notably affects the overall flow development, introducing an asymmetric evolution of the main flow features.

This work investigates the three-dimensional effects of a dielectric barrier discharge actuator as a stall recovery device. The actuator is installed on a NACA0015 airfoil with a 930 mm span and a 300 mm chord and the exposed electrode has periodic triangular tips specifically designed for the case under study. It was tested at progressively increasing velocities, up to 35 m/s corresponding to a Reynolds number of 7×105. The PIV technique has been used to characterize the flow morphology along the model span, identifying the flow region mainly sensitive to the plasma effect and suggesting possible design improvements to achieve a better efficiency.

A dielectric barrier discharge actuator (DBD) is considered and studied as a stall recovery device. The DBD is installed on the nose of a NACA0015 airfoil with chord × span 300 × 930 mm. The geometry of the exposed electrode has periodic triangular tips purposely designed for the case under study. Wind tunnel tests have been carried out over a range of airspeeds up to 35 m/s with a Reynolds number of 700 k. The flow morphology has been characterized by means of the particle image velocimetry technique, obtaining velocity fields and pressure coefficients. By exploring different planes along the model span, the three-dimensional effect of the DBD has been reconstructed, identifying the flow region mainly sensitive to the plasma actuation. Finally, the actuator effectiveness has been quantified accounting for the power consumption data, leading to defining further design improvements in view of a better efficiency.

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