This paper describes an experimental analysis of the bu¨et phenomenon on a two-dimensional (2D), transonic, and laminar airfoil at a Reynolds number around 3 · 10 6 . Investigations are carried out in ONERA£s S3Ch transonic wind tunnel. The experimental setup allows to vary the Mach number, the angle of attack, and the state of the boundary layer upstream of the shock which can be turbulent or laminar depending on the presence of arti¦cial tripping. Bu¨et occurs when either the angle of attack or the Mach number is set above a given threshold, which depends upon the particular airfoil, and, as shown here, on the state of the boundary layer. Above the threshold, the boundary layer / shock interaction destabilizes, causing the oscillation of the entire §ow ¦eld. In the turbulent case, the shock wave moves back and forth over a signi¦-cant portion of the chord at a frequency of about 75 Hz corresponding to a chord based on Strouhal number S t ≃ 0.07, in agreement with previous researches on this phenomenon. In the laminar case, a similar unsteady situation occurs but at a frequency much higher, about 1130 Hz, which corresponds to a Strouhal number of about S t ≃ 1. Flow oscillations are limited to the shock foot, the shock itself moving only lightly. The turbulent and laminar bu¨et thresholds are provided. An attempt to apply the classical feedback loop scenario to explain the unsteadiness of the §ow in the laminar case is carried out but shows a deceptive agreement with the experimental data. Two other mechanisms of unsteadiness are additionally explored, one based on vortex shedding behind the airfoil and the other on the possible breathing of the laminar separation bubble, which give valuable insights into the §ow physics.
Unsteady Pressure Sensitive Paint is applied to measure and investigate the shock dynamics on a transonic laminar airfoil which exhibits a rapid shock oscillation at a normalized frequency, based on chord and freestream velocity, close to one. The main motivation is to assess the capability of the PSP technique to capture the pressure fluctuations imposed at the upper surface of the wing by the oscillating shock wave, in the light of the recent developments performed at ONERA of the unsteady Anodized-Aluminum PSP method. Beyond frequency resolution, it is found that three main difficulties arise when applying PSP for the present unsteady laminar flow. The first concerns the roughness of the PSP layer, which must remain low to prevent unwanted transition of the laminar boundary layer. The second deals with the correction of illumination as compressibility and wing deformations may affect it. The third is about the correction of the temperature sensitivity of the PSP, which is mandatory in the present high speed flow situation. These three questions are carefully addressed. In addition, measurements for the case when the boundary layer at the upper surface of the airfoil is made fully turbulent by forced tripping are carried out for reference. In this case the dynamics is much slower, with a normalized frequency close to 0.05. The PSP pressure fields offer a rich database to analyze the flow features. In particular the spatial sampling provided by the PSP helps to qualify the twodimensional nature of the shock movements and to question the wave propagation process implied by these unsteadiness.
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