The influence of surface bumps on boundary-layer transition was experimentally investigated in the present work. The experiments were conducted in a (quasi-) two-dimensional flow at low to high subsonic Mach numbers and chord Reynolds numbers up to 10 million in the low-turbulence Cryogenic Ludwieg-Tube Göttingen. Various streamwise pressure gradients relevant for natural laminar flow surfaces were examined. Quasi-two-dimensional bumps, with a sinusoidal shape in the streamwise direction, fixed length and three different heights, were installed on a two-dimensional flat-plate model. The model was equipped with temperature-sensitive paint for non-intrusive transition detection and with pressure taps for the measurement of the surface pressure distributions. Boundary-layer transition was shown to occur at a more upstream location with increasing bump height-to-length ratio. This was mainly due to the local adverse pressure gradient on the downstream side of the bump, which was particularly pronounced in the case of the bump with the largest height-to-length ratio, thereby inducing boundary-layer separation (as verified via oil-film visualizations). In the case of the bump with the smallest height-to-length ratio, bumpinduced transition was found to be dependent on global pressure gradient, Mach number and Reynolds number; however, the influence of these parameters on transition induced by bumps with larger height-tolength ratios was significantly reduced. The sensitivity of boundary-layer transition to the effect of the bumps was shown to be more pronounced with stronger global flow acceleration and at smaller Mach numbers.

Knowledge on the boundary-layer transition location at large chordReynolds numbers is essential to evaluate the performance of airfoils designed for modern wind-turbine rotor blades. In the present work, a temperature-sensitive paint was used to systematically study boundary-layer transition on the suction side of a DU 91-W2-250 airfoil. The experiments were performed in the High-Pressure Wind Tunnel Göttingen at chord Reynolds numbers up to 12 million and angles-of-attack from -14° to 20°. The coefficients of airfoil lift, drag, and pitching moment were also obtained after integration of the pressure distributions measured on the surface and in the wake of the wind-tunnel model. The global information obtained via temperature-sensitive paint not only enabled the analysis of the change in the transition location with varying angle-of-attack and chord Reynolds number, but also provided an explanation for the evolution of the aerodynamic coefficients measured at stall and post-stall conditions. The stability of the laminar boundary layers investigated in the experiments was analyzed according to linear stability theory. The results of the stability

The aerodynamic performance of airfoils and blades designed for modern wind-turbine rotors, which have diameters of the order of hundred meters, must be examined at chord Reynolds numbers matching those of practical applications. In general, such high Reynolds numbers cannot be achieved in conventional wind tunnels. Moreover, knowledge on the boundary-layer transition location is essential to evaluate airfoil and blade performance at these flow conditions. This work presents an experimental methodology that can be applied at flow conditions reproducing those of real wind-turbine rotor blades and simultaneously provides aerodynamic coefficients and transition locations. The experimental methodology consists of: the Temperature-Sensitive Paint (TSP) technique for global, non-intrusive and reliable transition detection; conventional pressure measurements for the determination of the aerodynamic coefficients; and the High Pressure Wind Tunnel Göttingen (DNW-HDG) to run the experiments at Reynolds numbers matching those of real applications. The obtained results can be used to verify airfoil and blade performance and to validate numerical predictions. In the present work, the experimental methodology was applied to systematically investigate the aerodynamic performance of an airfoil designed for the mid-span sections of modern wind-turbine rotor blades. The examined chord Reynolds numbers were as high as 12 million and the angle-of-attack ranged from -14° to +20°. The presented methodology was here demonstrated to be mature for productive testing.

The influence of suction on step-induced boundary-layer transition has been experimentally investigated in the Cryogenic Ludwieg-Tube Goettingen at large chord Reynolds numbers (up to 16 • 10 6), Mach numbers from 0.35 to 0.77 and various streamwise pressure gradients by means of temperature-sensitive paint. Surface imperfections, implemented as combination of gap and forward-facing step, caused transition to occur at a location more upstream than in the case of a smooth surface (i.e. without gap and step). For this combination of imperfections, it was demonstrated for the first time in experiments that suction, achieved passively by exploiting the pressure difference between upper and lower side of the model, induced a movement of transition to a more downstream location than without suction, and in most cases even more downstream than on the smooth configuration at the same test conditions. Thus, the effect of suction was to even overcompensate the adverse effect of the combination of gap and forward-facing step on boundary-layer transition for the investigated test conditions.

Knowledge on the boundary-layer transition location at large chordReynolds numbers is essential to evaluate the performance of airfoils designed for modern wind-turbine rotor blades. In the present work, a temperature-sensitive paint was used to systematically study boundary-layer transition on the suction side of a DU 91-W2-250 airfoil. The experiments were performed in the High-Pressure Wind Tunnel Göttingen at chord Reynolds numbers up to 12 million and angles-of-attack from -14° to 20°. The coefficients of airfoil lift, drag, and pitching moment were also obtained after integration of the pressure distributions measured on the surface and in the wake of the wind-tunnel model. The global information obtained via temperature-sensitive paint not only enabled the analysis of the change in the transition location with varying angle-of-attack and chord Reynolds number, but also provided an explanation for the evolution of the aerodynamic coefficients measured at stall and post-stall conditions. The stability of the laminar boundary layers investigated in the experiments was analyzed according to linear stability theory. The results of the stability