A natural laminar flow (NLF) airfoil is designed to reduce drag by expanding laminar flow areas. In-depth knowledge of transition performance is essential for its aerodynamic design. The k-ω-γ-Re_θ framework, which consists of the SST k-ω turbulence model and γ-Re_θ transition model, is employed to simulate transitional flows around an NLF wing RAE5243 airfoil. The transition performances of the RAE5243 airfoil under various values of turbulent intensity, temperature, angle of attack, and Mach number are simulated and compared. The results show that the rise of inflow turbulent intensity will promote an earlier transition on both the suction and pressure sides. The influence of wall temperature on transition is limited. The rise of angle of attack will lead to an earlier transition on the pressure side but a later transition on the suction side. With the rise of Mach number, the transition happens earlier under a zero and positive angle of attack but later under a negative angle of attack. In addition, the correlation of transition onset locations with respect to turbulent intensity, surface temperature, angle of attack, and Mach number is established based on numerical results.

The transition characteristics of a modified RAE5243 airfoil were investigated using a wind tunnel test and numerical simulations. Transition detection is of great significance for the assessment of drag reduction. In wind tunnel tests, transition location can be detected by infrared thermography. However, in subsonic and transonic wind tunnel tests, the temperature difference between the laminar flow region and turbulent flow region is small. Moreover, the test models are usually made of metals, which make the transition location hard to identify. Combined with infrared thermography, a carbon nanotube heating coating powered by electricity was used to detect the transition location of a modified RAE5243 airfoil wing. The effects of heating power, angle of attack (AOA), and Mach number were studied. The results show that heating power has no impact on transition location. As the AOA increases, the transition location moves forward. With an increase in Mach number, the transition location moves forward first and then backward, and it reaches its most forward point at Ma = 0.75. The results of our numerical simulations indicate that, at Ma$\ge$ 0.75, a shock wave appears on the wing, and the transition is closely related to the shock wave rather than the adverse pressure gradient.

Advanced measurement technology on boundary layer transition is an effective means to study the flow mechanism and the performance of the cascade. In this research, to investigate the boundary layer transition of transonic VKI-RG turbine cascade, a noncontact measurement technique named temperature-sensitive paint (TSP) that is capable of quantitatively measuring the surface temperature of a model is used. Under the conditions of outlet Mach number (Ma₂) = 0.4, 1.03, and 1.2, the transitional location of a cascade is accurately measured by TSP technique. Moreover, a numerical study on the transitional location of a planar cascade was conducted with the conditions of Mach number (Ma) and attack angle (α). The numerical results show that the transitional location moved forward with the increase in Ma₂ and α at Ma₂ < 1; otherwise, the transitional location moved backward with increasing Ma₂ (Ma₂ > 1). Lastly, a strong adverse pressure gradient was formed behind the shock wave, which led to the beginning of the transition.