The scale difference between the real flight vehicle and the experimental model results in the Reynolds number effect, which makes it unreliable to predict the aerodynamic characteristics of flight vehicles by wind tunnel testing. To understand the mechanism of Reynolds number effects on the aerodynamic characteristics of the supercritical wing that is commonly used in transport aircraft in more detail, surface pressure wind tunnel tests of a transport aircraft reference model with a wing-body configuration were conducted in the European Transonic Windtunnel (ETW) at different Reynolds numbers. There are 495 pressure taps in total equipped on the surface of the test model with the Mach numbers ranging from 0.6 to 0.86 and Reynolds number varying from 3.3 × 10⁶ to 35 × 10⁶. In addition, an in-house developed CFD tool that has been validated by extensive experimental data was used to correct the wing deformation effect of the test model and achieve detailed flow structures. The results show that the Reynolds number has a significant impact on the boundary layer displacement thickness, surface pressure distribution, shock wave position, and overall aerodynamic force coefficients of the transport aircraft in the presence of shock wave and the induced boundary layer separation. The wind tunnel data combined with flow fields achieved from CFD show that the essence of the Reynolds number effect on the aerodynamic characteristics of transport aircraft is the difference of boundary layer development, shock wave/boundary layer interaction, and induced flow separation at different Reynolds numbers.

A simultaneous surface pressure and displacement measurement method that integrates pressure-sensitive paint (PSP) and binocular stereophotogrammetry is proposed. The assays were completed on the Φ4 m rotor test stand at China Aerodynamic Research and Development Center (CARDC). A single-shot lifetime approach was utilized to acquire the instantaneous pressure field on a rotor blade coated with PSP. At the same time, the PSP feature points were used to obtain the 3D coordinates of stereo cameras, which yielded the blade displacement field. The experimental results showed that the displacement measuring accuracy was better than 0.2 mm, and the pressure measurement accuracy was not affected, with Standard Deviation (STD) values below 700 Pa. The advantages of the proposed system are its simple structure, low cost, high accuracy and high test efficiency, which will offer a practical solution for the exploration of fluid–structure interplay. Hence, such a system is a prospective for the wind tunnel tests of helicopter rotor blades.

Pressure-sensitive paint (PSP) is an important wind tunnel testing technology. Compared with conventional PSP, the performance and test accuracy of cryogenic PSP are still immature. Therefore, investigating how to improve the pressure sensitivity of cryogenic PSP and reduce the interference of temperature effect is of great significance. By studying the PSP luminescence characteristics at different time points, temperatures, and pressures, some interesting phenomena have been discovered. When the temperature reaches 323 K, PSP can accelerate aging, leading to significant and irreversible changes in coating performance. Additionally, the effect of temperature on the luminescence characteristics of PSP shows significant differences over time. This unusual phenomenon may be related to the microstructure change in PTMSP (PtTFPP) coatings over time. In the beginning, PTMSP coating has high activity and spacing between PtTFPP luminescence centers, which change significantly with the microstructure as the temperature decreases. This might result in a stronger concentration quenching of PtTFPP, which counteracts the expected enhancement of luminescence efficiency typically caused by the temperature decrease. After 72 h, the microstructure of the coating tends to be stable, and the effect of temperature on the fluorescence characteristics of PSP becomes a thermal quenching law similar to that of traditional PSP. This discovery can provide a more precise basis for correcting the temperature effect for cryogenic PSP coatings with varying service lives.

For studying the complex flow structure of a high speed cavity, including shear layer, separation, reattachment and vortex, the single-color fluorescent oil flow visualization system in the 0.6m×0.6m trisonic wind tunnel of China Aerodynamic Research Development Centre was enhanced. Different portions of the cavity such as its bottom wall, left and right side walls, front and rear walls and area outside the cavity utilized different oil films with different fluorescent particles to visualize the flow mixing more effectively in the cavity. Specialized ultraviolet light sources are used to enhance the oil flow image contrast. Additionally, to ensure the oil film follows the surface streamlines and indicates the skin friction lines more accurately, the thickness and viscosity of the oil film are well controlled. A high-speed cavity wind tunnel test was conducted to validate the improved oil flow visualization system at Mach number ranging from 0.6 to 2.0. The experimental result shows that the improved oil flow system can effectively visualize complex flow patterns of the cavity flow at all Mach numbers. Colour fluorescent oil flow visualization system established in the present study provides a capability to visualize the flow mixing phenomenon in cavity flows with a high image contrast.

This study investigates the effect of Reynolds number on unsteady buffet characteristics of the OAT15A supercritical airfoil under transonic conditions (Ma = 0.73, AOA = 3.5°) using DDES based on the SST k-ω turbulence model coupled with the γ-Reθ transition model. Results show that, compared with fully turbulent conditions, the free-transition cases exhibit a more downstream shock position and higher lift. Under fully turbulent conditions, higher Reynolds numbers drive the shock downstream and enhance its stability. Under free-transition conditions, the shock moves downstream at low Reynolds numbers but shifts upstream at high Reynolds numbers due to changes in the transition location. During the unsteady buffet cycle at low Reynolds numbers, the lift increases as the shock moves downstream and the separation region shrinks. The lift reaches its maximum when the separation is minimal, corresponding to a quiet flow state with weak acoustic emission. As the lift decreases, a large separation region forms behind the shock, forcing the shock upstream and reducing the lift to its minimum. At high Reynolds numbers, the buffet cycle changes: the shock becomes more stable; trailing-edge vortex shedding intensifies; lift oscillation amplitude decreases; and buffet frequency increases.

Shock buffet is one of the critical issues affecting the aerodynamic performance, flight quality, and flight safety of large aircraft. To overcome the limitations of traditional experimental measurement methods, such as insufficient capability in capturing flow features and high cost, an integrated experimental system tailored for extreme cryogenic and high-Reynolds-number conditions is developed based on the conventional tuft technique. This system comprises “preparation of low-flow-disturbance fluorescent mini-tufts, high-efficiency large-area tuft taping, automatic generation of digital streamline, and flow topology analysis”. Furthermore, a technique for assessing the transonic shock buffet onset using dynamic flow visualization with fluorescent mini-tufts is proposed. This paper takes a typical supercritical airfoil as the research object. First, through high-precision numerical simulations, it reveals that low-energy, unstable boundary-layer separation is the core driving force for the development and maintenance of shock buffet, and that flow separation characteristics serve as an important basis for determining the shock buffet onset. Subsequently, experimental validation is conducted in a 0.3 m high-Reynolds-number transonic wind tunnel. Using a dual-excitation-band composite light source, simultaneous measurements of pressure-sensitive paint (PSP) and fluorescent mini-tuft patterns are realized. The experimental results show that under extreme conditions, characterized by a wide total temperature range of 110 K to 280 K and strong scouring at Mach numbers from 0.6 to 0.9, the fluorescent mini-tufts (approximately 0.05 mm in diameter) exhibit excellent flow-following capability without any detachment. The digitized flow patterns of the fluorescent mini-tufts, obtained via computer image recognition algorithms, clearly reveal the location and area of boundary-layer separation. The trends show good agreement with the cryogenic PSP results, providing an important reference for determining the shock buffet onset.