The unsteady characteristics of the second throat of a transonic wind tunnel have an important influence on the design and test of the wind tunnel. Therefore, the forced oscillation characteristics were studied by a numerical simulation method. The governing equation was the viscous compressible unsteady Navier–Stokes equation. Under the sinusoidal pressure disturbance of the computational domain exit, the shock wave presents a clear forced oscillation state, and the shock wave periodically changes its position. Under a pressure disturbance of 1%, the shock wave displacement reaches 150 mm. Additionally, overshoot occurs when the shock moves upstream or downstream. The shock-boundary layer interference is very sensitive to the motion characteristics of the shock wave, resulting in a transformation of the flow field symmetry. The flow field downstream of the shock wave exhibits periodic structural changes. Compared with the pressure change at the outlet, the pressure change near the shock wave has a phase delay. The increasing disturbance near the shock wave shows a clear amplification effect. The pressure disturbance near the shock wave had an obvious amplification effect, and its fluctuation amount reached 16% under the pressure disturbance of 1%. The variation trend of the second throat wall force, wavefront Mach number, and Mach number in the test section with time is similar to that of the downstream disturbance, but it does not have a complete follow-up effect, which indicates that the pressure disturbance can propagate into the test section through the boundary layer or the shock gap. Nevertheless, the second throat choking can still control the Mach number stability of the test section. The dynamic characteristics of shock oscillation are related to the amplitude and frequency of the applied pressure disturbance. The shock displacement decreases with the increase in the excitation frequency. When the excitation frequency is higher than 125 Hz, the flow field basically does not change.

Supersonic semi-flexible wall nozzles with the ability to continuously vary the Mach number have become a key configuration for advanced supersonic wind tunnels, with their contour design typically based on the Sivells method. However, this method faces two persistent challenges in its application: (i) suboptimal flow field uniformity, particularly during continuous Mach number variation, and (ii) localized Mach number depression (“dimples”) along the nozzle axis. This paper presents an enhanced design methodology to overcome these challenges. The proposed improvements incorporate two key elements: (1) exact analytical solutions for transonic flow to replace the conventional source-flow approximation used in the Sivells method at the throat section, and (2) a Bézier-optimized characteristic method for expansion contour design. Experimental results demonstrated that the improved design method significantly enhanced performance even in a small 0.3 m-scale wind tunnel: flow field uniformity reached 0.0010–0.0053 in fixed-geometry mode (Ma = 1.2–3.0), while dynamic flow uniformity was maintained at 0.002–0.010 in continuous morphing mode (Ma = 1.5–3.0). These research outcomes provide a reliable technical approach for designing supersonic wind tunnel semi-flexible nozzles.

A static pressure probe is a crucial tool for measuring static pressure fluctuations, and its internal cavity structure can significantly affect the accuracy of the data obtained. This study investigates the impact of the static pressure probe cavity on the frequency response characteristics of fluctuating pressure using both experimental and numerical simulations. The results are validated by comparing them with the behavior of a second-order system. Our findings indicate that the internal cavity of the static pressure probe acts as a second-order underdamped system. This system amplifies fluctuating pressure signals at frequencies below the characteristic frequency while attenuating those above it. The frequency response characteristics of the probe’s cavity are similar to those of a Helmholtz resonator. Among various factors, the diameter of the pressure tap within the cavity has the most significant effect on the system’s characteristic frequency and amplification ratio. By optimizing the design of the static pressure probe’s cavity dimensions, the precision of fluctuating pressure data below the system’s characteristic frequency can be improved. In the research, based on the Helmholtz resonance equation, we provide a semi-empirical formula for predicting the characteristic frequency of fluctuating pressure in a static pressure probe. Furthermore, by leveraging the mechanism of the static pressure probe’s cavity as a second-order underdamped system, we propose a method for rapid and accurate calibration of the static pressure probe’s fluctuating pressure measurements.

To meet the requirement for the low disturbance of wind tunnel flow in supersonic boundary layer transition simulation tests, a low-disturbance supersonic research wind tunnel with a nozzle size of 0.3 × 0.15 m2 and test Mach number of 3.0 was built. It was then used to study key design technologies for increasing the size and Reynolds number of low-disturbance supersonic wind tunnel. The wind tunnel adopted a “blowdown + ejection” layout to achieve a wide pressure range of 50–450 kPa. A “sintered wire mesh + honeycomb + damping mesh” structure was adopted in the settling chamber with the flow straightening method to achieve velocity pulsations of less than 0.3% and pressure pulsations of less than 0.1% at the exit. A comprehensive design technique for the supersonic laminar nozzle was adopted, which included nozzle boundary layer suction and highly polished surface nozzle processing. At a total pressure of 450 kPa, a low-disturbance supersonic test area with an axial length of about 9 cm was realized. The airflow pressure pulsations in this area were less than 0.15%, and the Reynolds number based on the characteristic length reached 1.3 × 106.