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The objective of this study is to enhance the methodology for the design of a supersonic wind tunnel, improving the process with advanced computational techniques. The supersonic wind tunnel is intended to operate within a flight envelope of Mach 2.5 to 4 and altitudes between 18 and 20 km; this study focuses on the operative condition of Mach 3.5. The research is based on computational fluid dynamics, enabling a deeper understanding of fluid flow phenomena that can deteriorate the operability of the wind tunnel. Additionally, a detailed mesh independence study has been conducted to ensure the reliability and robustness of the computational results. These new analyses allowed for a more comprehensive optimization in the state of the art of tunnel geometry and operational conditions, further enhancing the ability to sustain supersonic flow for extended durations. Particular attention was given to the second throat, which plays a crucial role in the overall performance of the facility, especially during the start-up process. Its design has been refined to improve efficiency by reducing the minimum starting pressure.

This paper presents an impressive design of a convergent divergent (C-D) nozzle using the method of characteristics for a Mach number 2 test section. The nozzle’s geometry was meticulously crafted in SolidWorks, and its performance was evaluated through a CFD simulation in Ansys Fluent R22 software. Results showed excellent agreement between the simulation and analytical data, with the Mach number ranging from 1.78 to 2. The study also compared turbulence modeling techniques, concluding that the k-omega model produced superior results. The supersonic wind tunnel achieved remarkable efficiency, completing a run at 1.8 Mach number in just 6 seconds. Overall, the study showcased exceptional accuracy and meticulousness.

Transonic/supersonic wind tunnels are indispensable equipment for advanced aircraft to operate across subsonic, transonic, and supersonic regimes. The deformation of the flexible nozzle is the key to accurately controlling the Mach number of transonic wind tunnels. However, solving the deformation of flexible wall plates remains challenging due to the highly nonlinear relationship between wall loading and deformation, as well as the lack of simple yet effective mathematical models under complex boundary conditions. To accurately describe the deformation of flexible wall plates and improve computational efficiency, this study systematically investigates the deformation characteristics of flexible walls in two orthogonal directions and proposes an orthogonal beam function (OBF) model for characterizing small-deflection deformations. For large-deflection deformations in a flexible wall, an elliptic integral (EI) solution is introduced, and the OBF model is correspondingly modified. Experimental validation confirms that the OBF model effectively describes large-deflection deformations in a flexible wall. This research contributes to solving large-deflection deformation in flexible wall plates, enhancing both computational efficiency and accuracy.

A new fuel injection system with a solenoid actuator was developed to reduce jet stabilization times and was compared to a conventional fuel injection system using a fast-acting valve. The jet stabilization time was examined from free-jet images obtained using a flow visualization technique. Using a supersonic wind tunnel, the jet establishment image and trajectory for supersonic flows were obtained. The jet stabilization times were obtained at fuel tank pressures of 500, 700, and 900 kPa and injection orifice diameters of 1 and 2 mm. For all cases, the jets were stabilized at 4–12 ms, which was shorter than the stabilization time (approximately 300 ms) of the conventional system. The injection pressure measurements and jet trajectory comparison showed that the new system could inject fuel at the target pressure more effectively than the conventional fuel system.

The present study proposes to design, fabricate and investigate a miniature supersonic wind tunnel at an affordable cost, safe to operate, and reliable for education purposes. The configuration consists of a high pressure reservoir, and mini tunnel which will be used for providing the supersonic flow over aerodynamic models. To design the mini supersonic wind tunnel, the method of characteristics is used to ensure that any shocks that arise in contact with the surface do not impact the uniformity and velocity of the flow through the duct. To achieve the supersonic flow, the nozzle contour of the supersonic portion of the wind tunnel is built in such a way that the shocks nullify each other and have zero effect on the velocity of the flow. For compressible flow visualization, an affordable Schlieren setup consisting of a torch, mirrors, and a mobile camera is used for flow visualization. Finally, the visualization results of the present study is in good relation with the theoretical calculations. The present facility is also calibrated by measuring the oblique shock angle in the test section, in the range of Mach number 1.6 to 1.8.

The dynamic response of a thin buckled panel in a supersonic wind-tunnel experiment is investigated. Measured time histories of the panel displacement and velocity show co-existing, nonlinear responses with features of periodic and chaotic oscillations. Fully coupled computational analyses are conducted in order to study and interpret the aeroelastic phenomena observed during the experiments. A computationally efficient modeling framework is formulated with a nonlinear structural reduced-order model and enriched piston theory aerodynamics for the mean flow. The simulations predict the onset of the chaotic motions observed in the experiments, albeit with an approximately 21% increase in the oscillation amplitude. A linearized equation governing the distance between neighboring solutions is derived and used to compute the largest Lyapunov exponent in order to prove the existence of chaos. A modified Riks analysis highlights the co-existence of multiple equilibrium positions which predisposes the nonlinear system to chaos. The system’s sensitivity to cavity pressure, temperature differential, and initial conditions is also investigated. Variation of the cavity pressure and temperature differential yields additional regions of dynamic activity that were not explored during the experiments.

The National Institute of Metrology Thailand (NIMT) has continuously expanded its facilities to support aerodynamic research and innovation, addressing both subsonic and supersonic testing needs. This development includes the establishment of a Force-Balanced Piston Gauge (FPG) for high-precision low differential pressure measurement, which is crucial for determining airspeed and pressure distribution in subsonic wind tunnel testing. Additionally, a portable Eiffel-type subsonic wind tunnel has been developed to facilitate the testing and calibration of pitot tube and hot-wire anemometers, ensuring accurate airflow measurements while providing flexibility for on-site applications. To enhance dynamic pressure calibration capabilities, NIMT has also developed a shock tube for calibrating pressure sensors under rapidly changing conditions and studying high-speed shock wave propagation in supersonic aerodynamic applications. Furthermore, NIMT has supported the development of a supersonic wind tunnel to strengthen Thailand’s research capabilities in supersonic flow and shock wave phenomena, enabling more advanced aerodynamic studies and aerospace applications. Looking ahead, NIMT plans to establish a space vacuum chamber, designed to serve as a test stand for aerospace mini thrusters, further expanding Thailand’s capabilities in space-related metrology and propulsion system research. These advancements contribute to strengthening Thailand’s metrological infrastructure, supporting cutting-edge research and development in aeronautics, aerospace, automotive engineering, and various industrial applications.

Understanding how protrusions, such as fins attached to flat or streamlined bodies, affect aerodynamics, especially in high-speed contexts, is vital for aerospace applications. These protrusions significantly influence overall aerodynamics and require a comprehensive understanding for accurate analysis and prediction of aerodynamic performance. This understanding is particularly critical in supersonic flight, where even minor aerodynamic disturbances can impact vehicle stability and efficiency. Therefore, a thorough understanding of protrusion-induced flow phenomena is essential for advancing aerospace engineering and improving supersonic vehicle performance and safety. The present paper focuses on the complex supersonic flow over a vertical fin, using a combination of experimental and computational methods. The study aims to understand how variations in fin height influence the behavior of the Lambda shock and any resulting changes in shock length. Specifically, the paper investigates different fin height-to-diameter (H/D) ratios ranging from 0.5 to 1.5 in steps of 0.25. To achieve this, both experimental testing in a supersonic wind tunnel and numerical simulations using the commercial CFD tool ANSYS-FLUENT are employed. Through this dual approach, the paper seeks insights into the characteristics of the Lambda shock and its effects on key aerodynamic parameters, such as shock strength and drag coefficient. By thoroughly investigating these aspects, the paper contributes to a deeper understanding of the complex flow phenomena associated with supersonic flow over vertical fins, potentially guiding the design and optimization of aerospace vehicles. The outcomes indicate that a fin height of 12 mm (H/D=1.0) provides the best balance in terms of pressure distribution, Lambda shock length, and drag coefficient, making it the optimal choice for enhancing aerodynamic stability and performance in supersonic conditions.

The transient response of a plate and a cavity is investigated in a supersonic wind tunnel start experiment where the freestream flow inside the test section reaches turbulent flow at Mach 2. Experimentally measured plate displacement time history shows flutter onset, transition to limit cycle oscillation, and stabilization at a static deformed state during the 30 s run. To analyze and interpret the measured plate response, a fully coupled aero-thermal-acousto-elastic analysis is carried out. A theoretical–computational model is formulated with a nonlinear structural plate model, acoustic pressure equation for the stationary fluid in a cavity, and the first-order Piston Theory aerodynamics. A linear stability analysis is performed that includes the nonlinear added stiffness due to an initial deformation to investigate the combined effects of freestream coupling and temperature differential on system stability. Also, direct time integration of the nonlinear fluid structural equations of motion is performed using experimentally measured flow parameters as inputs. All stability transitions are captured using the theoretical model with good agreement with experiment for transitions from no flutter to flutter/limit cycle oscillations (LCO) although the theoretical LCO amplitude is approximately 50 % larger than measured. The system’s sensitivity to cavity coupling, temperature differential, thickness calibration, static pressure differential, and turbulent pressure fluctuations are investigated. Lastly, snap-through buckling analyses in response to periodic and quasi-static excitations are conducted.

3-D DNS code based on Message Passing Interface (MPI) is developed to study channel turbulence. To solve 3-D conservative Navier-Stokes equations, the WENO-SYMBO3 method and the third-order Runge-Kutta method are adopted for spatial derivative and temporal integration, respectively. The WENO-SYSBO3 method is able to lower the dissipation errors during the discretization process and simulate turbulent flows more accurately. Meanwhile, as a shock-capturing scheme, it can accurately capture flow discontinuities. Therefore, it has unique advantages in compressible flow simulations. Typical flow problems, i.e. supersonic wind tunnel with a front step, supersonic uniform flow around a cylinder, and 2-D Poiseuille flow are conducted to validate the code’s performance. Channel turbulence was simulated through the code. The outcomes reveal that the average temperature within the wall boundary layer increases, the average density decreases, and both the turbulence intensity and the correlation coefficient increase significantly.