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The presence of large-scale coherent structures in various wall bounded turbulent flows, often called superstructures in turbulent boundary layers (TBLs), has been of great interest in recent years. These meandering high- and low-momentum structures can extend up to several boundary layer thicknesses in the streamwise direction and contain a relatively large portion of the layer's turbulent kinetic energy. Therefore, studying these features is important for understanding the overall dynamics of turbulent boundary layers and for the development of flow control strategies or near-wall flow modifications. However, compared to the extensive number of incompressible investigations, much less is known about the structural characteristics for compressible turbulent boundary layer flows. Therefore, in this investigation turbulent boundary layers developing on a flat plate with zero pressure gradient (ZPG) over a range of Reynolds numbers and Mach numbers are considered in order to examine the effect of compressibility on superstructures. More specifically, measurements are performed on a flat plate model in the Trisonic Wind Tunnel Munich (TWM) for the Mach number range $0.3 \\leq Ma \\leq 3.0$ and a friction Reynolds number range of $4700 \\leq Re_{\\tau } \\leq 29\\,700$ or $11\\,730 \\leq Re_{\\delta _2} = \\rho _e u_e \\theta ^*/\\mu _{w} \\leq 74\\,800$. Velocity fields are recorded using planar particle image velocimetry methods (PIV and stereo-PIV) in three perpendicular planes. Using multi-point correlation and spectral analysis methods it was found that the most energetic frequencies have slightly longer streamwise wavelengths for the supersonic case when compared to the subsonic case. Furthermore, a distinct increase in the spanwise spacing of the superstructures was found for the supersonic cases when compared to the subsonic and transonic turbulent boundary layers.

Recent experience has shown that test results of standard wind tunnel models under off-design conditions could be a useful aid in preparations of some nonstandard wind tunnel tests. However, reference data for such conditions do not exist, or they are scarce. Therefore, off-design transonic wind tunnel tests of the HB-2 standard models were executed in the VTI T-38 wind tunnel as a supplement to the supersonic tests of the same models under design-intent conditions, for which reference results were available. New tests were conducted so that test envelopes partially overlapped with those from available supersonic reference data. Good agreements of results with references were confirmed in the overlapped ranges, so it was assumed that, by implication, the obtained results were also valid in the transonic range of conditions, with an observation that the effects of sting diameter were much more pronounced in the transonic range than in the supersonic one. HB-2 models were tested in two sizes, using two different wind tunnel balances for each model, so that the results can be used with more confidence.

The Tri Sonic wind tunnel under consideration for this study will include...

Re-entry capsules' success depends significantly on dynamic and static stability, particularly before deploying the main parachute. Determining the range of dynamic instability and investigating the underlying causes is crucial for designing the entry capsule's control system. Dynamic stability is analyzed in this study based on pitch moment coefficients obtained from forced oscillation experiments conducted in the trisonic wind tunnel for the Orion entry capsule. The results reveal that pressure fluctuations at the aftbody of this model begin at Mach 2. The findings and other research results emphasize the significant role of the aftbody geometry in generating dynamic instability at low supersonic speeds due to its interaction with vortex flow. The results also demonstrate that increasing the Mach number to 2.2 would result in a near zero-pressure coefficient on the capsule's aftbody, which implies that there is no acting force on the aftbody. The results show that as the freestream Mach number increases from M∞= 1.8 to M∞= 2.2, the pressure on the aftbody remains unchanged during the pitching motion due to approaching the shear layer towards the body and consequent shrinking of the aftbody vortex. Furthermore, the sensitivity of dynamic stability to the mean angle of attack was investigated. It is shown that a slight increase of approximately 5 degrees in the mean angle of attack can considerably enhance the re-entry capsule's dynamic stability.

This paper outlines an experimental investigation conducted at the INCAS trisonic wind tunnel, focusing on the determination of pitch damping coefficient. The model used for this investigation is the Basic Finner Model, a standard model for dynamic tests which consists in a cone-cylinder body with four rectangular fins. The study aims to evaluate the influence of various parameters—including the Mach number, angle of attack, reduced frequency, center of rotation, and roll angle—on pitch damping coefficient. The employed method for determining these coefficients is the free oscillation method which consists in measuring the model oscillation in free stream after an initial perturbation. In order to perform these dynamic tests in the wind tunnel, a dedicated rig was developed to initiate the model’s oscillation using a linear servo-actuator and to record its oscillation using a strain gauge. The results obtained from the experiments illustrate how each parameter impacts the pitch damping coefficient, highlighting the precision of the measurements. The paper’s conclusion presents that the developed rig and the method used provide accurate results, and the variation in different parameters can change the damping coefficient.

Purpose The purpose of this study is to determine the impact of two parallel boosters fixed to the ILR 33 AMBER 2 K core rocket stage on its aerodynamic characteristics in the subsonic and transonic regimes and for M = 2.3. Design/methodology/approach Wind tunnel tests of the rocket model were carried out in a trisonic wind tunnel using a six-component internal balance. Three rocket model configurations were investigated. Findings The results of the presented studies showed that the presence of boosters causes a significant increase in the total rocket drag, which depends on both the Mach number and the rocket flight phase. Experimental tests of the rocket model allowed to determine the difference in drag coefficient between active and passive flight versus Mach number. It was found that, in the case of a deviation from the rocket’s flight direction, the aerodynamic coefficients strongly depend on the location of the boosters in relation to the direction of the deviation. Practical implications Studies of the rocket model aerodynamic characteristics allow the assessment of the influence of parallel boosters on rocket performance, which is important when the decision of a rocket staging type is taken. Originality/value The presented wind tunnel test results of the rocket model equipped with the two parallel boosters are an original contribution to the rocket research results presented in the literature.

Wind tunnels remain an essential element in the design and development of flight vehicles. However, graduates in aerospace engineering tend to have had little exposure to the demands of industrial experimental work, particularly at high speed, a situation exacerbated by a lack of up-to-date reference material. In an attempt to fill this gap, this paper presents an overview of the current and near-term status and usage of transonic industrial wind tunnels. The review is aimed at recent entrants to the field, with the aim of helping them make the step from research projects in small university facilities to commercial projects in large industrial facilities. In addition, a picture has emerged from the review that contradicts received wisdom that the wind tunnel is in decline. Globally, the industrial transonic wind tunnel is undergoing somewhat of a renaissance. Numbers are increasing, investment levels are rising, capabilities are being enhanced, and facilities are busy.

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.

The existing analytical methods for determining and correcting the effect of wind tunnel walls on experimental data are based on the hypothesis of potential flow. A useful simplification of the boundary conditions used to describe the perforated walls was to consider the wall as homogeneous, the solid and free portions not being treated separately, but as an equivalent permeable surface. The approximation of the wall behaviour during the experiment was possible by defining a porosity parameter. The purpose of this paper is to estimate the porosity parameter for the perforated walls of a trisonic wind tunnel by evaluating the pressure distributions measured on the walls of the test section.

Artificial intelligence (AI) can be used to optimize the prediction of pressure fluctuations over the external surfaces of aerospace launchers and minimize the number of wind tunnel tests. In the present research, various machine learning (ML) techniques capable of predicting the acoustic load were tested and validated. The methods included decision trees, Gaussian Process Regression (GPR), Support Vector Machines (SVMs), artificial neural networks (ANNs), linear regression, and ensemble methods such as bagged and boosted trees. These algorithms were trained using experimental data from an extensive wind tunnel test campaign conducted to support the design of a VEGA (Advanced Generation European Vehicle) launcher vehicle and provide wall pressure fluctuations in many configurations. The main objective of this study was to identify, among several algorithms, the most suitable method able to process such complex databases efficiently and to provide reliable predictions. Different statistical indices, including the root mean square error (RMSE), the mean square error (MSE), and a correlation coefficient (R-squared), were employed to evaluate the performance of the ML methods. Among all the methods, the bagged tree algorithm outperformed the others, providing the most accurate predictions, with low RMSE and high R-squared values across all test cases. Other methods, such as the ANNs and GPR, exhibited higher errors, indicating their reduced suitability for this dataset. The results demonstrate that ensemble decision tree methods are highly effective in predicting acoustic loads, offering reliable predictions, even for configurations outside the training database. These findings support the application of ML-based models to optimize experimental campaigns and enhance the design of aerospace launch vehicles.