Supersonic Wind Tunnel Research Articles

Unsteady Flow Characteristics in an Over-Under TBUnsteady Oscillation and hysteresis characteristics of high-speed duct during TBCC inlet mode transition under mildly throttled condition

This study aims to investigate the intricate dynamic characteristics of the high-speed duct during the over-under Turbine-Based Combined Cycle TBCC inlet mode transition process while operating in an off-design state under throttled conditions. A typical over-under TBCC inlet, designed for a working Mach number range of 0–6 with a transition Mach number of 3.5, is examined through experimental studies in a supersonic wind tunnel with a freestream Mach number of 2.9. The investigation focuses on the complex oscillatory flow and unique hysteresis observed in the mode transition process of the high-speed duct under the mildly throttled condition, utilizing high-speed Schlieren and dynamic pressure acquisition system. The findings reveal that the high-speed duct undergoes four distinct oscillation stages akin to those in a higher throttled state during the mode transition, albeit with smaller dominant frequency and energy. Moreover, an irregular alternating “big/little buzz” mode is observed in the early stage of the large oscillation stage. Notably, the mildly throttled state exhibits three intriguing hysteresis properties compared to the unthrottled and higher throttled states. Firstly, hysteresis is observed in the shock train motion stage in the duct before unstart, along with the corresponding inverse process. Subsequently, hysteresis is noted in the unstart and restart of the high-speed duct, with a smaller hysteresis interval than in the unthrottled state. Finally, the hysteresis characteristics of oscillation mode switching and the corresponding inverse process are explored. Based on the analysis, the first two hysteresis phenomena are associated with the formation and dissipation of the separation bubble. The significant adverse pressure gradient constrains the cross-sectional capacity of the channel, rendering the disappearance of the separation bubble more challenging. The hysteresis in oscillation mode switching is linked to not only the channel cross-sectional capacity but also the state of the incoming boundary layer.

Fiber Bragg Grating-based sensor for measuring supersonic inlet terminal shock at different angles of attack

This study presents a multipoint pressure measurement system using fiber Bragg grating sensors to detect terminal shock positions. The system consisted of a fiber optic string encapsulated with multiple diaphragm-based fiber Bragg grating sensors, capable of detecting flow field characteristics at fourteen different axial positions of the inlet. The calibration results indicated a sensitivity of 46.47 pm/kPa, a linear correlation coefficient of 0.9997, and a maximum repeatability error below 1.44%. In supersonic wind tunnel tests, the system effectively measured pressure and located the terminal shock across varying angles of attack and throttling degrees. The maximum difference compared to conventional pressure sensors was 4.7%, and the terminal shock positioning error was limited to 1.2%. Overall, the system has robust multipoint measurement capabilities and is priced at only one-tenth of traditional sensors, showing great potential in the field of inlet measurements.

Evolution of supersonic cooling film flow over concave and convex surfaces with different radius curvatures

A cooling film (Mach 2.3) is injected into a supersonic wind tunnel's (Mach 3.8) flow field. As the curvature radius decreases, mixing layer destabilization is delayed on the CV (convex wall) and advanced on the CC (concave wall). When x = 20–60 mm, wall pressure is influenced by the cooling film, and when x = 100–220 mm, curvature dominates. As the curvature radius decreases, pressure on the concave wall increases more rapidly, while that on the convex wall decreases more swiftly. IP (impulse for bulk dilation) values on the CC-1500 and CV-1500 walls are approximately twice those on the CC-3000 and CV-3000 walls, respectively.

Influence of Rarefaction Degree and Aft-Body Geometry on Supersonic Flows

During atmospheric entry, super-/hypersonic vehicles cross distinct atmospheric layers characterized by large density variations and thus experience different flow regimes ranging from free molecular, transition, slip, to continuous regimes. Due to the distinct modeling strategy between these regimes and complex physical phenomena appearing near the vehicles (boundary-layer/shock interaction, base-flow recirculation, etc.), assessing their aerodynamic properties may be difficult. The present work focuses on supersonic flows around sharp-base geometries in both continuous and slip-flow regimes and aims at highlighting the influence of both rarefaction degree and base geometry on the vehicles’ aerodynamic features. For this purpose, three axisymmetric cone-cylinder geometries with right-angled, rounded, or flared rear parts are considered. Flow visualization, pressure, and drag measurements are carried out at Mach number Ma=4 and Knudsen numbers ranging from Kn=4×10−4 to 1×10−2 in the supersonic rarefied MARHy wind tunnel. The experimental data are compared with numerical results of simulations performed with a continuous-flow Navier–Stokes (NS) solver and two rarefied flows codes: a discrete-ordinate Bhatnagar–Gross–Krook (K) solver and a direct simulation Monte Carlo (SPARTA) solver. While the NS solver overestimates frictional drag as Kn rises, the rarefied K and SPARTA results show satisfactory agreement with experimental data. The latter numerical results highlight the main effects of rarefaction: as Kn increases, shocks become more diffuse, skin friction strengthens (leading to a significant increase in drag coefficients), and the extent of the base-recirculation decreases. Regarding the aft-body geometry, its influence on the base recirculation vanishes with increasing Kn.

Design of a Wind Tunnel Nozzle Section Trolley Control System Based on Backstepping Method

Abstract To accommodate different types of supersonic wind tunnel tests, frequent replacements of nozzle sections and test sections are required, involving multiple complex technical processes. To optimize operational procedures and improve efficiency, a control system for the wind tunnel nozzle section trolley based on the backstepping method was designed. The system uses radio frequency identification (RFID) signals for station positioning and intelligently adjusts the speed and direction of the control motor, thus driving the trolley to move back and forth automatically. Through the simulation of the motor motion control model, the system effectively reduces random disturbances in the motion system caused by load changes, achieves precise positioning of wind tunnel plugins and nozzle assemblies, enhances system operating accuracy, reduces production cycle times, and realizes a one-button positioning function for all components, playing a significant role in improving wind tunnel test efficiency and operational quality.

Synergizing computer‐aided design, commercial software, and cutting‐edge technologies in an innovative nozzle test apparatus for an engineering laboratory course

AbstractThis study explores compressible flow, a field reliant on mathematical models for effective teaching. Using laboratory experiments as pedagogical tools, we introduce a compact nozzle test apparatus that integrates cutting‐edge technologies—additive manufacturing (AM), pressure‐sensitive paint, the Schlieren system, image processing, and computational fluid dynamics (CFD)—in a compressible flow laboratory course. Commercial software, including MATLAB, SOLIDWORKS, ANSYS Fluent, and LabVIEW, facilitates the incorporation of these technologies. The research outlines the course structure, objectives, and details of student projects. Through a comparative analysis of experimental results, analytical calculations, and CFD simulations, we showcase the successful integration of AM in pedagogical practices for compressible flow, addressing critical concerns like nozzle strength and surface roughness. Statistical data from student projects offer practical insights, ensuring accuracy in experimental applications. The laboratory's design and detailed lists of components and costs provide a meaningful comparison with a supersonic wind tunnel, considering manufacturing expenses, operational costs, spatial requirements, and noise levels. The assessment of lab report grades underscores the approach's efficacy in conveying compressible flow concepts successfully, facilitated by modern computers. In summary, our study presents a comprehensive, efficient, and technologically advanced approach to teaching compressible flow within a concise framework.

Nanoengineering-Enhanced Capillary Cooling Achieves Sustained Thermal Protection for Ultra-High Heat Flux and Temperature.

Extreme thermal conditions with heat flux densities exceeding 1 MW·m-2 or temperatures reaching up to 1000 °C are prevalent in various situations. However, the ability of thermal protection either depends on specialized materials or is currently limited with existing cooling schemes. Herein, we propose an innovative cooling scheme that relies on evaporation-driven capillary flow enhanced by nanoengineering-designed porous structures with common materials. Experimentally-obtained capillary flow cooling curve identifies critical heat flux corresponding to evaporation-driven flow stage, where coolants cool the surface and subsequent vapor impedes heat transfer from thermal boundaries. Nanoengineering provides opportunities for enhanced capillary flow, which proves to endow bronze, TC4, and Al2O3 with thermal protection ability 50%-180% higher than that without nanoengineering-designed. Our scheme achieves critical heat flux up to 2.0-3.1 MW·m-2, and performs thermal dissipation capacity almost twice higher than inherent latent heat of coolant. Furthermore, in a supersonic wind tunnel with total temperature reaching up to 1792 K, our scheme effectively protects surfaces by cooling them to surface temperatures below 500 K. Nanoengineering-enhanced capillary cooling gives access to the application of common materials for high-temperature and high-heat-flux environments and paves the way for the development of lightweight, long-lasting, and large-scale solutions for thermal protection. This article is protected by copyright. All rights reserved.

Failure tolerant functionally graded fiber-reinforced UHTCs exposed to hypersonic flows

A functionally graded ZrB2 ceramic containing increasing content of short carbon fiber, 0–60 vol%, was tested in a supersonic arc-jet wind tunnel by exposure to flow in simulated air. The composite, that exhibited increased failure tolerance due to the combined toughening effect of the fiber bridging and the residual stresses accumulated between scales of different composition, successfully survived 2120 K for 300 sec, whereas scale spalling occurred upon reaching 2930 K. Overall, in both conditions, carbon fiber remained protected and with unchanged morphology, providing the necessary structural stability for surviving the shocks encountered by the hottest components of new generation aerospace vehicles.

Representation of Thermomechanical Damage in Fiber-Reinforced Ceramic Composites at High Thermal Gradients

Ceramic matrix composites (CMCs) are the core strategic materials for the thermal structure of new-generation hypersonic aircraft. The unclear understanding of the cross-scale thermomechanical behavior and failure mechanism of materials and structures in high-temperature transient environments is a key scientific issue that restricts the reliable design and safe operation of hypersonic aircraft. In this study, a high-temperature test module in a supersonic wind tunnel with integrated various test systems was independently designed to systematically investigate the thermal mechanical damage evolution of CMCs used in astronautics applications under high-temperature transient conditions. Furthermore, a numerical method for fluid-heat-solid coupling based on an iterative solution strategy was developed and compared with an algorithm based on a wall modification strategy in terms of computational efficiency and accuracy. Besides, a comparative analysis was conducted with the temperature field distribution of the sample obtained in the hypersonic wind tunnel integration test to verify the correctness of the developed thermal-fluid-solid coupling calculation method. Finally, through thermoelastic coupling computational models, the high-thermal-gradient-induced thermomechanical damage mechanism of the CMCs was revealed, providing support for further optimizing the material system and thermal structure design of the CMCs.

High-temperature transient-induced thermomechanical damage of fiber-reinforced ceramic-matrix composites in supersonic wind tunnel

This article is based on the supersonic directly connected wind tunnel. Through a specially designed experimental chamber, combined with infrared temperature measurement, high-speed camera, etc., in-situ monitoring of composite materials under airflow at Ma 3.0 with a total temperature of 950 ∼ 1473K was carried out. The dimensional analysis method was used to propose dimensionless parameters to characterize the thermal coupling caused by high-speed airflow thermal shock. Research has shown that the thermal coupling effect of supersonic airflow causes uneven temperature inside the material, and the thermal stress caused by temperature gradient changes (including increasing and decreasing processes) is the main reason for material damage. The damage of ceramic matrix composites under thermal shock mainly manifests as a decrease in surface roughness, surface fiber fracture and a decrease in elastic modulus. In addition, the study also found that there are damage thresholds for the thermal shock effect of airflow at different total temperatures, which helps to further understand the thermomechanical damage mechanism and degradation law of composite structure under high-temperature transient conditions.

Three-dimensional characteristics of separation vortex generated by crossing shock/boundary layer interaction

Three-dimensional characteristics of separation vortex generated by crossing shock/boundary layer interaction

Relationship between shape of turbulent grid and boundary layer developed on wind tunnel wall in the supersonic wind tunnel

Relationship between shape of turbulent grid and boundary layer developed on wind tunnel wall in the supersonic wind tunnel

Design and CFD Simulation of Supersonic Nozzle by Komega turbulence model for Supersonic Wind Tunnel

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.

Experimental study on boundary layer of internal flow visible supersonic nozzle

The high-frequency pulsation noise generated by the turbulent boundary layer on the wall of a Laval nozzle can significantly affect the quality of the flow field at the nozzle outlet. In this study, a supersonic wind tunnel with visible internal flow is designed and fabricated to observe the development and evolution of the boundary layer on the contraction and expansion surfaces of a Laval nozzle, as well as to study the flow field inside the supersonic nozzle. The subsonic, transonic and supersonic profiles of the nozzle are designed by bicubic curve, Hall method and classical characteristic line method respectively. The results of numerical calculation and total pressure measurement show that the flow field at the nozzle outlet of the wind tunnel is uniform and stable, and the deviation of Mach-number-root mean square is better than the qualified level of China's national military standard. Nanoparticle-tracer based planar laser scattering (NPLS) technology is used to carry out the flow display test of the internal flow visual supersonic nozzle, and the fine structure image of the whole flow field in the nozzle is obtained. The image clearly shows the development and evolution of the boundary layer in the nozzle. The interface between boundary layer and main stream and the wall curve of nozzle transition region are extracted by image processing technology. The fractal dimension of the extracted boundary layer contour is calculated, thereby establishing the corresponding relationship between the fractal dimension and the boundary layer state, and determining the transition position of the boundary layer. The results show that the transition position of the nozzle profile is closer to downstream than that of the nozzle straight wall. The fractal dimension can qualitatively judge the flow state of the boundary layer; however, it is necessary to distinguish between laminar boundary layers and hairpin vortices in the initial transition stage by considering the thickness of boundary layer.

Visualization of sidewall vortices in rectangular nozzle supersonic blowdown wind tunnel

This study focuses on the details of the geometry and dynamics of sidewall vortices observed in supersonic wind tunnels with a rectangular cross section of the nozzle and the test section. The formation of sidewall vortices limits the accuracy of the data measured during wind tunnels' testing due to a reduced area of uniform core flow results. Most of the test data presented in this work are generated using Mie scattering visualization for M = 4 flow, with CO2 seeded up to 7% mole fraction. The Mie scattering results are complemented by data from fast pressure sensor and schlieren visualization. It is shown that the formation of vortices is caused by a transverse pressure gradient realized in the supersonic nozzle due to the gas under-expansion. The vortex external mixing layer is strongly perturbed in time but remains globally geometrically similar with streamwise distance. The vortex-generated dominant flow disturbances are in the frequency range of f = 10–50 kHz, doubling the magnitude of baseline power spectral density. The authors' viewpoint is that sidewall vortex generation is a more generic phenomenon than was thought previously.

Tehnika smanjenja opterećenja na unutrašnjim aerovagama tokom supersoničnih prelaznih pojava

Design of slender supersonic missiles requires a comprehensive experimental support in the form of wind tunnel data for a wide range of flight parameters (angle of attack and Mach number). However, in supersonic wind tunnel testing, the problem exists of transient loads at the times of the starting and stopping of the supersonic flow. In the environments of pronounced transient loads, characteristic of blowdowin wind tunnels like the T-38 of the Military technical Institute in Belgrade, it is necessary to provide control of the use of internal wind tunnel balances in the permitted design load ranges. The presented technique is related to the definition and implementation of a methodology for reducing the transient loads on wind tunnel balances in supersonic wind tunnel tests. By limiting the clearance between the model and its tail support (sting) to a magnitude which permits normal tests, but results in model-support contact during the excessive loads, part of the loads is transferred to the support sting, relieving the balance. The technique improves control over the wind tunnel test process, improves measurement accuracy and prevents damage to sensitive instrumentation (wind tunnel balances).

Simulation Test of The Aerodynamic Environment of A Missile-Borne Pulsed Laser Forward Detection System at High Flight Speed

When a missile-borne pulsed laser forward detection system flies at supersonic speed, the laser beam will be distorted by the uneven outflow field, resulting in a significant reduction in ranging accuracy. In this paper, the impact of high flight speed on a pulsed laser detection system is studied. First, a new ray tracing method with adaptive step size adjustment is proposed, which greatly improves the computational efficiency. Second, the aerodynamic environment of a munition flying at high speed is simulated by an intermittent transonic and supersonic wind tunnel to obtain the schlieren data of the flow field at various Mach numbers. The schlieren data present a shock wave structure similar to that of the simulation. In addition, the variation patterns of the pulsed laser echo waveform of the model under different aerodynamic conditions are studied to evaluate the detectability and operational stability of the laser detection system under static conditions. The test results match the simulation results well, and the two offer relatively consistent shock wave structures, which verifies the correctness and effectiveness of the flow field simulation model. The test echo waveforms are in good agreement with the simulated echo waveforms; the relative errors between the peak values of test and simulated echo waveforms at various Mach numbers do not exceed 20%, and the correlation coefficients between the test and simulated echo waveforms all exceed 0.7, indicating high correlations between the two.

Supersonic Wind Tunnel Flow Characterization Using Planar Laser CO2 Rayleigh Scattering

A complementary experimental and computational study was performed to characterize the test section flowfield of a variable-Mach-number supersonic wind tunnel. A planar laser Rayleigh scattering technique using CO2 nanocrystals was used to capture cross-sectional flow visualizations at various test section streamwise locations. Reynolds-averaged Navier–Stokes simulations of the wind tunnel flow were performed using a k-ω shear stress transport turbulence model to compare against the flow visualizations. Mach 3.20 and 3.55 flow conditions were investigated in the wind tunnel with an empty test section. The Rayleigh signal profile near the wind tunnel walls was compared against boundary-layer thickness data from shadowgraph visualizations, pitot survey measurements, and numerical simulation predictions to guide the interpretation of the scattering results. A brief discussion regarding the nature and potential source of certain unexpected scattering patterns in the images was provided. The most distinguishable flowfield characteristics observed in the Rayleigh scattering visualizations were the asymmetric thickness of the boundary layers along the test section walls and the appearance of two floor lobes near the flow-path corners. The results of the numerical simulations showed favorable agreement with the experimental data and were used to provide insight into the mechanism that caused the boundary-layer features observed. This study demonstrated the usefulness of CO2 Rayleigh scattering to obtain flow visualization data of great value to better understand flow quality in supersonic wind tunnels with two-dimensional nozzles.

Intelligent reconstruction algorithm of hydrogen-fueled scramjet combustor flow based on knowledge distillation model compression

The real-time perception of the combustion flow of a scramjet can help to quickly evaluate the working state and provide a new opportunity for the intelligent reconstruction of the data-driven combustion flow field. However, the working process of supersonic combustors is extremely short, usually of the order of milliseconds, and the inference speed is slow due to the large number of parameters of traditional deep learning flow field reconstruction models within a very short time. Currently, knowledge distillation is known as an effective model compression method. In this paper, a neural network model based on symmetric structure cascade (NNSSC1) is proposed. By means of symmetric structure cascade, features of different levels are spliced and fused, which can effectively improve the reconstruction performance of the model. Secondly, to accelerate the application of the neural network model in the flight test, a student model with a simple structure is constructed. By distillation learning, the NNSSC model is used as the teacher model, combining the high performance of the teacher model and the high efficiency of the student model, and the combustion flow field is reconstructed with high speed and high precision based on the supersonic combustor wall pressure. In a direct connected supersonic pulse combustion wind tunnel with an inflowing Mach number of 2.5, the model was trained and tested on a dataset constructed in a hydrogen fuel scramjet experiment with different equivalent ratios. This method achieves high-precision and efficient reconstruction of complex flow fields in the combustor based on sparse wall pressure, providing more reliable data for accurately evaluating the flow and combustion status of the combustor of a scramjet.

Establishment and validation of a relationship model between nozzle experiments and CFD results based on convolutional neural network

The acquisition of experimental data in a supersonic wind tunnel often faces challenges of complexity and high costs. Furthermore, there are limitations in the control of experimental conditions and measurement techniques. In this study, experiments were conducted on the start-up and shutdown processes of a single expansion ramp nozzle (SERN), and a dataset was established in conjunction with computational fluid dynamics (CFD). By utilizing experimental data and CFD simulation results, we employed a deep learning-based approach, utilizing one-dimensional convolutional neural networks (CNN), to establish a relationship model between the experimental and CFD results, with the aim of using CFD results to predict experimental outcomes. We provide detailed descriptions of the experimental methods, numerical simulation methods, and CNN model, and conduct training and testing of the model. The results demonstrate that the CNN model is capable of accurately predicting the wall pressure distribution of the SERN, with low prediction errors. In interpolation and extrapolation cases, the model exhibits good accuracy and generalization ability, with coefficients of determination above 96%. Compared to CFD calculations, the CNN model proves to be more reliable and accurate in predicting SERN performance. Additionally, we establish a comparative model that does not utilize CFD data to verify the superiority of the proposed CNN model. The results show that the CNN model with CFD data improves performance on average by 14%. This approach reduces reliance on experiments to a certain extent, lowering experimental costs and complexity. It not only saves the resources and time required for experiments but also compensates for the limitations in controlling experimental conditions and measurement techniques. Therefore, this research provides a viable solution to overcome the challenges and limitations of acquiring experimental data.