High-enthalpy Wind Tunnel Research Articles

Experimental and numerical investigation of the effects of porous bleed systems on a model scramjet inlet under high backpressure conditions

In this study, we conducted experimental and numerical investigations to assess the effects of backpressure and configurations of porous bleeds on a scramjet inlet's performance at Mach 5. Experiments in a high-enthalpy wind tunnel, using a backpressure controller and a porous bleed with 1 mm holes and 10 mm lateral spacing, were paired with Reynolds-averaged Navier-Stokes simulations to explore the mitigation of flow separation and the prevention of inlet unstart under various backpressure conditions. The simulations provided insights into how backpressure levels and porous bleed geometries impact internal flow structures, as well as the critical backpressure ratio at which inlet unstart occurs. Additional simulations varied the bleed system's hole diameter and spacing, identifying configurations that optimize aerodynamic performance by enhancing total pressure recovery and Mach number at the isolator exit, while reducing bleed mass flow rates. The study also quantified the impact of these configurations on engine thrust, determining that certain bleed system designs significantly improve thrust by efficiently suppressing the interaction between the core flow and corner recirculation zones. This study examined the internal three-dimensional flow structure characteristics under high back pressure and provided insights into porous bleed configuration capable of efficiently suppressing inlet unstart.

Sub-filter-scale shear stress analysis in hypersonic turbulent Couette flow

Direct numerical simulations of hypersonic turbulent Couette flows are performed for top-wall Mach numbers of 6, 7 and 8, inspired by non-reactive high-enthalpy wind tunnel free-stream conditions, with the goal of analysing the physical processes driving the sub-filter-scale (SFS) stresses to inform development of large-eddy simulation techniques for hypersonic wall-bounded flows. Semi-local scaling laws collapse mean profiles and second-order turbulent statistics well, in spite of the strong wall-normal gradients of temperature and density. On the other hand, the SFS shear stresses exhibit an unexpected profile characterized by a region of pronounced shear stress deficit, which becomes more pronounced for higher Mach numbers. Instantaneous visualizations suggest that the SFS shear stress deficit is induced by the counter-gradient resolved momentum transport driven by residual velocity motions at the interface between high-density low-speed streaks being ejected away from the wall, and low-density high-speed ones replacing the displaced fluid, qualifying this as a compressibility effect. It is shown that the SFS shear stresses are primarily driven by second-order interactions between residual velocities, in spite of their triply nonlinear nature. This, in turn, motivated a statistical quadrant analysis revealing the presence of a SFS shear stress deficit, that is, SFS processes driving momentum transport of the resolved field towards the top wall. Additionally, higher-Reynolds-number simulations reveal that there is an upper limit of spatial filter widths to resolve large-scale structures, and such deficit is observable at any Reynolds number when a reasonable spatial filter is applied.

Velocity measurements in particle-laden high-enthalpy flow using non-intrusive techniques

This research aims at analysing the particle-laden flow of the hypersonic high-enthalpy wind tunnel L2K, situated in Köln at the German Aerospace Center (DLR). In the L2K wind tunnel, Martian atmosphere can be created, and the facility can simulate heat load conditions encountered during atmospheric entry of Martian missions. In the tests, a simplified Martian atmosphere (97% CO2 and 3% N2) was used. The high-enthalpy flow was loaded with micrometric particles of magnesium oxide. The particles’ mean velocity was measured with a 2D–2C particle image velocimetry (PIV) system, in the region right downstream the nozzle expansion of the wind tunnel. The work proves the possibility of creating a high-enthalpy particle-laden flow for thermal protection systems (TPS) testing with simulated Martian atmosphere. Average particle velocities of around 2000 m/s are measured and compared with the numerical simulation of the wind tunnel’s particle-free flow, and with the flow velocity measured with tunable diode laser absorption spectroscopy (TDLAS). The study also highlights some unexpected results and features of the high-enthalpy particle-laden flow and proposes some theories for the causes of such effects, which include agglomeration due to melting, and gravitational effect.

UV spectroscopy of artificial meteors (200–400 nm)

The high energy of meteoroid entering the Earth atmosphere presumably results in UV radiation. However, ground-based observations are impaired by the atmospheric absorption below 400 nm. Artificial meteors are produced in a high enthalpy wind tunnel, and observed with a [200–400] nm fiber-fed spectrometer in order to analyze for the first time the UV emission of meteors. Similarly to visible observations, several atomic lines of Fe and Mg are detected. Contrary to observations in the visible wavelength range, Si is also clearly detected in all tested samples. Carbon is not detected in atomic lines. As the strongest emission lines are detected between 220 and 330 nm, we recommend that future meteor dedicated space-based UV instruments focus on this particular wavelength interval.

Magnetic field control of high-enthalpy shock wave/boundary-layer interactions using a fully implicit thermochemical non-equilibrium solver

Magnetohydrodynamic (MHD) heat shield systems have been proven to be highly effective in hypersonic blunt-nosed bodies. Thus, it is natural to consider the application of MHD to more complex geometries that induce shock wave/boundary-layer interactions associated with locally high heat load, which is a lack of deep research, especially for the large-size model in high-enthalpy conditions. To investigate the possibility and mechanism of the control of shock wave/boundary-layer interactions using local magnetic fields, the partially ionized flow of the hollow cylinder/flare coupling of these fields was numerically analyzed using a hypersonic thermochemical nonequilibrium solver. It consisted of an 11-species reaction model and a Park two-temperature model based on the low magneto-Reynolds assumption. To address the numerical stiffness introduced by the large difference in the characteristic time of the thermochemical non-equilibrium and grid refinement in the boundary layer and interaction zone, a fully implicit block lower–upper symmetric Gauss–Seidel algorithm was developed. This algorithm improved the accelerating rate and the solver was validated using cases involving MHD control of flow around a typical reentry vehicle, and the measured double cone and hollow cylinder/flare data for a high-enthalpy wind tunnel. After that, a parametric study of high-enthalpy flow over the hollow cylinder/flare model was conducted and MHD control was divided into three categories. Among these, N-type control led to the largest reduction in the peak value of the total wall heat flux near the neck region, which was obtained using a uniform magnetic field. Specifically, the peak heat flux and the peak skin friction coefficient were reduced by 20.0% and 48.0%, respectively. Then, two local MHD interaction parameters were newly proposed for mechanism analysis and can serve as an effective indicator for MHD controllability. The numerical results demonstrated that the Lorentz force in the counter-streamlined direction induced by the external magnetic field and ionized flow were key factors that influenced the control of the local heat flux. The results of this study lay the foundation for the design of MHD experiments in a high-enthalpy wind tunnel.

Catalytic wall effects for hypersonic nozzle flow in thermochemical non-equilibrium

This paper investigates the effects of wall catalytic activity in hypersonic nozzle flows. Non-equilibrium Navier–Stokes-Fourier equations were here applied within a density-based algorithm in the OpenFOAM framework. The numerical model was validated using different test conditions conducted at NASA EAST facility and General Electric Research Laboratory. After confirming the reliability of the numerical model, a numerical investigation on the wall catalytic effects on the flow parameters using the Scirocco Plasma Wind Tunnel was performed. Four nozzles with different expansion ratios were extensively tested. It was verified that higher expansion ratio nozzles experience a greater wall catalytic effect on the flow properties. Particularly, doubling the expansion ratio causes the difference between a non-catalytic and a fully catalytic wall to increase in 21% for Mach number and 25% for trans-rotational temperature and static pressure. Furthermore, the results of this investigation allow us to define a new criterion by which to consider if the wall catalytic effects are or not important in the nozzle flow fields. It was found that catalytic effects in the core flow are negligible when simulating nozzles with expansion ratio below 186 and Mach number below 8.6. As the expansion ratio increases, the thermal boundary layer thickness also increases and the atoms recombination that occur in that region have more influence on the core flow. In these cases, it is then necessary to take into account the wall catalytic activity to achieve a better estimation of the flow properties at the nozzle exit and in the test cell. • Catalytic wall effects in high enthalpy wind tunnels are numerically investigated. • Numerical simulations of hypersonic nozzles in thermochemical non-equilibrium are performed. • The recombination reactions at fully catalytic walls changes the nozzle core flow. • Catalytic wall effects are negligible for expansion ratio below 186 and Mach number below 8.6.

Experimental and Numerical Study of Hypersonic Flow over Backward-Facing Step

Experimental and numerical studies were conducted to investigate the flow over a backward-facing step with finite aspect ratio. Measurements using pressure-sensitive paint, infrared thermal imaging, oil-flow visualization and schlieren imaging were conducted in the Hypersonic and High Enthalpy Wind Tunnel at the University of Tokyo. The results show the complex three-dimensional structure of the flow over a backward-facing step. The numerical simulation was also conducted, and the simulation results reveal that the fluid flow from the open side of the model has significant three-dimensional effect on the flow features. Comparison between CFD and experimental results is also made in this paper. A recirculation bubble found after the step, generated by the interaction of side expansion and recirculating flow, is observed in CFD results. A symmetric local peak pressure and heating region named ‘twin peak’ is also founded by both numerical simulation and experiments.

Measurement of the gas flow stagnation temperature by the method of two identical thermocouples in the short-duration aerodynamic facilities

Investigations of the temperature and heat flux distributions over the external surface of flying vehicles are important for the development of hypersonic planes. For obtaining correct results, experiments are performed in short-duration high-enthalpy wind tunnels. Determination of the stagnation temperature of the flow in such facilities is a challenging task due to short start-up times and high flow temperature. In this paper a new method is proposed for determining the stagnation temperature of a gas flow in the short-duration aerodynamic wind tunnels with stepwise changes in its parameters. The method is based on using two identical thermocouples with different initial temperatures. It is shown that in the case of jump-like changes in flow parameters the instrument function of the difference of the thermocouple readings coincides with the instrument function of each thermocouple, and the instrument function itself is actually the normalized temperature difference. The instrument function obtained in the study allows temperature retrieval by the method of deconvolution. This approach is highly sensitive to random noise of experimental data; therefore, regularization methods are applied for high-frequency noise suppression. A procedure for choosing the regularization parameter for temperature reconstruction is proposed. The research novelty of the article is that the proposed method allows exact reproduction of the stepwise function and instrument function without any additional assumptions, while the measurements and dynamic calibration are performed in one experiment. Applicability of the method for obtaining the stagnation temperature in the short-duration high-enthrall wind tunnel IT-302 M ITAM SB RUS is demonstrated.

Instability and mode transition analysis of a hydrogen-rich combustion in a model afterburner

Combustion instabilities were investigated experimentally for a hydrogen-rich combustion in a model afterburner installed at the end of a high-enthalpy wind tunnel. Air was supplied at 0.3 MPa and 950 K. The combustion instabilities were studied with the time-resolved measurements of a near-infrared (NIR) emission from water molecules over 780 nm using a high-speed video camera. Pressure was also measured in the combustor. The pressure and the NIR images were analyzed by data-driven approach, which include the fast Fourier transform (FFT), the wavelet transform, the dynamic mode decomposition (DMD) and the Gaussian process latent variable methods (GP-LVM). Thermoacoustic instability was observed under a rich condition, and the amplitude of the pressure oscillation was the maximum at the overall equivalence ratio of approximately 2.4 or 2.7 as a result of the FFT. The combustion dynamics were investigated in detail for an experimental run at the equivalence ratio of 2.4. A pressure spectrogram indicated a flame–vortex interaction with a Strouhal number of 0.5 (2300 Hz), thermoacoustic instability (560 Hz), and their transitions with the wavelet transform. For NIR images, the same tendency was also observed in the spectrogram of the modes obtained by the Gabor-filtered DMD, which could clearly resolve the high-order harmonic modes of the flame–vortex interaction and the thermoacoustic instability. Furthermore, NIR images were analyzed with GP-LVM to study the evolution of the combustion dynamics in a three-dimensional latent space. Recurrence plots with the Euclidean distance function were used to visualize the evolutions of the combustion dynamics. A limit cycle behavior of the flame–vortex interaction was clearly observed, whereas the limit cycle of the thermoacoustic instability showed more complicated behaviors. The transition behaviors of the instabilities were observed in the recurrence plots in detail, indicating that the flame–vortex interaction excited the fourth harmonic mode of the thermoacoustic instability, followed by the basic mode.

Temporally and Spatially Resolved Pressure and Temperature Maps in Hypersonic Flow

Anodized aluminum pressure sensitive (AAPSP) and temperature sensitive paints (AATSP) were applied to a 30° compression corner model in the Mach 7.1 Hypersonic and High Enthalpy Wind Tunnel at the University of Tokyo, Kashiwa Campus. The measurement technique was able to capture spatially resolved unsteady hypersonic flow phenomena over the surface of the model at a rate of 10 Hz for pressure measurements and 30Hz for temperature measurements. High heating due to oblique shocks and flow stagnation was observed at the leading edge of the model. Shock-shock interactions, confirmed by schlieren imagery were clearly visible in the pressure distribution. Görtler vortices consistent with boundary layer growth over a concave wall were also visible in both the pressure and temperature maps produced by the AAPSP and AATSP. The tests demonstrate the effective use of AAPSP and TSP for spatially resolved measurements in hypersonic flow. The temporal resolution of the measurement was limited by camera technology, and can be improved without modifications to the luminescent sensors.

Improved representation of destructive spacecraft re-entry from analysis of high enthalpy wind tunnel tests of spacecraft and equipment

The assessment of casualty risk in destructive re-entry has historically been performed purely by simulation using heating correlations, which only have verification on basic shapes, and estimated phenomenology for fragmentation. As the application of space debris mitigation requirements is expected to result in a higher number of re-entries, this has brought a stricter enforcement of the casualty risk guidelines to proposed missions. In turn, this has increased the interest in designing spacecraft to demise in re-entry in order to allow uncontrolled re-entries to be performed. Initial testing of spacecraft materials and basic structures has demonstrated that the destructive re-entry tools do not capture the correct physics to be able to assist design in a meaningful way, and therefore, some means to improve the representativeness of the tools is required. The current understanding of the phenomenology of the fragmentation and demise processes is limited. As a consequence, it is vital to perform appropriate tests in order to improve the capability of the tools to assist the design process. To this end, a set of destructive tests on spacecraft materials, structures and components has been performed in an arc-heated supersonic wind tunnel. These tests include the first destructive wind tunnel tests ever performed on a complete nano-satellite and a reaction wheel. From these tests, it has been determined that the failure of aluminium structures is highly dependent upon the behaviour of the protective metal oxide layer, and that this can be catastrophic in nature. The tests on the nano-satellite have shown that the structure can be supported by stainless steel spacers between the electronics cards, and that glass fibre reinforced plastic PCBs are more resistant to melting than had been anticipated. The reaction wheel test has shown that the connections between parts are critical to the fragmentation and demise processes, as the glued housing separates quickly, well before melt temperature is reached at the joint. It has also demonstrated the importance of radiative cooling, as the flywheel and ball-bearing unit have survived a test at over 800 kW/m2 with little damage.

Numerical simulation of a CO2, CO, O2, O, C mixture: Validation through comparisons with results obtained in a ground-based facility and thermochemical effects

Various issues of numerical simulation of high-enthalpy non-equilibrium flows of a gas mixture of CO2, CO, O2, O, and C are discussed. Navier–Stokes equations are considered within the framework of one-, two-, and three-temperature approximations with allowance for chemical reactions. In the one-temperature approximation, chemical reactions are modelled by the Arrhenius law. In the multi-temperature approximations, simulations are performed with the use of the Treanor–Marrone model for carbon dioxide dissociation and the Park model for dissociation of diatomic gas molecules, which take into account vibrational non-equilibrium of molecules. The Arrhenius law is also used for calculating recombination reactions and also forward and backward exchange reactions. Translational-vibrational and vibrational-vibrational energy exchange is simulated by a simplified Landau–Teller model. Some examples of flows with the use of the above-mentioned approximations are calculated. It is demonstrated that the one-temperature approach ensures an adequate reproduction of experimental data on the shock standoff distance, heat flux, and pressure coefficient on the cone, which are obtained in a high-enthalpy wind tunnel. The calculation of the flow around a cylinder under Martian re-entry conditions reveals a significant effect of thermochemical non-equilibrium.

Aerodynamic Design of a Laval Nozzle for Real Gas Using Hodograph Method

We propose an approach for the design of the subsonic part of plane and axisymmetric Laval nozzles for real gases. The proposed approach is based on the hodograph method and allows one to solve the inverse design problem directly. Real gas effects are taken into consideration using the chemical equilibrium model. We present nozzle contours computed with the proposed method for a stoichiometric methane-air mixture. Results confirm that real gas effects have a strong influence on the nozzle shape. The described method can be used in the design of nozzles for rocket engines and for high-enthalpy wind tunnels.

Numerical inverse method for calculating heat flux in temperature-sensitive-coating measurement on a finite base

The one-dimensional unsteady heat transfer of a polymer layer on a finite base is studied. A numerical inverse method is developed to calculate heat flux from surface-temperature images obtained using temperature-sensitive coating particularly in high-enthalpy wind tunnels. It is found that the effect of the finite thickness of the base is significant as time increases such that the semi-infinite base assumption made in many methods is no longer accurate in heat-flux calculation. The method developed in this work can take into account the effects of the finite thickness of a base and the temperature-dependencies of the thermal properties of the materials. This method has been validated through simulations and processing the experimental temperature-sensitive-paint (TSP) images.

東京大学柏キャンパス極超音速高エンタルピー風洞( 空力実験設備:現状と新展開 第4回)

東京大学柏キャンパス極超音速高エンタルピー風洞( 空力実験設備:現状と新展開 第4回)

Evolution of disturbances in the shock layer on a flat plate in the flow of a mixture of vibrationally excited gases

The results of the numerical and experimental investigations of the evolution of the disturbances in a hypersonic shock layer on a flat plate streamlined by a flow of the mixture of vibrationally excited gases are presented. The experimental study was conducted in the hot-shot high-enthalpy wind tunnel IT-302 of the ITAM SB RAS. The numerical simulation was carried out with the aid of the ANSYS Fluent package using the solution of the unsteady two-dimensional Navier−Stokes equations with the incorporation of the user-created modules and enabling the consideration of the vibrational non-equilibrium of the carbon dioxide molecules within the framework of the model of the two-temperature aerodynamics. It was obtained that an increase in the carbon dioxide concentration in the mixture with air leads to a reduction of the intensity of pressure disturbances on the surface. The efficiency (up to 20 %) of the method of sound absorbing coatings in the vibrationally excited flows of the mixture of the carbon dioxide and air has been shown.

Numerical Study of a Three Dimensional Interaction between two bow Shock Waves and the Aerodynamic Heating on a Wedge Shaped Nose Cone

This paper presents a numerical investigation on the three-dimensional interaction between two bow shock waves in two environments, i.e. ground high-enthalpy wind tunnel test and real space flight, using Fluent 15.0. The first bow shock wave, also called induced shock wave, which is generated by the leading edge of a hypersonic vehicle. The other bow shock wave can be deemed objective shock wave, which is generated by the cowl clip of hypersonic inlet, and in this paper the inlet is represented by a wedge shaped nose cone. The interaction performances including flow field structures, aerodynamic pressure and heating are analyzed and compared between the ground test and the real space flight. Through the analysis and comparison, we can find the following important phenomena: 1) Three-dimensional complicated flow structures appear in both cases, but only in the real space flight condition, a local two-dimensional type IV interaction appears; 2) The heat flux and pressure in the interaction region are much larger than those in the no-interaction region in both cases, but the peak values of the heat flux and pressure in real space flight are smaller than those in ground test. 3) The interaction region on the objective surface are different in the two cases, and there is a peak value displacement of 3 mm along the stagnation line.

Spatial variation in similarity spectra decompositions of a Mach 1.8 laboratory-scale jet

The primary source of noise from supersonic jets is turbulent mixing noise. Tam et al. [AIAA Paper 96-1716 (1996)] proposed similarity spectra for a two-source model of turbulent mixing noise corresponding to noise from omnidirectional fine-scale turbulence structures and directional large-scale turbulent structures. These empirical similarity spectra have been compared with reasonable success to the spectra of both military and laboratory-scale jets. Most application have looked at the variation in angle: fine-scale similarity spectra matches sideline of the jet nozzle, large-scale in the maximum radiation lobe, and a combination needed in between. A similarity spectra decomposition of from an ideally expanded, Mach 1.8 laboratory-scale jet allows for a spatial comparison between the near and far-field spectra. The sound from the convergent-divergent nozzle was collected at the Hypersonic and High-Enthalpy Wind Tunnel at Kashiwa Campus of the University of Tokyo at a variety of near, mid, and far field locations. Comparison of similarity spectra decompositions over distance yield insights into the sound propagation from this supersonic jet. In addition, results from a preliminary wavenumber eduction technique to build a wavepacket-based equivalent source model of the large-scale turbulent mixing noise. [Work supported by Office of Naval Research grant.]

Effect of the mode composition disturbances in a high-enthalpy wind tunnel on wave processes in the hypersonic viscous shock layer

A numerical and experimental study of the spectral distribution of disturbances hotshot wind tunnel was performed. It was found that disturbances are the superposition of fast acoustic waves propagating along axial direction and oblique slow acoustic waves which have less amplitude. Numerical simulation of disturbances evolution in the nozzle of the wind tunnel was carried out by solving 2D Navier-Stokes equations using the ANSYS Fluent program package. Experiments were carried out in the IT-302M high-enthalpy hotshot wind tunnel ITAM SB RAS.

Unified approach for conjugate heat-transfer analysis of high speed air flow through a water-cooled nozzle

ABSTRACTThis article presents a unified approach to solve steady-state conjugate heat-transfer problem including simultaneously gas, liquid and solid regions in just one 3D domain, distinguished by their particular properties. This approach reduces approximation errors and the time to solve the problem, which characterise iterative methods based on separated domains. The formulation employs RANS equations, realisablek-ε turbulence model and near-wall treatment model. A commercial CFD code solves the pressure-based segregated algorithm combined with spatial discretisation of second order upwind. The problem consists of a convergent-divergent metallic nozzle that contains cooling channels divided in two segments along the wall. The nozzle wall insulates the high-speed hot air flow, dealt as perfect gas, from the two low-speed cold water flows, dealt as compressed liquid, both influenced by transport properties dependent of the local temperature. The verification process uses three meshes with increasing resolutions to demonstrate the independence of the results. The validation process compares the simulation results with experimental data obtained in high-enthalpy wind tunnel, demonstrating good compliance between them. Results for the bulk temperature rise of the water in the second cooling segment of the nozzle showed good agreement with available experimental data. Numerical simulations also provided wall temperature and heat flux for the gas and liquid sides. Besides, distribution of temperature, pressure, density and Mach number were plotted along the nozzle centerline showing a little disturbance downstream the throat. This phenomenon has been better visualised by means of 2D maps of those variables. The analysis of results indicates that the unified approach herein presented can make easier the task of simulating the conjugate convection-conduction heat-transfer in a class of problems related to regeneratively cooled thrust chambers.