Hypersonic Flow Conditions Research Articles

Improving heat transfer prediction across catalytic SiC interface through a multiscale framework leveraging machine learning and data fusion techniques

Accurate aerothermal prediction under hypersonic flow conditions is crucial for any thermal protection materials design and engineering. The surface catalytic effect, where dissociated atoms recombine into their molecular forms at the air-solid interface, is an exothermic reaction and plays an important role in high-enthalpy aerodynamic environment prediction. Taking SiC, a widely recognized high-temperature thermal protection material as an example, a multiscale fusion approach for aerothermal prediction is proposed. The approach utilizes reactive molecular dynamics method to construct an interface catalysis model, and the Bayesian maximum entropy to integrate experimental and simulational data into an optimized database. Subsequently, using the radial basis function neural network algorithm, a machine learning-based reaction kinetics model with precise analysis of surface catalysis is trained. The obtained catalytic recombination efficiency is used as a boundary input for computational fluid dynamics simulation, enabling rapid and accurate prediction of hypersonic thermal environment. The result shows that the predicted surface heat flux by this novel approach is consistent with the benchmark studies, but comes with significantly reduced computation time. The stagantion heat flux predictions based on the assumptions of full catalytic wall, finite rate catalytic wall (from the proposed multiscale fusion method), and non-catalytic wall are determined to be 8.65×106 W/m2, 7.51×106 W/m2, and 4.02×106 W/m2, respectively. Comparing these with the wind tunnel benchmark value of 7.48×106 W/m2, the error from the proposed multiscale fusion method is 0.30 %, indicating significant enhancement in prediction accuracy through the proposed upscaling method. The new approach could not only reveal complex exothermic reaction processes at the interface, but also enhance the prediction accuracy at much higher computational efficiency, providing an alternative for multiscale modeling of complex flow and heat transfer at the interface.

Aerodynamic Drag Coefficient Prediction of a Spike Blunt Body Based on K-Nearest Neighbors

Spike blunt bodies are a method to reduce drag when a body moves at speeds above sound. Several numerical works based on computational fluid dynamics (CFD) have deeply studied fluid performance and highlighted its advantages. However, most documentation focuses on modifying spike physical properties while keeping constant supersonic or hypersonic flow conditions. In recent years, machine learning models have emerged as viable tools to predict values in almost any field, including aerodynamics. In the case of CFD, many models have been explored, such as support vector regression, ensemble methods, and artificial neural networks. However, a simple and easy-to-implement method such as k-Nearest Neighbors has not been extensively explored. This work extrapoled k-Nearest Neighbors to predict the drag coefficient of a spike blunt body for a range of supersonic and hypersonic speeds based on drag data obtained from CFD analysis. The parametric study of the spike blunt body was performed considering body diameter, spike length, and freestream Mach number as input variables. The algorithm presents proper predictions, with errors less than 5% for the drag coefficient and considering a minimum of three neighbor nodes. The k-NN was compared again Kriging model and k-NN presents a better accuracy. The above validates the flexibility of the method and shows a new area of opportunity for the calculation of aerodynamic properties.

Surrogate multi-objective optimization of missile RCS performance in hypersonic rarefied flow of the near space

This study focuses on optimizing the lateral jet efficiency of THAAD-like (Terminal High Altitude Area Defense) missiles operating under hypersonic rarefied flow conditions. We employ the DSMC-QK algorithm to simulate the three-dimensional lateral jet flow field, accounting for thermochemical non-equilibrium effects. The analysis investigates how the force/momentum amplification coefficient varies with the angle of attack, jet pressure ratio, jet Mach number, and jet gas composition. Subsequently, we develop an artificial neural network (ANN) proxy model using the pyrenn toolbox, achieving an average prediction error of 0.866% and a maximum error of 1.60%. Utilizing this ANN model, we perform single- and multi-objective optimizations with a genetic algorithm to determine the optimal jet parameters. The results reveal that in multi-objective optimization, the proportion of helium in the jet gas composition increases, leading to a slight reduction in the force amplification coefficient but a substantial 61.4% decrease in the mass flow rate. This demonstrates that a judicious selection of jet gas composition can significantly reduce mass flow while maintaining high jet efficiency, thus achieving efficient lateral jet control.

Reduced-Order Modeling of Steady and Unsteady Flows with Deep Neural Networks

Large-eddy and direct numerical simulations generate vast data sets that are challenging to interpret, even for simple geometries at low Reynolds numbers. This has increased the importance of automatic methods for extracting significant features to understand physical phenomena. Traditional techniques like the proper orthogonal decomposition (POD) have been widely used for this purpose. However, recent advancements in computational power have allowed for the development of data-driven modal reduction approaches. This paper discusses four applications of deep neural networks for aerodynamic applications, including a convolutional neural network autoencoder, to analyze unsteady flow fields around a circular cylinder at Re = 100 and a supersonic boundary layer with Tollmien–Schlichting waves. The autoencoder results are comparable to those obtained with POD and spectral POD. Additionally, it is demonstrated that the autoencoder can compress steady hypersonic boundary-layer profiles into a low-dimensional vector space that is spanned by the pressure gradient and wall-temperature ratio. This paper also proposes a convolutional neural network model to estimate velocity and temperature profiles across different hypersonic flow conditions.

Low-frequency large-scale unsteadiness analysis of lateral jet interactions on a slender cylinder in hypersonic flows using fast-responding pressure-sensitive paint

Lateral jet interactions, commonly encountered in high-speed vehicles, are typical complex three-dimensional (3D) shock wave/boundary layer interactions. The spatiotemporal features of surface pressure are crucial for understanding such complex interactions. In this study, fast-responding pressure-sensitive paint measurements are implemented at a sampling rate of 3000 Hz and a spatial resolution of 0.35 mm/pixel to investigate global pressure fields in lateral jet interactions on a slender cylinder under hypersonic flow conditions. Tests are conducted at three angles of attack (0°, 5°, and 10°) and two sideslip angles (0° and 2.5°) at Mach 6 and a Reynolds number of 1.35 × 107/m. The mean pressure results indicate the presence of a 3D high-pressure region beneath the bow shock feet, the area and pressure level of which depend on the cylinder attitude. Moreover, significant low-frequency broadband unsteadiness associated with the large-scale motion of bow shock is found in the high-pressure region. Statistical, coherence, and modal analyses are performed to clarify the characteristics of the large-scale unsteadiness. The results indicate that the dominant pressure mode is induced by streamwise oscillation of the bow shock, which exhibits two subregions with opposite pressure-fluctuation variations. Furthermore, the effects of the cylinder attitude on the unsteadiness extent and intensity are discussed.

A hybrid CFD-RMD multiscale coupling framework for interfacial heat and mass simulation under hyperthermal ablative conditions

Under hypersonic flow conditions, the heat and mass transfer at the gas-solid interface of thermal protection material becomes highly complicated due to the strong chemical and thermal non-equilibrium phenomena, associated with the real gas effect, the high temperature effect and various heterogeneous reactions occurring at the interface or inside the TPM. Interfacial heat and mass transfer is strongly coupled with non-linear reactions at different spatiotemporal scales. Based on the traditional macroscopic Computational Fluid Dynamics (CFD) and microscopic Reactive Molecular Dynamics (RMD) methods, a hybrid CFD-RMD multiscale simulation framework is established in this work to improve the accuracy of aerothermal prediction by considering the heat and mass transfer at the gas-solid interface under hyperthermal non-equilibrium conditions. To validate the proposed multiscale framework, phenolic resin composite is investigated in this work with particular interest in the pyrolysis gas injection effects. Benchmarking ablation studies are conducted under three representative altitudes of 40, 60 and 80 km, respectively, and the recession rates of 7.47 × 10−4, 1.53 × 10−3 and 1.08 × 10−3 mm/s are obtained which are comparable with the experimental results. Much higher computational efficiency of the hybrid CFD-RMD multiscale framework is achieved comparing to a full RMD simulation method at the same spatial and temporal level. Not only to the pyrolysis study, the established framework can also potentially be extended to other gas-solid interfacial reaction situations, which is of great help in improving the mechanism understanding of various physics-derived microscopic models for macroscopic predication of hypersonic aerothermal characteristics.

Implementation of various-fidelity methods for viscous effects modeling on the design of a waverider

In the present study, the implementation of three viscous effects modeling methods at hypersonic flow conditions is investigated. The design of a hypersonic passenger transport waverider configuration is employed as a case study. The methods, each of which has a different fidelity level, are evaluated in terms of their capability to predict aerodynamic performance, as well as their effect on the final geometric design of the waverider. More specifically, for the same mission profile and design constraints, three different waverider configurations (Cases A, B and C) are designed with the osculating cone methodology, and each employing a different viscous effect modeling method. These three configurations are designed (a) assuming a purely inviscid flowfield, (b) using Eckert's Reference Temperature and (c) Van Driest's viscous effects modeling semi-empirical methods respectively, each subsequent one of higher complexity and fidelity than the previous. An optimization loop, employing the Nelder-Mead simplex algorithm, is also incorporated into the overall design process. Each configuration is evaluated with all three viscous effects modeling methods and the analytical results are compared against those of CFD modeling. The CFD analyses are performed by solving the steady-state three-dimensional Euler and Reynolds Averaged Navier Stokes (RANS) equations. Overall, the three configurations are evaluated in terms of aerodynamic performance, design/optimization computational cost, and accuracy between the analytical and CFD results. All inviscid analytical calculations demonstrate very good agreement with the Euler CFD results. Indicatively, the lift-to-drag predictions deviate by as much as 2% between the two approached. The Van Driest's method predictions result to 5%-6.5% lower lift-to-drag ratios compared to the RANS CFD, and to higher wall shear stress values. Contrarily, the Ref. Temperature method predictions are in a much closer agreement to the RANS CFD results, leading to a slight overprediction (0%-2.4%) of lift-to-drag ratios. The incorporation of viscous effects in the design process reduces the waverider's aerodynamic efficiency by about 8.5%-11% and results in a 46%-98% increase in computational time, compared to the inviscid analysis.

Aerodynamic Shape Optimization of a Symmetric Airfoil from Subsonic to Hypersonic Flight Regimes

Hypersonic flight has been the subject of numerous research studies during the last eight decades. This work aims to optimize the aerodynamic performance of a two-dimensional baseline airfoil (NACA0012) at distinct flight regimes from subsonic to hypersonic speeds. A mission profile has been defined, where four points representing the subsonic, transonic, supersonic, and hypersonic flow conditions have been selected. A framework has been implemented based on high-fidelity RANS computational fluid dynamics simulations. Gradient-based optimizations have been conducted with the objective of minimizing the drag. The optimization results show an overall improvement in aerodynamic performance, including a decrease in the drag coefficient of up to 79.2% when compared to the baseline airfoil. In the end, a morphing strategy has been laid out based on the optimal shapes produced by the optimization.

Simulations of Hypersonic Boundary-Layer Transition over a Flat Plate with the Spalart-Allmaras One-Equation BCM Transitional Model

Transitional flow has a significant impact on vehicles operating at supersonic and hypersonic speeds. An economic way to simulate this problem is to use computational fluid dynamics (CFD) codes. However, not all CFD codes can solve transitional flows. This paper examines the ability of the Spalart–Allmaras one-equation BCM (SA-BCM) transitional model to solve hypersonic transitional flow, implemented in the open-source CFD code Eilmer. Its performance is validated via existing wind tunnel data. Eight different hypersonic flow conditions are applied. A flat plate model is built for the numerical tests. The results indicate that the existing SA-BCM model is sensitive to the freestream turbulence intensity and the grid size. It is not accurate in all the test cases, though the transitional length can be matched by tuning the freestream intensity. This is likely due to the intermittency term of the SA-BCM model not being appropriately calibrated for high-velocity flow, though if the model can be recalibrated it may be able to solve the general high-velocity flows. Although the current SA-BCM model is only accurate under certain flow conditions after one calibration process, it remains attractive to CFD applications. As a one-equation model, the SA-BCM model runs much faster than multiple-equation flow models.

Surrogate-based aerodynamic shape optimization of a morphing wing considering a wide Mach-number range

Aerodynamic shape design considering both subsonic and hypersonic performance is a challenge in the field of aerospace. Many studies have shown that morphing wing is a promising technology to address this challenge. In this article, a new morphing mechanism which changes both the plane shape (such as span and sweep angle) and the profile (such as chord length and relative thickness) of the wing is proposed and a multi-objective optimization design for this new morphing mechanism is conducted based on surrogate-based optimization (SBO) algorithm. In the optimization design, three different configurations of the morphing wing which are the 36-, 45- and 70-degree sweep angle state, respectively and four different flow conditions which are subsonic flow (0.25 Ma), transonic flow (0.85 Ma), supersonic flow (1.2 Ma) and hypersonic flow (6 Ma) are considered. 36-degree sweep angle state is used for subsonic flow, 45-degree sweep angle state is used for transonic flow and supersonic flow and 70-degree sweep angle state is used for hypersonic flow. Subsonic lift and transonic, supersonic, hypersonic lift-to-drag ratios are objectives of this optimization design which has up to 64 design variables. To reduce computational cost of high-fidelity CFD (computational fluid dynamics) simulations and find the global optimum, Kriging surrogate model combined with a parallel infill-sampling method and a multi-round strategy are employed in this study. At last, the optimized morphing wing is evaluated and compared with the baseline and a fixed wing which is a moderate sweep wing. Results show that the optimized morphing wing features a very good aerodynamic performance improvement over the flow regimes from subsonic to hypersonic flow conditions.

Material Response Modeling of Melt Flow-Vapor Ablation for Iron

The ability to accurately predict melt-dominated ablation of meteoritic materials is important for risk assessment of near-Earth asteroids and survivability of reentering space debris. Despite previous computational studies that involve melt ablation for glassy materials, the application of incompressible gas boundary-layer theory to melt/vapor formation has largely been applied to limited surface heating rates and vapor-dominated regimes. A melt ablation formulation near stagnation regions, for crystalline and glassy materials, is implemented here and validated with ground experiments performed for iron under hypersonic flow conditions. The mathematical model does not neglect temperature gradients in the molten layer or assume domination of either melt or vapor formation. The implementation of this comprehensive melt/vapor ablation numerical model helps provide domain-independent engineering recession estimates. By using computational fluid dynamic-driven surface properties and radial basis function interpolation-driven dynamic meshes to predict stagnation-point melt ablation, it is shown that melt initiation is predicted within 10% of pyrometer data and that simulated recession is dominated by melt flow contributions at the extreme operating limits of a hypersonic arc-jet test facility. The melt thickness shows a strong parabolic dependence on heat transfer coefficient variation. Shape change validation results of test articles accounting for this variation are discussed.

Competing effects of surface catalysis and ablation in hypersonic reentry aerothermodynamic environment

Under hypersonic flow conditions, the complicated gas-graphene interactions including surface catalysis and surface ablation would occur concurrently and intervene together with the thermodynamic response induced by spacecraft reentry. In this work, the competing effects of surface heterogeneous catalytic recombination and ablation characteristics at elevated temperatures are investigated using the Reactive Molecular Dynamics (RMD) simulation method. A Gas-Surface Interaction (GSI) model is established to simulate the collisions of hyper-enthalpy atomic oxygen on graphene films in the temperature range of 500–2500 K. A critical temperature Tc around 900 K is identified to distinguish the graphene responses into two parts: at T < Tc, the heterogeneous surface catalysis dominates, while the surface ablation plays a leading role at T > Tc. Contradicting to the traditional Arrhenius expression that the recombination coefficient increases with the increase of surface temperature, the value is found to be relatively uniform at T < Tc but declines sharply as the surface temperature increases further due to the competing ablation effect. The occurrence of surface ablation decreases the amounts of active sites on the graphene surface for oxygen adsorption, leading to reduced recombination coefficient from both Langmuir-Hinshelwood (L-H) and Eley-Rideal (E-R) mechanisms. It suggests that the traditional Computational Fluid Dynamics (CFD) simulation method, which relies on the Arrhenius-type catalysis model, would result in large discrepancies in predicting aerodynamic heat for carbon-based materials during reentry into strong aerodynamic thermal environment.

Reduced order models for heat flux and pressure distributions on space debris afterbodies

The prediction of the on-ground risk caused by re-entering space debris requires accurate and computationally affordable models to compute wall heat flux and pressure coefficient Cp at every stage of the reentry. Such models already exist for the wind area of the debris, i.e the area directly impinged upon by fictitious lines parallel to the incoming flow. But the heat flux and Cp are often neglected in the shadow area, even though they can be relatively high for certain shapes of debris. The present work focuses on developing reduced order models for heat flux and Cp distributions in the shadow area of space debris, for hypersonic continuous flow conditions. We identified four phenomena which appear in the shadow area and cause relatively high levels of these aerodynamic quantities (attached flow, detached flow with fluid reattachment, detached flow with solid reattachment and shock–shock interactions). Models were developed for the heat flux and Cp distributions caused by attached flows on the lee side of cylindrical geometries using the Proper Orthogonal Decomposition (POD) and interpolation method. Using this method, it is possible to reduce the number of required data by efficiently exploring the parameters domain of variation. The sample points were chosen thanks to an adaptive design of experiments, and the input data for the models were obtained by 3D Navier–Stokes computations of the flow around cylinders at incidence. The analysis of the computational results highlighted the influence of the Reynolds number based on the cylinder diameter ReD,∞ on the heat flux and pressure distributions. Reduced-order models were created for three input parameters (length L, diameter D and angle of attack α) and fixed incoming flow conditions, using the POD and interpolation method, and then were extended to other upstream hypersonic continuous flow conditions by more classical approaches based on non-dimensional quantities. Finally, the reentry trajectories of two different cylindrical debris were computed with and without the new models developed, to demonstrate their influence on the integrated heat flux received by the debris during their reentry, and therefore on their survival rate.

Criteria for hypersonic airbreathing propulsion and its experimental verification

Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years' development. Two critical issues are identified to be the barriers for the ramjet-based engine that has been taken as the most potential concept of the hypersonic propulsion for decades. One issue is the upstream-traveling shock wave that develops from spontaneous waves resulting from continuous heat releases in combustors and can induce unsteady combustion that may lead to engine surging during scramjet engine operation. The other is the scramjet combustion mode that cannot satisfy thrust needs of hypersonic vehicles since its thermos-efficiency decreases as the flight Mach number increases. The two criteria are proposed for the ramjet-based hypersonic propulsion to identify combustion modes and avoid thermal choking. A standing oblique detonation ramjet (Sodramjet) engine concept is proposed based on the criteria by replacing diffusive combustion with an oblique detonation that is a unique pressure-gain phenomenon in nature. The Sodramjet engine model is developed with several flow control techniques, and tested successfully with the hypersonic flight-duplicated shock tunnel. The experimental data show that the Sodramjet engine model works steadily, and an oblique detonation can be made stationary in the engine combustor and is controllable. This research demonstrates the Sodramjet engine is a promising concept and can be operated stably with high thermal efficiency at hypersonic flow conditions.

Experimental study of forward-facing cavity with energy deposition in hypersonic flow conditions

Controlling nose surface peak heat fluxes is crucial to the design of hypersonic vehicles. In this study, we report a novel technique of convective heat flux reduction on the nose surface of a spherically blunted cone by employing a forward-facing cavity combined with heat energy deposition inside the cavity. The heat deposition is achieved by the exothermic reaction of a chromium film coated on the cavity surface. The experiments are performed in hypersonic shock tunnels using air as the test gas at free stream stagnation enthalpy conditions of 2.2 ± 0.08 MJ/kg (H1), 3.2 ± 0.018 MJ/kg (H2), and 5.4 ± 0.02 MJ/kg (H3), for a geometry of cavity length to diameter ratio of 1. Schlieren images of the flow are captured using a high-speed camera and a high-power pulsed diode laser light source. The surface heat flux measurements were performed using calibrated platinum thin film sensors. We observed that the heat deposition altered the cavity flow field significantly by lowering the flow oscillation frequency and increasing the shock standoff distances with higher oscillation amplitudes. The overall surface mean heat flux reduction is increased from ≈13% to ≈49% compared to the blunt body geometry, whose nose radius is 30 mm and enhanced with reference to the cavity without heat deposition from ≈15% to ≈35%. Chromium film surface reactions are studied using X-ray Photoelectron spectroscopy, and the results confirm that the exothermic surface reactions of the Cr thin film are attributed to the formation of oxides and nitrides of Cr.

Analysis of Morphing Modes of Hypersonic Morphing Aircraft and Multiobjective Trajectory Optimization

Aiming at improving the range of wing-body combination aircraft at hypersonic flow conditions and exploring the application of morphing technology in hypersonic aircraft, morphing aircrafts with different morphing modes have been proposed. The aerodynamic characteristics and wing efficiency of morphing aircraft with different morphing modes have been studied for a further application. In order to verify the significance of morphing technology on the trajectory, the trajectory of the glide phase has been optimized through multiobjective optimization method. Through a 3 degree-of-freedom dynamic model and a heat flux model, the range of glide trajectory and the total heat of the leading edge of the wing are calculated and selected as optimization objectives in the multiobjective optimization problem. The optimization variables include the Mach numbers when the aircraft is morphing (morphing timing) and the angle of attack of different phases. Based on the multiobjective evolutionary algorithm based on decomposition, the multiobjective trajectory optimization problem is solved and a uniform Pareto Front is obtained, and through the analysis of typical solutions, it can be seen that a compromise is made to balance the two objectives. Through the comparison of morphing and non-morphing aircraft, morphing aircraft can fly further with a smaller total heat of the leading edge of the wing. Also, it seems that the variable sweep wing morphing mode has a better overall performance. The result in this paper is a verification of the application prospects of the morphing technology under hypersonic environment, which will provide a reference for further application of morphing technology in hypersonic aircraft.

Continuum Simulations of Hypersonic Flows in Chemical and Thermal Nonequilibrium

The ability of the Navier–Stokes equations to capture the effects of strong chemical and thermal nonequilibrium on gas composition in a manner similar to the direct simulation Monte Carlo (DSMC) method is tested with the introduction of a nonequilibrium chemistry model. Although this chemistry model has the ability to reasonably reproduce measured Arrhenius rates in conditions of thermal equilibrium, it results in reaction rates that vary significantly for the most commonly used Arrhenius rates in conditions of thermal nonequilibrium. The nonequilibrium chemistry model is implemented in the three-temperature Navier–Stokes computational fluid dynamics (CFD) solver, Data-Parallel Line Relaxation (DPLR), and tested on high-altitude, hypersonic flow conditions. The results of the simulations show that the model predicts a greater amount of NO compared with previous Navier–Stokes using nonequilibrium quasi-classical trajectory derived rates and remains consistent with DSMC computations over altitudes ranging from 53.5 to 87.5 km. The most commonly used chemistry model for Navier–Stokes solvers was found unable to match this performance, indicating the importance of including nonequilibrium effects when modeling chemically reacting hypersonic flowfields.

A surrogate modeling approach for reliability analysis of a multidisciplinary system with spatio-temporal output

Reliability analysis of a multidisciplinary system is computationally intensive due to the involvement of multiple disciplinary models and coupling between the individual models. When the system inputs and outputs are varying over time and space, the reliability analysis is even more challenging. This paper proposes a surrogate model-based method for the reliability analysis of a multidisciplinary system with spatio-temporal output. The transient characteristics of the multidisciplinary system output under time-dependent variability are analyzed first. Based on the transient analysis, surrogate models are built for individual disciplinary analyses instead of a single surrogate model for the fully coupled analysis. To address the challenge introduced by the high-dimensionality of spatially varying inter-disciplinary coupling variables, a data compression method is first employed to convert the high-dimensional coupling variables into low-dimensional latent space. Kriging surrogate modeling is then used to build surrogates for the individual disciplinary models in the latent space. Based on the individual disciplinary surrogate models, reliability analysis of the coupled multidisciplinary system under time-dependent uncertainty is investigated. Further, epistemic uncertainty sources, such as data uncertainty and model uncertainty, lead to uncertainty in the reliability estimate. Therefore, an auxiliary variable approach is used to efficiently include the epistemic uncertainty sources within the reliability analysis. An aircraft panel subjected to hypersonic flow conditions is used to demonstrate the proposed method. The analysis involves four interacting disciplinary models, namely, aerodynamics, aerothermal analysis, heat transfer, and structural analysis. The results show that the proposed method is able to effectively perform reliability analysis of a multidisciplinary system with spatio-temporal output.

Drag and Heat Reduction Performance for an Equal Polygon Opposing Jet

AbstractAn equal polygon opposing jet can withstand huge wave drag and serious aerodynamic heating in hypersonic flow conditions. The opposing jet is able to change the flow field structure, and then it improves the aerodynamic characteristic of the hypersonic vehicle. In order to get more information about the flow field characteristics of the opposing jet, the schemes with equal polygons for the opposing jet were designed and their properties with different polygons have been investigated numerically in the paper. Also, the numerical method has been validated against the available experimental data in the open literature. The obtained results show that the drag-reduction performance is best when the number of jet angles (N) is 7, and its value reaches 26.4%. At the same time, its wall maximum heat flux is the smallest and the performance for heat protection is the best. Moreover, the maximum heat flux can be decreased by 60.6%. N has a slight influence on the position of the shock wave. When N is big en...

Research on electron density measurements for high enthalpy flows

The high enthalpy shock tunnel, which can provide the high temperature gas conditions for hypersonic flight, is an effective ground facility for the research of reentry problems. It also has the ability of studying real gas effect. The hypersonic test flow condition, with the stagnation enthalpy of 16 MJ/kg and stagnation temperature of 7900 K, was achieved in the JF-10 high enthalpy shock tunnel. The electron density of the flow-field for the 10 degree wedge model was measured using the electrostatic probe, which can obtain sufficient spatial resolution without influencing the flow structure. Results showed that the electrostatic probe has high sensitivity and the spatial electron density can be preferable acquired. The constant bias method can get the dimensionless electron density coupling with the flow field parameters, but the new developed sweeping circuit can get the quantitative value on the basis of effectively reducing the induced noise in the circuit.