Hypersonic Flow Field Research Articles

Exploring the drag reduction caused by a combined spike and jet setup for a hypersonic blunt object at low elevations

Abstract When hypersonic vehicles fly at low altitudes during the end of the trajectory, they often face extremely high dynamic pressure and greater resistance. Studying the drag reduction features of the hypersonic vehicle at low altitudes is highly important. The Reynolds-averaged N-S equations are solved using the finite-volume method with the SST k-ω turbulence model. Numerical simulations are conducted on three-dimensional hypersonic flow fields involving spike and combined spike and jet configurations at low altitudes with hypersonic speeds to investigate drag reduction effects and analyze flow field characteristics. The data indicates that the presence of a spike can effectively reduce drag during low-altitude hypersonic flight. Enhancing the spike length, its drag reduction efficiency first increases and then slightly decreases. Enhancing the Mach number, the length of the spike that achieves the best drag reduction efficiency becomes longer. A spike and jet combination can enhance drag reduction efficiency when flying at an angle of attack. Enhancing the spike length or the pressure ratio in this configuration can enhance its effectiveness in reducing drag. The findings of this study provide practical guidance for developing drag-reduction strategies for hypersonic aircraft flying at low altitudes.

Aero-optical effects of a hypersonic vehicle based on the refractive index step ray tracing method

The aero-optical effect caused by the high-speed flight of aircraft can affect the accuracy of astronomical navigation, so studying the laser transmission characteristics in high-speed flow fields is essential. This study adopts a 3D spherical blunt cone model as the physical aircraft model, incorporates the latest reanalysis data from the National Centers for Environmental Prediction as atmospheric parameters, and uses Fluent software to establish a hypersonic aircraft external flow field model. The ray tracing error is calculated using a method based on the refractive index step size. Results indicate that, under the conditions of flying altitude of 5 km and incident angle of 45°, the proposed ray tracing method has better accuracy than the traditional method. Compared with the Runge–Kutta method and the method based on geometric step size, the proposed ray tracing method reduces the error by 46.4% and 12.5%, respectively, at the flight Mach number of 5.

A novel graph modeling method for GNN-based hypersonic aircraft flow field reconstruction

Detection of flow fields constitutes a critical role in the advancement and innovation of hypersonic aircraft. Under hypersonic conditions, aircraft aerodynamics manifest a multitude of complex phenomena, including intense turbulence and fluctuations, transitions within the boundary layer, interactions between shock waves and boundary layers, as well as the effects of high-temperature gas. Thus, the surveillance of hypersonic aircraft flow fields is imperative not only for flight safety but also for the progression of hypersonic technologies. Given the practical limitations that restrict sensors installed to only key areas for detection, we propose a GNN-based method for hypersonic aircraft flow field reconstruction from limited sensors. Moreover, we have introduced a novel graph modeling technique for enhancing the reconstruction motivated by that solely the most significant edges within the graph are efficacious for field reconstruction. The methodology encompasses the following steps, modeling graphs from diverse perspectives, and transforming them into minimum spanning trees. These sparse graphs are then integrated with learnable weights optimized by automatic differentiation. Our method has been tested on an open-source turbulence dataset, the two components of graph pruning and weight optimization bring about 34% and 10% improvements on average. Furthermore, our approach has been validated in the reconstruction of hypersonic aircraft flow fields for at least 10% error reduction among all the situations.

Hypersonic inlet flow field reconstruction dominated by shock wave and boundary layer based on small sample physics-informed neural networks

High Mach number inlets face complex challenges such as shock wave/boundary layer interference, adversely impacting the aerodynamic characteristics and stable operating boundaries. Traditional computational fluid dynamics (CFD) simulations for performance and flow field analysis are slow, and data-driven deep learning technologies struggle with predicting flow fields in complex aerodynamic phenomena like shock wave/boundary layer interactions, boundary layer separation, and Mach disks. This study introduces a rapid, robust method for reconstructing hypersonic viscous inlet flow fields using inlet geometry design parameters. This method is called shock wave and boundary layer dominated two-dimensional multiphysical field reconstruction model based on multi-scale sensory field fusion residual physical information neural network (PIMSRF_ResNet), ensuring high-precision training data. Using the optimal Pareto solution set from multi-objective optimization with a small sample intelligence method, 150 sets of multi-physics fields under Ma10 conditions with varying geometric design parameters are calculated. Compared to traditional pure data-driven models, PIMSRF_ResNet's structural similarity in the test set reaches 0.9773, with a peak signal-to-noise ratio close to 30 dB and a correlation coefficient exceeding 98%. These experimental results demonstrated that the proposed method is a promising tool for real-time prediction of complex internal flow fields dominated by high Mach number shock waves and boundary layers.

Three-Dimensional Radiation Analysis of the Hypersonic Jules Verne Reentry Spectra

The automatic transfer vehicle (ATV) Jules Verne, developed by the European Space Agency, played a crucial role in resupply missions to the International Space Station from 2008 to 2015. Following its resupply missions, Jules Verne was programmed to execute a planned destructive reentry into Earth’s atmosphere. To better understand the spectra generated by the Jules Verne, three-dimensional hypersonic flowfield and radiation analyses are performed at the 75 km trajectory point of the ATV. Results are compared to measurements obtained at 78 and 73 km by a miniature Echelle spectrograph operated from an airborne platform. Incorporating three-dimensional effects into the radiation analysis results in an overall reduction of the total radiation spectra and a decrease in atomic and molecular features, leading to a closer alignment between the observed and simulated values. Comparable effects are observed by making minor adjustments to the internal mode temperatures and modifying the model for the population of excited state number densities.

Investigation of the influence mechanism of hypersonic flow field environment on thermionic emission efficiency

Electron Transpiration Cooling (ETC) is a novel thermal management method that utilizes thermionic emission to cool the leading edge surface of a hypersonic vehicle. Establishing a thermionic emission model that accounts for the hypersonic flow field environment is imperative for accurately predicting the thermal protection efficiency of ETC. A thermionic emission dual-sheath model was established, considering the impact of hypersonic non-equilibrium flow field particle parameters on thermal electron emission. The virtual cathode threshold was obtained through numerical analysis, considering the influence of particles parameters and surface thermoelectric material parameters in a non-equilibrium flow field. Distribution of sheath potential and particle number density in plasma sheath region were performed at a range of conditions. Thermionic emission efficiency approached 80 % as the plasma ion density was 5 × 20 m−3 and the work function was 2 eV. The results showed that plasma ions have different degrees of neutralization effect on thermal emission electrons depending on the flow field and surface material characteristics, ultimately causing varying surface thermionic emission efficiency.

Noise Generated in a Scramjet Combustor

This paper presents the noise levels generated inside a two-dimensional scramjet combustor at Mach 7.3 freestream using the focused laser differential interferometric (FLDI) technique. FLDI measurements were taken at the front, middle, and rear of the combustor’s center plane under unfueled, combustion-suppressed, and combustion-on conditions. Spectral analysis of the FLDI signal shows density fluctuations inside the scramjet combustor are almost two times higher than the freestream and confirms the capability of this technique for measuring the disturbances in a complex hypersonic flowfield. Though combustion-induced pressure rise is low, a trend in the increase in the normalized density fluctuations/noise levels due to supersonic combustion is observed across the measurement locations for the combustion-on condition. However, the overall increment in the noise level is modest between the three conditions and across the three probing locations. This infers that, for the current scramjet geometry, fueling rates, fuel–air mixing, and supersonic combustion are not the dominant noise sources in the scramjet combustor. Instead, the base hypersonic flowfield inside the scramjet combustor is the main contributor to the noise field. Noting that there is relatively large uncertainty in the measured noise levels and that the combustion-induced pressure rise is modest in the present study, further investigation is needed to check the applicability of these results to the scramjet engines with flow paths having more complex flowfields, fueling schemes, and fueling rates.

Long-lived nitric oxide molecular tagging velocimetry with 1 + 1 REMPI.

The successful demonstration of long-lived nitric oxide (NO) fluorescence for molecular tagging velocimetry (MTV) measurements is described in this Letter. Using 1 + 1 resonance-enhanced multiphoton ionization (REMPI) of NO at a wavelength near 226 nm, targeting the overlapping Q1(7) and Q21(7) lines of the A-X (0, 0) electronic system, the lifetime of the NO MTV signal was observed to be approximately 8.6 µs within a 100-Torr cell containing 2% NO in nitrogen. This is in stark contrast to the commonly reported single photon NO fluorescence, which has a much shorter calculated lifetime of approximately 43 ns at this pressure and NO volume fraction. While the shorter lifetime fluorescence can be useful for molecular tagging velocimetry with single laser excitation within very high-speed flows at some thermodynamic conditions, the longer lived fluorescence shows the potential for an order of magnitude more accurate and precise velocimetry, particularly within lower speed regions of hypersonic flow fields such as wakes and boundary layers. The physical mechanism responsible for the generation of this long-lived signal is detailed. Furthermore, the effectiveness of this technique is showcased in a high-speed jet flow, where it is employed for precise flow velocity measurements.

Experimental study on shock interaction control of double wedge in high-enthalpy hypersonic flow subject to plasma synthetic jet

The hypersonic shock-shock interaction flow field at double-wedge geometries controlled by plasma synthetic jet actuator is experimentally studied in a Ma = 8 high-enthalpy shock tunnel with the purpose of exploring a novel technique for reducing surface heat flux in a real flight environment. The results demonstrate that increasing the discharge energy is advantageous in eliminating the shock wave, shifting the shock wave interaction point, and shortening the control response time. The oblique shock wave can be completely removed when the actuator's discharge energy grows from 0.4 J to 11.5 J, and the displacement of the shock wave interaction point increases by 124.56%, while the controlled response time is shortened by 30 μs. Besides, the reduction in diameter of the jet exit is firstly proved to have a negative impact on energy deposition in a working environment with incoming flow, which reduces the discharge energy and hence decreases the control effect. The shock wave control response time lengthens when the jet exits away from the second wedge. Along with comparing the change in wall heat flux at the second wedge over time, the control effect of plasma synthetic jet actuator with and without inflation is also analyzed. When plasma synthetic jet works in inflatable mode, both the ability to eliminate shock waves and the shifting effect of the shock wave interaction point are increased significantly, and the wall heat flux is also reduced.

A stability-enhanced approach for aerodynamic heating computation in high-speed flows

Near-space vehicles face severe aerothermal environments due to hypersonic speeds. Accurately predicting hypersonic aerodynamic heating is crucial for the design of spacecraft’s thermal protection systems, aerodynamic shapes, and propulsion systems. Numerical simulations have become an important tool for computing aerothermal environments under extreme conditions. However, there is a contradiction between the robustness and low dissipation characteristics of existing numerical methods for hypersonic flow fields. This paper proposes an effective dimension hybrid approach that enhances the robustness of AUSM-family schemes while maintaining their low dissipation characteristics. Numerical experiments demonstrate the proposed approach’s effectiveness and potential for practical engineering applications.

Flow Field Reconstruction of 2D Hypersonic Inlets Based on a Variational Autoencoder

The inlet is one of the most important components of a hypersonic vehicle. The design and optimization of the hypersonic inlet is of great significance to the research and development of hypersonic vehicles. In recent years, artificial intelligence techniques have been used to improve the efficiency of aerodynamic optimization. Deep generative models, such as variational autoencoder (VAE) and generative adversarial network (GAN), have been used in a variety of flow problems in the last two years, making fast reconstruction and prediction of the full flow field possible. In this study, a hybrid multilayer perceptron (MLP) combined with a VAE network is used to reconstruct and predict the flow field of a two-dimensional multiwedge hypersonic inlet. The obtained results show that the VAE network can reconstruct the overall flow structure of the hypersonic flow field with high accuracy. The reconstruction accuracy of complex flow structures, such as shockwaves, boundary layers, and separation bubbles, is satisfactory. The flow field prediction model based on the MLP-VAE hybrid model has a strong generalization and generation ability, achieving relatively accurate flow field prediction for inlets with geometric configurations outside the training set.

Experiment and simulation of electron density distribution in discharge plasma at hypersonic speed

Pulsed discharge can generate high density and high dynamic plasma, which has promising application prospects in the field of stealth technology for high-speed aircraft. To study the evolution process of pulsed discharge plasma jet in a hypersonic flow field, the pulsed discharge experiment was performed in a hypersonic wind tunnel with 8 M in this paper. The plasma evolution process and electron density were measured by a high-speed schlieren device and spectrum acquisition system. A shock wave appeared after the blast wave generated by the discharge interacted with the external flow field. In the region below the shock wave, the plasma jet flowed downstream and produced a plasma layer. The electron density of the jet increases with the injected energy, and the peak density reaches 5.28 × 1015 cm−3. Due to the limitations of experimental measurements, based on the Navier–Stokes equations and the air dissociation and ionization model, including 11 components and 20 chemical reactions, a simulation for the experimental process was performed. At the injected energy of 495 and 880 mJ, the difference between the simulated electron density and the experimental value is 16.09% and 15.34%, respectively. The thickness of the plasma layer initially increases and then decreases over time, with higher injected energy leading to a thicker layer. Specifically, when 880 mJ of energy is injected, the plasma layer can reach a maximum thickness of 6.69 cm. The collision frequency fluctuates around 1 GHz, and the collision frequency at the upper edge of the plasma layer is large.

Flow Characteristic Study on Wake Flow of High-speed vehicles

Based on the general CFD software NNW-FlowStar of national numerical wind tunnel engineering, this paper adopts anisotropic unstructured hybrid mesh to simulate the wake field characteristics of aircraft in high-speed flight. NNW-FlowStar software in high resolution numerical format, gaussian type node gradient calculation method, massively parallel and a series of numerical algorithm in terms of development and improvement to ensure the fine simulation of the hypersonic flow field. The numerical method can clearly aircraft wake flow field, the main shock, the corner expansion area, shear layer and the shock wave, clearly shows the wake flow field structure. With the increase of Mach number, the wake flow is getting more and more strong, the backwater area range is more and more small, after the stagnation point, the shock wave is also much greater compressibility.

Numerical Analysis of Pressure Load on Hypersonic Wing

The flight vibration environment of hypersonic vehicles has always been a hot topic in the field of aerospace at home and abroad. In this paper, the simple wing model and the outer flow field grid of the wing are established by using fluent software. A series of environmental parameter conditions are set in the hypersonic flow field, and the steady and unsteady flow fields are analyzed. The distribution characteristics of the pressure load on the wing surface in the hypersonic flow field are obtained. The results show that the pressure load measured at the head position of the wing model is larger, and the pressure load at the fuselage position is smaller and changes more gently. With the increase of Mach number, the pressure load on the wing model will also increase. Then MatLab software is used to convert the time domain curve of wing surface pressure load into a power spectral density curve, which provides load input for subsequent random vibration response simulation on finite element software.

Numerical simulation of hypersonic plasma flow field disturbed by pulsed discharge

A plasma sheath will be generated around the hypersonic vehicle during reentry, and a large number of electrons in the plasma sheath will seriously affect the communication between the vehicle and the ground station. In order to reduce the electron number density of the hypersonic vehicle plasma sheath, a method of using pulsed discharge active actuation to regulate the plasma sheath is proposed. Based on the air dissociation and ionization model including 11 components and 32 chemical reactions, the reduction effect of pulsed discharge actuation on the electron density of plasma sheath is studied by numerical simulation. A first test is performed in which the pulsed discharge is compared with the plasma jets' experimental data. Then, a second test compared the plasma flow field around the RAMC-II vehicle with the flight test and NASA data. In these two tests, the simulation results are basically consistent with the experimental results. Finally, the effect of pulsed discharge with different energy density on the plasma sheath electron density is studied. The numerical results show that the interaction between the high-pressure aerodynamic actuation generated by the actuator and the plasma sheath produces an obvious shock wave, which blocks electrons from flowing downstream, reduces the velocity and pressure of the flow field behind the shock wave, and, finally, makes the electron density downstream of the actuator attenuate significantly, with the maximum attenuation amplitude of about 35%. Compared with the traditional method, the method proposed in this paper requires less space, load, and source power and has certain engineering feasibility.

Novel Engineering Methodology for Decoupled Aerothermal Analysis of Hypersonic Atmospheric Entry Flows

Analyses of thermal protection system response to atmospheric entry are generally performed in a decoupled manner where some terms in the surface balance equations are simplified. In such an approach, surface fluxes are calculated using a fluid dynamics solution of the hypersonic flowfield. These are nondimensionalized for use by a material response code, where correction terms are applied to account for the missing physics. A new method is presented here that directly includes ablation physics, thus removing the need for correction models. This approach includes the diffusion processes of ablative species in the boundary layer and nonequilibrium surface chemistry. The new approach is compared to the heritage methodology using a trajectory designed to allow molecular dissociation and vibrational energy contribution. The new methodology predicts surface temperatures with the same qualitative trend, with the largest disagreements in areas of the trajectory where nonequilibrium effects are expected to occur. The solid ablation flux and subsequent recession behavior are consistently lower for the new method, but this discrepancy in surface thermochemistry can be decreased by adopting kinetic models with more aggressive oxidation mechanisms. It is found that a large difference in computed recession does not necessarily equate to the largest difference in heat shield sizing.

Comparative analysis of internal energy excitation and dissociation of nitrogen predicted by independently developedab initiopotential energy surfaces

In this article we provide a comparative study of rate laws for internal energy excitation and dissociation of nitrogen. These thermochemical properties are obtained by studying molecular interactions on two independently developed $a\phantom{\rule{0}{0ex}}b$ $i\phantom{\rule{0}{0ex}}n\phantom{\rule{0}{0ex}}i\phantom{\rule{0}{0ex}}t\phantom{\rule{0}{0ex}}i\phantom{\rule{0}{0ex}}o$ potential energy surfaces developed at the University of Minnesota and NASA Ames Research Center. Furthermore, a comparison of a canonical hypersonic flow field is provided, where the only modeling input to the flow simulation are the $a\phantom{\rule{0}{0ex}}b$ $i\phantom{\rule{0}{0ex}}n\phantom{\rule{0}{0ex}}i\phantom{\rule{0}{0ex}}t\phantom{\rule{0}{0ex}}i\phantom{\rule{0}{0ex}}o$ potential energy surfaces.

Aerodisk Effect on Hypersonic Boundary Layer Transition and Heat Transfer of HIFiRE-5 Vehicle

The substantial aerodynamic drag and severe aerothermal loads, which are closely related to boundary layer transition, challenge the design of hypersonic vehicles and could be relieved by active methods aimed at drag and heat flux reduction, such as aerodisk. However, the research of aerodisk effects on transitional flows is still not abundant. Based on the improved k-ω-γ transition model, this study investigates the influence of the aerodisk with various lengths on hypersonic boundary layer transition and surface heat flux distribution over HIFiRE-5 configuration under various angles of attack. Certain meaningful analysis and results are obtained: (i) The existence of aerodisk is found to directly trigger separation-induced transition, moving the transition onset near the centerline upstream and widening the transition region; (ii) The maximum wall heat flux could be effectively reduced by aerodisk up to 52.1% and the maximum surface pressure can even be reduced up to 80.4%. The transition shapes are identical, while the variety of growth rates of intermittency are non-monotonous with the increase in aerodisk length. The dilation of region with high heat flux boundary layer is regarded as an inevitable compromise to reducing maximum heat flux and maximum surface pressure. (iii) With the angle of attack rising, first, the transition is postponed and subsequently advanced on the windward surface, which is in contrast to the continuously extending transition region on the leeward surface. This numerical study aims to explore the effects of aerodisk on hypersonic boundary layer transition, enrich the study of hypersonic flow field characteristics and active thermal protection system considering realistic boundary layer transition, and provide references for the excogitation and utilization of hypersonic vehicle aerodisk.

Research on a novel combined shock control mechanism for thermal protection and drag reduction in hypersonic compressible flow field

Aerodynamic heating and extreme shock drag are major problems faced in hypersonic compressible flow fields. This paper proposes a combined flow control mechanism that combines the advantages of spike, lateral jets and aero-disc. The effects of the combined mechanism in reducing aerodynamic heating and changing the flow field structure were studied by computational fluid dynamics. The physical model is a two-dimensional axisymmetric blunt body model with a novel combination mechanism. The numerical model is the coupling of the two-dimensional axisymmetric Reynolds averaging method and the SST model. The accuracy of the numerical calculation model and calculation method is verified by comparing with the experimental data in the literature. The results show that the new combination mechanism can effectively reduce the heat flow on the surface of the blunt body, and can be used as a new type of hypersonic compressible flow field thermal protection system (TPS). The research results show that appropriately reducing the Mach number of the lateral jet and increasing the total pressure ratio of the lateral jet can further expand the subsonic recirculation zone and reduce the intensity of the reattachment shock wave. This can effectively improve the thermal protection and drag reduction efficiency of the combined structure, but put forward higher requirements on the heat resistance and mechanical strength of the upper spike. At the same time, the influence of different jet angle on thermal protection performance, drag reduction effect and changing shock structure performance is further studied. The results show that the negative jet Angle is more conducive to the effect of cooling gas, while the positive jet Angle is more conducive to the effect of changing the shock structure in hypersonic flow field.

Direct measurement of skin-friction in shock/boundary-layer interaction flows

The direct measurement of skin friction at high speed is extraordinarily complicated due to the interference from various extraneous forces and the short test-time of high-speed facilities. The present study deals with the design and performance assessment of a skin friction sensor in a hypersonic flowfield with shock-boundary layer interaction, which is a predominant high-speed flow phenomenon. The experiment was conducted in the IIT Bombay Shock Tunnel (IITB-ST); the sensor was exposed to a flow with freestream Mach number, total enthalpy, and Reynolds number of 8.4, 0.69 MJ/kg, and 2.60 × 106 m−1, respectively. The shock-boundary layer interaction was induced by an oblique shock impinged on a flat plate, with a flow turn angle of 15°. The low magnitude average skin friction could be measured with an uncertainty of ±21%. The comprehensive study conducted confirmed the suitability of the skin friction sensor for high-speed applications in impulse facilities.