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Among the design requirements of hypersonic vehicles, reducing aerodynamic heating and drag force simultaneously is the main challenge. This paper proposes a novel spike aerodome-channel combination concept to realize the flow field reconstruction around the hypersonic blunt body. The novel configuration is investigated in the axisymmetric flow at a Mach number of 6 at zero angle of attack. The two-dimensional Reynold-averaged Navier–Stokes equations are numerically solved, and the shear-stress transport k–ω model is the turbulence model implemented in this study. Parameters such as spike length and lateral jet location are investigated to explore the drag and heat reduction performance and the flow control features. The obtained results show that the application of the novel spike aerodome-channel concept alters the flow field by eliminating or replacing the strong bow shock wave, and the design of hypersonic vehicles can benefit from the application of the proposed concept. The blunt body coupled with a frustum of cone-tipped spike-channel configuration provides a remarkable drag reduction effect of 20.71% with respect to the case without channel. Considering the effect of lateral jet location, the drag reduction performance of the case with LR = 0.75 is superior to that of the root jet case at the same spike length, and a considerable drag reduction of 28.93% is obtained with L/D = 2.4. In addition, longer spike length is beneficial for improving drag reduction performance, while excellent efficiency of heat protection is obtained in a certain spike length range. For the case of L/D = 1.6 with root jet, the peak Stanton number is significantly decreased by 33.51%. [ABSTRACT FROM AUTHOR]

Sustained hypersonic flight within the atmosphere causes aerodynamic heating, which poses huge challenges for the thermal protection systems of hypersonic aircraft. Therefore, the heat flux on the aircraft surface needs to be computed accurately during the aircraft design stage. Previous approaches have not been able to achieve simultaneous accuracy and efficiency when computing the heat flux. To deal with this problem, an efficient heat flux prediction method based on deep learning techniques, called SA-HFNet, is proposed in this paper. SA-HFNet tries to learn the relationship between the heat flux and the aircraft shape and flight conditions using deep neural networks without solving the Navier–Stokes equations. Unlike other intelligent methods, SA-HFNet can automatically become aware of changes in aircraft shape. As far as we know, it is the first intelligent method that is able to obtain the heat flux quickly and adapt to changes both in the global aircraft shape and in local shape deformation. Extensive experimental results show that SA-HFNet achieves promising prediction accuracy in less time compared with computational fluid dynamics methods. Furthermore, SA-HFNet has good generalization capability because it has the potential to predict the heat flux for previously unseen aircraft shapes. [ABSTRACT FROM AUTHOR]

Gas–surface interactions between thermal protection materials and high-enthalpy nonequilibrium flow are one of the greatest challenges in accurately predicting aerodynamic heating during supersonic flights. Finer microscopic details of flow properties are required for elaborate simulation of these interactions. Spectral insight, with quantum-state-specific characteristics, is provided in this work to investigate the physico-chemical processes in high temperature interface of a carbon/carbon (C/C) composite. The nonequilibrium air flow is produced by a 1.2 MW inductively coupled plasma wind tunnel at an enthalpy of 20.08 MJ/kg. The duration of each test is up to 100 s, and quartz is also tested for comparison. Spectral insights into the reaction mechanisms of the gas–surface interactions are acquired by the optical emission spectroscopy and laser absorption spectroscopy. Dynamic evolution of the chemical reaction pathways and thermal nonequilibrium are discussed based on the results of optical emission spectroscopy. Temporally and spatially resolved results of the translational temperature and number density of atomic oxygen are quantified by laser absorption spectroscopy. Controlling mechanisms in the surface chemistry are further analyzed in conjunction with the surface temperature, scanning electron microscopy, and energy dispersive spectroscopy. Reaction mechanisms on the C/C composite surface sequentially experience an oxidation-dominant, an intense competitive, a nitridation-dominant, and a recession dominant period. Distributions in the axial direction and dynamic characteristics of the translational temperature and number density of atomic oxygen are found closely related with surface swelling, recession, and chemical reactions. The results herein are consistent with each other and are instructive to further investigate the interface evolution on C/C composites. [ABSTRACT FROM AUTHOR]

A conjugated computational fluid dynamics model using convective thermal boundaries is proposed for the thermal analysis of the regenerative cooling system of an X-51A-like hypersonic aircraft in conditions closer to realistic scenarios. By X-51A-like, we mean that the convective boundaries of the cooling system are determined according to the structure and flight conditions (Mach 6 and an altitude of 20 km) of the X-51A Waverider. The aerodynamic heating, supersonic combustion, and convective heat transfer with an interior (fuel tank) are simulated from an engineering perspective. The flow, heat transfer, and pyrolysis characteristics of endothermic hydrocarbon fuels (EHFs) flowing in B-channels (bottom), S-channels (side), and T-channels (top) considering buoyancy effect are simulated and compared. A distinct temperature distribution difference in the solid region is observed between the cases with different boundary conditions. However, the cooling performance of EHFs is insensitive to the type of thermal boundary. Five indices, the maximum temperature, outlet temperature, n-decane conversion, total heat sink, and pressure drop, are compared to comprehensively assess the cooling performance. The cooling demands in the B-channels and S-channels are about 1.3 times those in the T-channels because of the additional cooling effect from the adjacent fuel tank. This study should be of great significance in the practical and systematic design of regenerative cooling systems. [ABSTRACT FROM AUTHOR]

An effective control strategy is studied in this work to reduce aerodynamic heating inside transitional hypersonic boundary layers, which is of both scientific significance and practical importance. The experiments show that the use of grooved wavy walls covered with an acoustic transparent film can successfully reduce the growth of Mack's second mode, which delays transition and results in significant reduction of aerodynamic heating. [ABSTRACT FROM AUTHOR]

The two main problems in hypersonic vehicles are wave drag and aerodynamic heating. This work proposes a novel method for drag reduction by combining spikes with the plasma synthetic jet actuator (PSJA). Numerical simulations are performed to better understand the drag reduction mechanism with an incoming flow at Mach 6. The results suggest that the flow field is affected primarily by the diffracted wave and synthetic jet. The maximum drag reduction reaches 47.7% with the plasma spike compared with the opposing jet. A better drag reduction effect is achieved when increasing the energy density of the PSJA, while the propagation velocity of the diffracted wave remains constant. A wider and faster jet is obtained with a larger PSJA orifice diameter. However, the control time of the jet shortens. A mode conversion occurs when the orifice diameter is 1 mm. Furthermore, the maximum drag reduction rate increases from 37.6% to 49.0% when the length diameter ratio (L / D) increases from 0.5 to 1.5. The effect of spike length on drag reduction decreases gradually at greater lengths. [ABSTRACT FROM AUTHOR]

One of the key challenges for accurate prediction of hypersonic aerodynamic heating is the exothermic uncertainty due to the complex surface catalytic recombination effect, which is caused by the strong interactions between highly non-equilibrium dissociated gas and the thermal protection material surface. Employing engineered surface morphology to improve thermal protection effects has been proposed, but its effects on surface catalytic recombination remain unclear. To address this problem, this work employs the reactive molecular dynamics method to investigate the surface adsorption and recombination characteristics of continuous impingement of atomic oxygen upon eight different nano-structured silica surfaces. A parametric study of the influences of the gas incident angles and the surface structural parameters, i.e., roughness factor and surface fraction, is conducted. The results show that the surface catalytic recombination performance is very sensitive to the incident angle of the incoming gas, and the presence of nanostructures increases the recombination rate. The influence of surface morphology shows a complicated feature, where nanostructures with moderated fin height and high surface fraction are beneficial for the inhibition of surface recombination effects, leading to reduced exothermic heat release. Such microscopic revelation of the surface morphology effect is helpful for accurate prediction of aerodynamic heat and provides guidance for the surface engineering of optimized morphology to achieve improved thermal protection effect. [ABSTRACT FROM AUTHOR]

Silica-based ceramic material is one of the most competitive options of thermal protection material. However, catalytic reaction mechanism of gaseous atoms on it is complex and confusing. To model catalytic recombination of dissociated oxygen atoms accurately in prediction of chemical non-equilibrium flow and aerodynamic heating, the reactive molecular dynamics method was adopted to simulate gas–surface interaction on the interface, and a series of post-processing methods were constructed to analyze recombination pathways of atomic oxygen on α-quartz. It was found that there are four types of adsorbates on α-quartz surface and five pathways to produce recombined oxygen molecules. Recombination pathways would change from Eley–Rideal recombination-dominated to molecule desorption-dominated with increase in temperature. Information extracted by current post-processing methods explains how and why the recombination coefficient changes with temperature. The post-processing methods can be further applied in analysis of catalytic recombination on other thermal protection materials. [ABSTRACT FROM AUTHOR]

A multi-physics thermochemical non-equilibrium model is established to study the flow characteristics of the plasma sheath around an atmospheric reentry demonstrator. This model includes the tight coupling of Navier–Stokes equations, 54 chemical reactions of air, and a four-temperature model. The processes of dissociation, ionization, and the internal energy exchanges of air components were successfully simulated during aerodynamic heating of the reentry vehicle. The distributions of plasma sheath temperature, the molar fraction of air species, stagnation pressure, surface pressure, and electron number density around the reentry vehicle were obtained at different flight altitudes. Additionally, to validate the numerical model developed in this study, the flow characteristics of the Radio Attenuation Measurement-C-II (RAM-C-II) vehicle are also simulated and then compared with corresponding experimental data. They show good consistency in general. It is found that when the vehicle is at a high flight altitude, there is a strong thermochemical non-equilibrium phenomenon around the vehicle. However, the plasma sheath tends to be in local thermal equilibrium at a low flight altitude. The distance from the shock layer to the stagnation point decreases with a decrease in reentry altitude from 90 to 65 km but increases with a decrease from 65 to 40 km. The electron number density in the shock layer is maximum. The distribution of the electron number density in the wake region differs significantly at different flight altitudes. [ABSTRACT FROM AUTHOR]

Acoustic waves existing in hypersonic boundary layers act as a heat pump that transfers energy from the sonic line to the wall causing the wall temperature to rise, which explains the newly identified aerodynamic heating related to Mack's second mode from the perspective of thermoacoustic effects. The analysis of data from direct numerical simulations shows that Mack's second mode, belonging to the family of trapped acoustic waves, is highly amplified in a Mach 6 boundary layer and becomes sufficiently strong to affect the mean wall-normal temperature gradient, and the energy transport in the wall-normal direction due to the thermoacoustic effect balancing the thermal conduction brought by the additional temperature gradient. [ABSTRACT FROM AUTHOR]

The need to increase the payload capacity of the rockets motivates the development of high-power rocket engines. For a chemical propulsion system, this results in an increasing thermal load on the structure, especially the combustion chamber and nozzle must be able to withstand the extreme thermal load caused by high-temperature and high-pressure combustion gas. In order to protect the structure from the effect of increasing heat flux, it is necessary to counteract such effect with more advanced thermal management technology. This requires us to accurately predict the aerodynamic heating of the structure by high-temperature and high-speed combustion gas. In this study, a high-temperature combustion gas tunnel developed in the laboratory is used to produce high-speed combustion gas. Combined with the results of numerical calculation, the flow and aerodynamic heating characteristics of air and hydrogen–oxygen combustion gas under the same total temperature and pressure are analyzed and compared. The comparison revealed that the combustion gas flow in the nozzle has higher static temperature, velocity, and smaller Mach number. When the combustion gas flows around the sphere, the shock standoff distance and stagnation pressure are smaller than those of air, and the wall heat flux is much larger than that of air. The active chemical reaction in the combustion gas makes the aerodynamic heating of the structure more severe. Finally, through the analysis of a large amount of data, a semi-empirical formula for the heat flux of the stagnation point heated by a high-speed hydrogen and oxygen equivalent ratio combustion gas is obtained. [ABSTRACT FROM AUTHOR]

Hyperthermal atomic oxygen (AO) bombardment to thermal protection system surface has been identified to impact the aerodynamic heating significantly, due to complex chemical reactions at the gas–solid interface, e.g., surface catalysis recombination, oxidation, and ablation. Previous investigations have focused on the surface effects of the AO collision process, while the influence of impacting gas characteristics remains unclear under various non-equilibrium aerodynamic conditions. This work conducts a reactive molecular dynamics (RMD) study of AO collisions over graphene surface, by considering the incoming gas at different translational energies (0.1 ≤ Ek ≤ 10 eV), incident angles (θ = 15°, 30°, 45°, 60°, 75°, and 90°), and O/O2 ratios (χO2 = 0.00, 0.25, 0.50, 0.75, and 1.00). The RMD results indicate that for AO normal incidence, the predominant reactive products of O2, CO, and CO2 molecules are produced due to the synergistic catalytic recombination and surface ablation reaction effects. A maximum recombination performance is identified around 5-eV AO incidence. For off-normal AO incidence, the recombination coefficient increases with the increase in incidence angle from 15° to 60° due to the larger perpendicular components of translational energy and then decreases smoothly. With the increase in O2 mole fraction, the surface reflection probabilities increase, which result in the decrease in both catalytic recombination and ablation activities. Via revealing the atomistic-scale mechanism of gas effects on the surface under hypersonic non-equilibrium conditions, this work sheds light for the future design and optimization of thermal protection materials. [ABSTRACT FROM AUTHOR]

The Fokker–Planck-type approximation of the full Boltzmann equation has aroused intense research interest due to its potential for the stochastic particle simulation of rarefied gas flows. The ellipsoidal statistical Fokker–Planck (ES-FP) model treats the evolution of molecular velocity as a continuous stochastic process, and it satisfies the basic requirements for a proper gas-kinetic model including the H-theorem and an adjustable Prandtl number. The ES-FP model can be numerically implemented with computational particles in a Monte Carlo manner. Two different particle ES-FP schemes are presented. The first scheme utilizes the exact stochastic integral solution of the Langevin equations corresponding to the ES-FP equation and couples free-molecular moves and intermolecular collisions. The second scheme is designed to enforce the conservation of momentum and energy during the numerical simulation based on the decoupled algorithm and the analysis of the specific conditions for the conservative property. Numerical tests are conducted to demonstrate the performances of different schemes. In the simulation of a homogeneous gas system, the ES-FP scheme without enforcement of conservation leads to unphysical variation in the momentum and loss in energy, whereas the conservative ES-FP scheme strictly maintains the momentum and energy of the system. For the Mach 6 flows over the leading edge of a flat plate and over a round-nosed blunt body, the non-conservative ES-FP scheme underestimates the shock angle and the shock standoff distance, makes inaccurate predictions of aerodynamic force and heating, and produces low-temperature anomalies in front of the shock waves. In comparison with the results given by the direct simulation Monte Carlo method, the results of the conservative ES-FP simulations show satisfactory accuracy for the flow fields as well as the distributions of pressure, friction, and heat flux on the wall surfaces. [ABSTRACT FROM AUTHOR]

The basic problems of transition in both incompressible and compressible boundary layers are reviewed. Flow structures in low-speed transitional and developed turbulent boundary layers are presented, together with almost all of the physical mechanisms that have been proposed for their formation. Comparisons of different descriptions of the same flow structures are discussed as objectively as possible. The importance of basic structure such as solitonlike coherent structure is addressed. For compressible flows, the receptivity and instability of boundary layer are reviewed, including the effect of different parameters on the transition. Finally, the principle of aerodynamic heating of hypersonic boundary layer is presented. [ABSTRACT FROM AUTHOR]

The influence of a wavy wall on the hypersonic boundary layer transition and the related aerodynamic heating is investigated on a flared cone at a range of unit Reynolds numbers. Experiments are conducted in a Mach 6 wind tunnel using Rayleigh-scattering flow visualization, fast-response pressure sensors, and infrared thermography. The results show that compared to the smooth-wall cone, the wavy-wall one can suppress the second-mode instability to a certain degree and eliminate the local heating spot before transition is completed (denoted as HS). These verify our recent work on aerodynamic heating that HS is caused by the second-mode instability. [ABSTRACT FROM AUTHOR]

The evolution of second-mode instabilities in hypersonic boundary layers and its effects on aerodynamic heating are investigated. Experiments are conducted in a Mach 6 wind tunnel using fast-response pressure sensors, fluorescent temperature-sensitive paint, and particle image velocimetry. Calculations based on parabolic stability equations and direct numerical simulations are also performed. It is found that second-mode waves, accompanied by high-frequency alternating fluid compression and expansion, produce intense aerodynamic heating in a small region that rapidly heats the fluid passing through it. As the second-mode waves decay downstream, the dilatation-induced aerodynamic heating decreases while its shear-induced counterpart keeps growing. The latter brings about a second growth of the surface temperature when transition is completed. [ABSTRACT FROM AUTHOR]