高超声速烧蚀产生的质量注入对转捩的影响

Aerodynamic heating-induced ablation generates distributed roughness elements (DRE) on the surfaces of hypersonic vehicles, significantly influencing boundary layer transition and aerodynamic performance. However, location-dependent DRE modulation mechanisms remain unresolved. This study employs surface-mounted PCB® pressure sensor arrays in a Mach 6 wind tunnel to elucidate the critical role of regularly spaced DRE starting locations in modulating hypersonic boundary layer transition over a flat plate. Experimental results reveal that leading-edge DRE starting at X = 0 mm induces significant second-mode instability wave suppression. In comparison, DRE starting at X = 1 mm (with the DRE height-to-local-boundary-layer-thickness ratio k/δref = 2.83) provokes immediate bypass transition to turbulence, while downstream DRE starting at X = 120 mm (k/δref = 0.26) minimally alters natural transition. Addressing the spanwise instability non-uniformity phenomenon in the flat-plate boundary layer, proper orthogonal decomposition analyses were performed and demonstrate that it exhibits fundamentally distinct organization mechanisms across different DRE cases, with mode 3 dominating under leading-edge DRE cases, whereas mode 1 prevails in smooth and downstream DRE cases. Importantly, DRE width variations exert negligible effects on the downstream boundary layer instability across all cases, whereas the k/δref is identified as the critical scaling parameter controlling transition behavior. These findings resolve location-dependent disparities in DRE physics and establish k/δref as a predictive metric for ablation-induced DRE effects, thereby providing critical insights for the design of thermal protection systems on hypersonic vehicles.

Roughness surfaces likely present on high-speed flight vehicles due to ablation can greatly impact the laminar-turbulent transition process. In this work, effects of a randomly distributed roughness patch with roughness Reynolds number Rekk=474 on the stability and transition in a Mach 6.5 boundary-layer flow over a flat plate have been investigated via stability analyses and direct numerical simulation (DNS). The roughness patch induces several streamwise streaks downstream. The streaks slightly stabilize the Mack mode instability, yet sustain strongly unstable shear-layer modes, achieving a significantly larger integrated growth rate than the smooth case. The most amplified shear-layer mode extracts energy primarily through the spanwise velocity gradient and develops nonlinearly into hairpin vortices residing on the strongest low-speed streak. The hairpin vortices eventually contaminate the whole flowfield, leading to a fully turbulent state. We further assess the influences of wall-temperature ratio and the roughness geometry on the flowfield and the pertaining instability characteristics. The results reveal that the high wall-temperature ratio weakens the streak amplitudes and shear-layer instabilities; while randomly distributed roughness tends to induce larger-amplitude streaks than the regular counterpart with the same Rekk, the flowfield of the former one can even be more stable than the latter. We find that the spanwise gradient of streamwise velocity should also be considered along with the streak amplitude in determining the strength of shear-layer instabilities.

Severe aerodynamic heating on hypersonic vehicle nose tips generates ablation-induced distributed roughness elements (DRE), critically impacting aerodynamic performance. This work aims to reveal the effects of nose-tip DRE on the nonlinear evolution of hypersonic boundary-layer instability and elucidate the associated transition mechanism. The DRE nose tips fabricated using additive manufacturing were applied on a standard [Formula: see text] half-angle cone, and various experimental techniques were employed. The results show that nose-tip DRE can significantly amplify boundary-layer disturbance levels compared to smooth cases. Detailed analysis revealed that while nose bluntness exerts a stabilizing effect, density fluctuation eigenfunctions indicate substantial distortion of near-wall base flow by the DRE. Bispectral analysis further identified accelerated nonlinear evolution of modulated high-frequency instability waves due to DRE. In the competition between bluntness-induced stabilization and DRE-driven destabilization mechanisms, the latter appears to dominate, resulting in a significant upstream shift of transition onset with irregular serrated features in the transitional region. This work uncovered the coexistence of high-frequency instability waves and streamwise vortex structures. The modulated Mack second mode does not govern the transition process. Instead, transition under DRE cases primarily arises from the downstream breakdown of periodic streamwise vortex structures induced by larger aft-located circumferential roughness elements.

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Surface ablation induced by aerodynamic heating is a common phenomenon for high-speed cruising vehicles, impacting surface geometry, temperature distribution, and mass injection, all of which play crucial roles in the perturbation evolution and boundary-layer transition. This paper presents a high-Reynolds-number asymptotic theory to formulate the impact of a local surface ablation on the Mack-mode evolution in high-enthalpy hypersonic boundary layers. The mean-flow distortion induced by ablation is formulated by the compressible triple-deck formalism, incorporating the chemical non-equilibrium effect. Simultaneously, the distortion of the Mack mode is formulated by the multi-scale analysis, with an amplification factor quantifying the overall impact of the ablation. The asymptotic model distinctly separates the effects of the mean-flow distortion and the Mack instability property. The amplification factor is attributed to two main factors: a local scattering effect at the ablation region, primarily contributed by the indentation, and a successive adjustment of the Mack growth rate, mainly contributed by the temperature distribution. The study reveals that the Mack mode experiences enhancement by ablation when its frequency falls below a critical threshold but is suppressed for higher frequencies. Remarkably, the critical frequency aligns closely with the most unstable frequency within the second-mode frequency band.

Pyrolysis gas injection during the ablation process will affect the hypersonic boundary-layer transition. This paper proposes a three-equation [Formula: see text] transition model that considers wall mass injection. First, the boundary conditions of wall mass injection are established and verified. Second, based on analysis of the compressible boundary-layer solutions under wall mass injection, new transition correlations are developed. This allows an injection factor transport equation to be constructed using strictly local variables, and this equation is coupled with the [Formula: see text] transition model. Additionally, to model the wall mass injection effects in the fully turbulent zone, the wall boundary conditions for the specific turbulence dissipation rate are amended. The results of numerical simulations show that the transition onset locations and the changes in the heat transfer rate are reasonably consistent with experimental data.

Accurate prediction of the thermal response of ablative materials is crucial for ensuring the reliability of thermal protection systems (TPS). Ablative materials exhibit surface roughness due to material properties and manufacturing processes, which can consequently affect the transition behavior of the hypersonic boundary layer, heating rate, and skin friction of thermal protection system, etc. In this work, a hybrid micro-continuum scale approach for predicting coupled thermal flow and heterogeneous chemical ablation process in porous media is developed. The proposed model is validated through investigating the ablation process of carbon fibrous porous media under the high-temperature molecular oxygen flow. The effect of incoming flow Péclet number, porosity, and the initial pore-scale structure on the flow and thermal ablation process is further explored. The results show that during the ablation process, both high Pe number and high porosity lead to a greater ablation recession rate. Though different pore-scale structures in the porous media with same porosity exhibit almost identical average ablation rate, they can result in distinct disuniform velocity distributions within the porous media, which have a greater impact on ablation surface roughness. Under convective-dominated inflow conditions, it gradually forms a high-permeability “wormhole-like” pathway for the binary gradient porosity structure, leading to a maximum surface roughness. This proposed hybrid micro-continuum scale simulation method can potentially provide valuable pore-scale insights during the ablation of porous medium, enhancing the prediction accuracy of the material thermal response for TPS applications.

The surface roughness during ablation significantly affects the hypersonic boundary-layer transition and heat transfer. In this study, an equation of transport for the roughness amplification factor [Formula: see text] is introduced based on the improved baseline hypersonic [Formula: see text] transition model by considering the effects of surface roughness. The roughness amplification factor [Formula: see text] is convected into the flow field to represent the disturbance induced by the surface roughness. It defines a region of influence of the surface roughness to locally modify the transition model and track changes in the momentum thickness to locally modify the criteria for the onset of transition. Moreover, the boundary condition of the wall is amended for the specific rate of turbulence dissipation to model the effects of roughness in the fully turbulent zone. The improved model can accurately predict the location of transition in roughness under different hypersonic flows.

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In this paper, results from two experiments performed at California Institute of Technology’s T5 free-piston reflected shock tunnel are compared to numerical stability computations conducted using various stability analysis tools. The goal of this comparison is to begin understanding the range of boundary-layer transition predictability using different stability approaches for high-enthalpy flows. The analysis is focused on the physics of the second-mode instability at high enthalpy and the role of high-temperature effects. Although the stability solvers considering thermochemical nonequilibrium were best at estimating the measured second-mode frequency ([Formula: see text] for shot 2990, [Formula: see text] for shot 3019), they overpredicted the most amplified frequency by approximately 16–23%. A moderate spread in the predicted most amplified frequency was also observed between the different solvers. The solvers estimated a most amplified frequency range of approximately 1450–1550 kHz for shot 2990 and approximately 1525–1650 kHz for shot 3019. There was also significant inconsistency observed in predicting the peak [Formula: see text]-factor magnitude, ranging from [Formula: see text] for shot 2990 and from [Formula: see text] for shot 3019.

This study is dedicated to delaying the transition induced by broad-spectrum disturbances in a Mach 5.86 hypersonic flat-plate boundary layer through local steady blowing–suction control. The optimal blowing–suction parameters, including the position and amplitude of a single blowing suction as well as the positions of the dual blowing–suction with fixed amplitudes, are determined through direct numerical simulation employing Bayesian optimization. It is revealed that the effective single blowing–suction is located at a certain distance downstream of the synchronization point of the fast/slow mode corresponding to the dominant second mode. The effective dual blowing–suction positions are distributed around the position of the single blowing–suction. Within the parameter range considered in the study of the dual blowing–suction, the transition delay effect tends to improve with an increase in the sum of control amplitudes. The analysis of the transition mechanism found that the baseline transition case involves the presence of broad-spectrum second modes, oblique second modes, and first-mode oblique waves, resulting in the coexistence of multiple canonical transition mechanisms. Analysis of the control mechanism reveals that the transition delay primarily depends on the significant attenuation of the second mode and the maintenance of [Formula: see text]-shaped vortices induced by the first mode along the streamwise direction. A detailed investigation of the linear and nonlinear mechanisms behind the transition delay is conducted using resolvent analysis and wavelet bispectrum analysis. Results show that the optimized blowing–suction can suppress both the broad-spectrum second and first modes. Furthermore, the wavelet bispectrum analysis indicates that the suppressed nonlinear interactions between the second mode and the first mode oblique wave result in the inhibition of the evolution of [Formula: see text]-shaped vortices to the hairpin vortices. Finally, the robustness of the control strategy is explored. As the broad-spectrum disturbances present stronger three-dimensional effects, the transition delay effect weakens but eventually converges to half of the optimal delay effect.

Abstract This paper focuses on the concept of delaying laminar–turbulent transition in hypersonic boundary layers by stabilising fundamental resonance (FR), a key nonlinear mechanism in which finite-amplitude Mack modes support the rapid growth of oblique perturbations. As a pioneering demonstration of this control strategy, we introduce surface heating applied exclusively during the nonlinear phase. Unlike traditional control methods that target the linear phase, the suppressive effect of surface heating on secondary instability modes during FR is evident across various Reynolds numbers, wall temperatures and fundamental frequencies, as confirmed by direct numerical simulations (DNS) and secondary instability analyses (SIA). To gain deeper insights into this control concept, an asymptotic analysis is conducted, revealing an almost linear relationship between the suppression effect and the heating intensity. The asymptotic predictions align overall with the DNS and SIA calculations. The asymptotic theory reveals that the suppression effect of FR is primarily influenced by modifications to the fundamental-mode profile, while mean-flow distortion has a comparatively modest yet opposing impact on this process. This research presents a promising approach to controlling transition considering the nonlinear evolution of boundary-layer perturbations, demonstrating advantages over conventional methods that are sensitive to frequency variations.

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The boundary layer transition on a compound delta wing for Mach 6 has been studied experimentally and numerically. The experiment was performed at Peking University quiet wind tunnel using the Rayleigh scattering flow visualization and infrared thermography. Direct numerical simulations, under the same flow conditions, are applied to analyze the transition mechanism. The results show that the traveling cross flow vortices first appear near the leading edge of compound delta wing. These vortices modulate the mean profile of the flow due to which a rope-like structure appear in the streamwise direction, which is typical of Mack's second-mode. As Mack's second-mode grows to a sufficiently large amplitude, it triggers secondary instability, which behaves as secondary finger like structures. At the end of the transition process, low-frequency waves are excited by Mack's second-mode through an interaction mechanism with their phase speed approaching each other. It is also found that increasing the unit Reynolds number greatly promotes the aerodynamic heating as well as local hot streaks appear on both sides of the compound delta wing in the streamwise direction. The appearance of hot streaks on the compound delta wing is strongly correlated with Mack's second-mode.

Abstract Cross-flow transition over a delta wing is systematically studied in a Mach 6.5 hypersonic wind tunnel, employing the Rayleigh scattering flow visualisation, high-speed schlieren and fast-response pressure sensors. Direct numerical simulations and analysis based on linear stability theory under the same flow conditions are applied to analyse the transition mechanism. Three unstable modes are identified: the travelling cross-flow instabilities, the second mode and the low-frequency waves. It is shown that the travelling cross-flow vortices first appear in the cross-flow region near the leading edge of the model. These vortices can modulate the mean profile of the flow, which benefits the growth of second mode. A phase-locked interaction mechanism transfers energy from the cross-flow instabilities to the high-frequency second mode, leading to amplification at the expense of the cross-flow instability. As the second mode grows to a critical amplitude, it triggers a $z$-type secondary instability within a similar frequency range, which introduces secondary finger-like structures connecting to the cross-flow vortex. It is further found that the generation of these finger-like structures is related to the expansion and compression of the second mode. These finger vortices further evolve along the streamwise direction into low-frequency waves and corresponding hairpin-like structures that finally trigger turbulence. An interaction mechanism likely exists between the secondary instability and the low-frequency waves, since their phase speeds are approaching each other. These observations of the interaction mechanism are consistent with those of previous studies on hypersonic boundary layers (Zhang et al., Phys. Fluids, vol. 32 (7), 2020, 071702; Li et al., Phys. Fluids, vol. 32 (5), 2020, 051701).

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This paper performs direct numerical simulations of hypersonic boundary layer transition over a Hypersonic Transition Research Vehicle (HyTRV) model lifting body designed by the China Aerodynamic Research and Development Center. Transitions are simulated at four angles of attack: 0°, 3°, 5°, and 7°. The free-stream Mach number is 6, and the unit Reynolds number is 107 m−1. Four distinct transitional regions are identified: the shoulder cross-flow and vortex region and the shoulder vortex region on the leeward side, the windward vortex region and the windward cross-flow region on the windward side. As the angle of attack increases, the transition locations on the leeward side generally move forward and the transition ranges expand, while the transition locations generally move backward and the transition ranges decrease on the windward side. Moreover, the shoulder vortex region moves toward the centerline of the leeward side. At large angles of attack (5° and 7°), the streamwise vortex on the shoulder cross-flow and vortex region will enable the transition region to be divided into the cross-flow instability region on both sides and the streamwise vortex instability region in the middle. In addition, the streamwise vortex also leads to a significant increase in cross-flow instability in their upper region, which can generate a new streamwise vortex instability region between the two transition regions on the leeward side. Furthermore, since the decrease in the intensity and the range for the cross-flow on the windward side, the windward cross-flow region tends to become narrow and ultimately disappears.

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Experimental and linear stability theory (LST) investigation of boundary layer transition on a flat plate was conducted with a flow of Mach number 5. The temperature distributions and second-mode disturbances on the flat plate surface at different unit Reynolds number (Reunit) values were captured by infrared thermography and PCB technology, respectively, which revealed the transition location of the flat-plate boundary layer. The PCB sensors successfully captured the second-mode disturbances within the boundary layer initially at a frequency of about 100 kHz, with a gradually expanding frequency range as the distance travelled downstream increased. The evolution characteristics of the second-mode instabilities were also investigated by LST and obtained for the second mode, ranging from 100 to 250 kHz. The amplitude amplification factor (N-factor) of the second-mode instabilities was calculated by the eN method. The N-factor of the transition location in the wind tunnel experiment predicted by LST is about 0.98 and 1.25 for Reunit = 6.38 × 106 and 8.20 × 106, respectively.

Laminar-to-turbulent transition in hypersonic boundary layer on a straight cone is numerically investigated in this study. High-fidelity simulation is performed with direct-numerical simulation (DNS) coupled with the linear stability theory (LST). This study focuses on the transition scenario of fundamental breakdown, driven by the two-dimensional Mack 2nd mode and a pair of oblique modes. The major instabilities in the hypersonic boundary layer are identified by LST and introduced at the DNS inlet. Current DNS computations successfully capture intrinsic transition phenomena, including aligned vortical structures and peak heat flux in the transition process, and complete transition to turbulent flow. Appropriate numerical dissipation associated with shock-capturing methods is investigated in this study because of the presence of a nose shock outside the boundary layer and compression waves from amplified instabilities inside the boundary layer. This numerical study is conducted with two shock sensors. A classical shock sensor generates excessive dissipation in the viscous boundary layer, which artificially delays the turbulent transition. The alternative sensor reduces the unintended dissipation, allowing flow to develop turbulence within the computational domain. Computational data are discussed with relevant experimental and theoretical data.

This study presents a design approach for a uniform porous surface to control laminar-to-turbulent transition in hypersonic boundary layers. The focus is on suppressing the Mack second mode, which is a dominant instability in hypersonic boundary layers. The Mack second mode is acoustic-wave-like in the ultrasonic frequency range and can be effectively attenuated by porous surfaces. Previous studies have explored porous surfaces, either by targeting a specific frequency or by adopting geometrically complex configurations for various frequencies. In contrast, the present study proposes a porous surface design that effectively stabilizes the Mack second mode over a wide frequency range, while maintaining structural simplicity. In addition, this porous surface design incorporates constraints associated with practical fabrication to enhance manufacturability. The absorption characteristics of porous surfaces are evaluated with an acoustic impedance model, and the stabilization performance is assessed with linear stability theory. The proposed porous surface design is compared with a conventional design method that focuses on the Mack second mode with a single frequency. Consequently, the proposed design methodology demonstrates robust and consistent suppression of the Mack second mode in a broad frequency range. This approach improves both stabilization performance and manufacturability with a uniform porous surface, contributing to its practical application in high-speed vehicles.

Current flight tests and wind tunnel experiments face challenges in obtaining long-duration hypersonic flow processes and thermal environments. Numerical simulation is an effective tool to study wall thermal effects—thermal radiation heat exchange and internal thermal conduction—and their influence on flow transition. This study employs linear stability theory (LST) and direct numerical simulation to analyze a 10-mm-thick C/SiC composite flat plate under three Mach numbers (6, 10, 15) at 30 km altitude, considering internal thermal conduction. The results of hypersonic base flow, flow stability, and transition are compared with and without surface radiation under long-duration conditions. Results demonstrate that the wall temperature ratio (Sw = Tw/T0) serves as a key indicator of wall thermal effects: larger Sw increments correlate with enhanced wall thermal effects. Surface radiation suppresses the temperature rise at the wall, resulting in decreased Sw increments. Sw (with/without radiation) decrease along the streamwise direction while increasing over time. Wall thermal effects increase boundary layer thickness and induce significant modifications to the base flow profile. Larger increments in Sw occur near the plate leading edge, indicating stronger wall thermal effects in this region. LST analysis shows that wall thermal effects shift the second-mode instability to lower frequencies, reduce maximum growth rates, and delay transition locations. Higher Mach numbers reduce the Sw increment, diminishing wall thermal effects and consequently the transition delay extent.

Transition delaying is of great importance for the drag and heat flux reduction of hypersonic flight vehicles. The first mode, with low frequency, and the second mode, with high frequency, exist simultaneously during the transition through the hypersonic boundary layer. This paper proposes a novel bi-frequency synthetic jet to suppress low- and high-frequency disturbances at the same time. Orthogonal table and variance analyses were used to compare the control effects of jets with different positions (USJ or DSJ), low frequencies (f1), high frequencies (f2), and amplitudes (a). Linear stability analysis results show that, in terms of the growth rate varying with the frequency of disturbance, an upstream synthetic jet (USJ) with a specific frequency and amplitude can hinder the growth of both the first and second modes, thereby delaying the transition. On the other hand, a downstream synthetic jet (DSJ), regardless of other parameters, increases flow instability and accelerates the transition, with higher frequencies and amplitudes resulting in greater growth rates for both modes. Low frequencies had a significant effect on the first mode, but a weak effect on the second mode, whereas high frequencies demonstrated a favorable impact on both the first and second modes. In terms of the growth rate varying with the spanwise wave number, the control rule of the same parameter under different spanwise wave numbers was different, resulting in a complex pattern. In order to obtain the optimal delay effect upon transition and improve the stability of the flow, the parameters of the bi-synthetic jet should be selected as follows: position it upstream, with f1 = 3.56 kHz, f2 = 89.9 kHz, a = 0.009, so that the maximum growth rate of the first mode is reduced by 9.06% and that of the second mode is reduced by 1.28% compared with the uncontrolled state, where flow field analysis revealed a weakening of the twin lattice structure of pressure pulsation.

Abstract The efficacy of steady large-amplitude blowing/suction on instability and transition control for a hypersonic flat plate boundary layer with Mach number 5.86 is investigated systematically. The influence of the blowing/suction flux and amplitude on instability is examined through direct numerical simulation and resolvent analysis. When a relatively small flux is used, the two-dimensional instability critical frequency that distinguishes the promotion/suppression mode effect closely aligns with the synchronisation frequency. For the oblique wave, as the spanwise wavenumber increases, the suppression effects would become weaker and the mode suppression bandwidth diminishes/increases in general in the blowing/suction control. Increasing the blowing/suction flux can effectively broaden the frequency bandwidth of disturbance suppression. The influence of amplitude on disturbance suppression is weak in a scenario of constant flux. To gain a deeper insight into disturbance suppression mechanism, momentum potential theory (MPT) and kinetic energy budget analysis are further employed in analysing disturbance evolution with and without control. When the disturbance is suppressed, the blowing induces the transport of certain acoustic components along the compression wave out of the boundary layer, whereas the suction does not. The velocity fluctuations are derived from the momentum fluctuations of the MPT. Compared with the momentum fluctuations, the evolutions indicated by each component's velocity fluctuations greatly facilitate the investigations of the acoustic nature of the second mode. The rapid variation of disturbance amplitude near the blowing is caused by the oscillations of the acoustic component and phase speed differences between vortical and thermal components. Kinetic energy budget analysis is performed to address the non-parallel effect of the boundary layer introduced by blowing/suction, which tends to suppress disturbances near the blowing. Moreover, viscous effects leading to energy dissipation are identified to be stronger in regions where the boundary layer is rapidly thickening. Finally, it is demonstrated that a flat plate boundary layer transition triggered by a random disturbance can be delayed by a blowing/suction combination control. The resolvent analysis further demonstrates that disturbances with frequencies that dominate the early transition stage are dampened in the controlled base flow.

Active mass injection serves as an effective thermal protection technique by significantly reducing wall heat flux. However, it inherently alters boundary layer stability characteristics, leading to substantial impacts on the laminar-to-turbulent transition process. Crucially, the underlying mechanisms governing how different injected gases modulate flow stability remain unclear. To systematically analyze the effects of different gas injections on flow stability, this study investigates gas-specific mass injection effects by employing a multicomponent Navier-Stokes solver to compute flow fields with air, argon, and nitrogen injections. The influence of mass injection on flow stability was analyzed using linear stability theory, with subsequent differentiation of the distinct effects attributable to various injectant properties. The study demonstrates that mass injection displaces the freestream gas, forming an injection layer near the wall and consequently increasing the boundary layer thickness. Herein, the main boundary layer retains properties similar to the original boundary layer, while the injection layer exhibits significantly reduced temperature and velocity gradients, resulting in decreased wall heat flux and skin friction. Linear stability analysis reveals that while mass injection excites multiple higher-order instability modes, the second mode remains dominant. Notably, mass injection reduces the unstable region of the second mode and significantly decreases the integrated disturbance amplitude, thereby suppressing transition. This stabilizing effect is more pronounced with lighter gases. The differences in injected gas properties are mainly reflected in the viscosity coefficient, thermal conductivity, relative molecular weight, and diffusivity. Among these, the boundary layer thickness is primarily affected by the viscosity coefficient, relative molecular weight, and diffusivity of the injected gas, while the temperature within the boundary layer decreases with increasing thermal conductivity and specific heat capacity of the injected gas. The influence of injected gas properties on flow stability manifests through two distinct pathways: (1) modification of the base flow profile, and (2) alteration of mixed gas properties. Specifically, the transport coefficients (viscosity and diffusivity) of the injected gas primarily affect instability characteristics through Pathway 1, while the specific heat capacity mainly operates via Pathway 2. The relative molecular weight exerts combined effects through both pathways.

This work presents recent advancements in the study of film cooling in hypersonic flows, considering experimental and numerical investigations, with the aim to characterize the wall-cooling performance in different coolant injection and baseflow conditions in a Mach number range 2–7.7. The study explores the mutual interaction between the injected coolant film and the boundary-layer flow, with emphasis on the effects of wall blowing on the boundary-layer characteristics, stability, and transition to turbulence, as well as the effect of transition on wall-cooling performance. Considered flow configurations include cases of effusion cooling in both wall-normal or slightly inclined and wall-parallel blowing, different types of coolant, cases of favorable pressure gradient compared to zero pressure gradient, as well as transpiration cooling cases at different blowing ratios and surface geometries. For the transpiration cooling case, experiments in different hypersonic wind tunnel facilities are presented for flat plate and cone geometries, with coolant injected through C/C porous samples, whereas numerical simulations of modeled porous injection are presented for a flat plate and a blunt cone, showing results for the boundary-layer receptivity with coolant injection and the associated effects on transition and cooling performance. A summary of the main findings is provided along with a critical analysis based on a comparative study to evaluate the effect of each configuration, injection strategy, and key parameters on the boundary-layer flow and the feedback on wall-cooling performance. Conclusions are drawn about potential directions of study for the further development and optimization of the film cooling technique for future hypersonic vehicles.

Hypersonic flight in the atmosphere is associated with high thermal flux impacting the vehicle surface. The nose, leading edges, and some elements of the engine typically require the implementation of highly refractory materials or an active thermal protection system to maintain structural stability during the vehicle mission. Carbon–carbon (C–C) composites are commonly considered for the application thanks to their unique thermal and mechanical properties. However, C–C composites’ ablation and oxidation under long cruise flights at high speeds (Mach number > 5) are the limiting factors for their application. In this paper, the results of an experimental study of C–C composite thermal ablation and oxidation with test article surface temperatures up to 2000 K are presented. The tests were performed under atmospheric conditions and hypersonic flow in the ND_ArcJet facility at the University of Notre Dame. The test articles were preheated with CW laser radiation and then exposed to M = 6 flow at stagnation pressures up to 14 bar. It was found that C–C composite oxidation and mechanical erosion rates are significantly increased in hypersonic airflow compared to those at ambient conditions and nitrogen M = 6 flow. Compared to atmospheric air, mass loss occurred at a rate of 1.5 orders of magnitude faster for M = 6 airflow. During high-speed flow conditions, rapid chemical oxidation and the mechanical destruction of weakened C-fibers likely cause the accelerated degradation of C–C composite material. In this study, a post-mortem microscopic analysis of the morphology of the C–C surface is used to explain the physical processes of the material destruction.

Aiming to efficiently simulate the transient process of transpiration cooling with phase change and reveal the convection mechanism between fluid and porous media particles in a continuum scale, a new two-phase mixture model is developed by incorporating the local thermal non-equilibrium effect. Considering the low-pressure and high overload working conditions of hypersonic flying, the heat and mass transfer induced by capillary and inertial body forces are analyzed for sub-cooled, saturated and super-heated states of water coolant under varying saturation pressures. After the validation of the model, transient simulations for different external factors, including spatially-varied heat flux, coolant mass flux, time-dependent external pressure and aircraft acceleration are conducted. The results show that the vapor blockage patterns at the outlet are highly dependent on the injection mass flux value and the external pressure, and the reduced saturation temperature at low external pressure leads to early boiling off and vapor blockage. The motion of flying has a large influence on the cooling effect, as the inertial force could change the flow pattern of the fluid inside significantly. The comparison of the results from 2-D and 3-D simulations suggests that 3-D simulation shall be conducted for practical application of transpiration cooling, as the thermal protection efficiency may be overestimated by the 2-D results due to the assumption of an infinite width length of the porous plate.

Supersonic film cooling serves as an effective thermal protection method for hypersonic vehicles. This study performs direct numerical simulations of supersonic film injection into a hypersonic turbulent boundary layer (Mach 6), investigating the influence of film Mach number Mac on cooling performance and flow characteristics. Mean flow field data reveal that increasing Mac extends absolute effective cooling length, but is less cost-effective in terms of effective cooling length per unit film mass flow rate. Further analysis reveals that the diminished cost-effectiveness results from earlier boundary layer transition relative to the effective cooling length, which enhances turbulent mixing in the effective cooling region. Turbulent kinetic energy (TKE) evolution demonstrates that in the wall-jet region, the turbulent intensity at the lowest Mac (Mac=2) displays a one-peak pattern that differs significantly from the two-peak distributions under higher Mac conditions, which stems from the lowest film mass flow rate that causes rapid complete mixing. Bypass transition occurs in the boundary layer beneath the film. Although the absolute location of the transition onset is insensitive to Mac, the downstream turbulence development shows altered characteristics with increasing Mac, as evidenced by the peak TKE levels and spanwise length scales of turbulent structures. Interestingly, greater scale disparity between the turbulent structures near the wall and in the mixing layer is observed as Mac increases, which is attributed to the reduced spanwise scales of near-wall streaks. The increase in Mac results in decelerated turbulence development, extending the absolute streamwise distance required for the film to be completely dissipated.

Chemical kinetic schemes have been developed for hypersonic flows with ablative carbon and carbonaceous components; however, experimental data for the validation of these schemes are limited. Therefore, in this study, we use a carbon-ablation chemical kinetics model to identify changes in the refractive index field near a hypersonic vehicle as well as other experimentally observable metrics that can be detected in future experiments conducted in a high-enthalpy wind tunnel. The combined use of a zero-dimensional kinetics model, two-dimensional hypersonic flow simulations, and a refractive index model confirm that the level of carbon present in Mach 24 hypersonic flow significantly affects the refractive index, electron density, and shock wave location. All three metrics can be used for an analysis of ablation products in a high-enthalpy wind tunnel.

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Transition of hypersonic boundary layer flow is an important problem in aerodynamic research. To obtain the natural transition position using the flow stability analysis method, it is necessary to analysis the neutral curve and the perturbation growth rule. Accurate calculation of the basic flow is the basis and prerequisite for stability analysis. This paper proposes that surface radiation should be considered in the analysis of hypersonic boundary layer flow stability due to the high temperatures at the vehicle surface caused by aerodynamic heating. In order to obtain the influence of surface radiation on the flow stability, we assume that the surface satisfies the energy balance and the gas satisfies calorimetrically perfect gas properties. To simplify the analysis, gas radiation has been neglected. The boundary layer analysis of a hypersonic flat plate with and without radiation has been performed by a combination of flow stability theory and direct numerical simulation. It is found that high temperature radiation on the surface causes energy to transfer outwards. The results showed that surface temperature decreases along the streamwise direction and is no longer an adiabatic surface temperature. Instability interval decreases and the maximum growth rate increases. The range of the instability region expressed in dimensionless frequency values moves to lower frequency as the Mach number increases. Maximum growth rate shows a tendency to increase and then decrease as the Mach number increases.

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An implicit coupling framework between hypersonic nonequilibrium flows and material thermal response is proposed for the numerical simulation of ablative thermal protection materials during its flight trajectory. Charring ablative materials, when subjected to aerodynamic heating from hypersonic flows, undergo complex processes such as ablation and pyrolysis, involving heterogeneous and homogeneous chemical reactions. These multi-physical phenomena are simulated by a multicomponent material thermal response (MTR) solver that takes into account the complexity of component of pyrolysis gases. The species concentrations are calculated to improve the accuracy of transport and thermophysical parameters of pyrolysis gases. The MTR solver implements implicit time integration on finite difference discretization form to achieve higher efficiency. The numerical solutions of hypersonic flows and material thermal response are coupled through a gas-surface interaction interface based on surface mass and energy balance on the ablating surface. The coupled simulation employs the dual time-step technique, which introduces pseudo time step to improve temporal accuracy. The explicit coupling mechanism updates the interfacial quantities at physical time steps, which achieves higher computational efficiency, but introduces time discretization errors and numerical oscillations of interfacial quantities. In contrast, the implicit coupling mechanism updates the interfacial quantities at pseudo time steps, which reduces the temporal discretization error and suppresses numerical oscillations, but is less efficient. In addition, a simplified ablation boundary based on steady-state ablation assumption or radiation-equilibrium assumption is proposed to approximate solid heat conduction without coupling the MTR solution, providing quasi-steady flow solutions in the presence of ablation.

: The aim of this work is to investigate the effects of thermochemical non-equilibrium in the shock layer and boundary layer of hypersonic flows past blunt bodies entering planetary atmospheres. Ablative thermal protection systems change the mixture composition of the boundary layer with significant impact on the surface heat flux. In this context, a vibrationally resolved state-to-state approach is employed in order to understand the effect of non-equilibrium of the molecular energy level population on the surface heat flux.

Experiments on hypersonic boundary-layer instability of a fin–cone configuration were conducted in a Φ 0.5 m Mach 6 Ludwieg tube tunnel. Infrared thermography and high-frequency pressure sensors were used to measure the transition front and instability waves under four different nose bluntness conditions. On the leeward surface, transition is delayed near the centerline due to expansion waves generated by the double-cone structure. The region close to the corner is strongly influenced by the horseshoe vortex, whereas instability waves around 110 kHz manifest as the flow moves away from it. In contrast, transition on the windward surface occurs earlier and broadband high-frequency instability waves of 160–300 kHz are present near the corner. Increasing nose bluntness strongly suppresses transition away from the fin root, especially near the centerline and on the fin-off cone side, but has a relatively limited impact on the shock-interaction regions near the fin–cone corner. Transition on the fin surface remains insensitive to nose bluntness variations. This work elucidates the distinct transition behaviors across different regions of a complex fin–cone configuration and their differential responses to nose bluntness, providing valuable insights for the aerodynamic design and transition prediction of hypersonic vehicles.

In the present study, the response of a hypersonic turbulent boundary layer at an inflow of Ma∞ = 6 and Re∞ = 16·106 1/m to a smooth and rough surface along a sharp cone is examined. The model consisted of three segments with exchangeable parts to consider smooth and rough surfaces with a roughness topology of square bar elements with a nominal wavelength of four times the height of the elements. In selected regions of interest, the flow field was measured by particle image velocimetry (PIV) which enabled analysis of mean velocity fields and Reynolds stresses. Van Driest transformed smooth wall mean velocity profiles showed the expected incompressible behavior and compared well to previous investigations. A combination of an integral and fitting approach is discussed to enable inner scaling of the rough wall profiles, which showed the expected shift below the smooth wall profile. The smooth wall turbulence profiles from PIV agreed to artificially filtered DNS in case of the streamwise component. Turbulence profiles above the smooth and rough wall agreed to within measurement accuracies. Additionally, two−point correlations were used to investigate turbulent structures above the smooth and rough wall. Both, length scales and orientations of the correlations, showed high level of agreement between smooth and rough walls, with only differences close to the wall. Furthermore, uniform momentum zones could be identified with similar behavior along both smooth and rough walls. Information from turbulence data support outer layer similarity, whereas mean velocity profiles show an increase in Coles wake parameter for the rough wall data. This might be influenced by transitional roughness effects.

In nonequilibrium slip and jump conditions, normal gas velocity and temperature gradients are used to calculate the gas slip velocity and temperature at the surface, respectively. Gökçen et al. (Computational fluid dynamics near the continuum limit, AIAA Paper No. 87-1115, 1987, and Gökçen and MacCormack, Nonequilibrium effects for hypersonic transitional flows using continuum approach, AIAA Paper No. 89-0461, 1989) stated that the tangential velocity and temperature of the gas molecules before a collision with the surface could be interpreted as the macroscopic tangential velocity and temperature of the gas molecules at the so-called near-wall distances auλ and aTλT away from the surface, respectively. The coefficients au and aT are the order of unity. In the present work, new forms of the slip and jump conditions are proposed by modifying the Gökçen slip and jump conditions to include the coefficients (au, aT). Numerical investigations are comprehensively conducted to determine the numerically proper values (au, aT) for the hypersonic rarefied gas flows. Cases such as the circular cylinder in cross-flow and sharp and blunted leading edge wedge are considered in the present work, with nitrogen as the working gas. The simulation results show the significant effects of the coefficients (au, aT) on the accuracy of the slip velocity and surface gas temperature predictions, and the values of au = 1.2 and aT = 1.1 show good agreement with the direct simulation Monte Carlo data.

Hypersonic near-continuum flow over a flat plate with micro-scale roughness is studied using the kinetic direct simulation Monte Carlo method on roughness module configurations with different relative roughness (h) values and roughness densities (RN) under a matrix of freestream parameters (Mach number Ma∞, Reynolds number Re∞, temperature T∞, and Knudsen number Kn∞). An open-source Stochastic PArallel Rarefied-gas Time-accurate Analyzer code, which enables Cartesian grid adaption and efficient parallelization, is utilized for the rough-plate flow simulations. Flowfield analysis reveals that the local patterns inside the roughness modules evolve starting from closed (two vortices) via transitional ultimately to open (one vortex) by an increase in h, with co-existing shrinkage of high-density zones and attenuation of density peaks. The surface quantities are significantly influenced by the flowfield characteristics, and a local association between the peak heat flux and the peak pressure is identified. Non-dimensional peak heating and pressure correlation laws for the local peak heat flux and pressure coefficients in terms of two length-scale transformations are proposed, enabling the capture of local heating and pressure extrema on rough plates with varying h and RN conditions under different Ma∞, Re∞, and T∞ parameter values. The peak heat flux and pressure coefficients can be described by analogous correlating equations expressed by first-order-polynomial or power functions. An increase in the rarefaction degree (Kn∞) deviating from the near-continuum regime causes the correlation laws to fail.

An experimental study on controlling hypersonic boundary layer transition using a surface arc plasma actuation array was conducted. First, base boundary transition characteristics were analyzed in virtue of various sensors and high-speed schlieren, and the transition criterion was established based on the critical value of schlieren spatial power spectral density resolution. Then, the influence of three different actuation frequencies (8, 34, and 55 kHz) was studied based on linear stability theory analysis. The impact of different actuation frequencies on the dominant unstable waves in the boundary layer was analyzed, and a transition criterion under the regulation of plasma actuation was proposed. Finally, the proper orthogonal decomposition method was used to analyze the influence of different actuation frequencies on the unstable characteristic structures. Based on the research, the efficacy of plasma actuation array in promoting transition is verified, the corresponding regulation mechanism is summarized, and transition regulation mechanization is refined.

Cavities on the surfaces of hypersonic vehicles cannot be avoided easily. Moreover, they can trigger boundary layer transition under certain conditions. However, little progress has been reported on boundary layer transitions induced by a three-dimensional (3-D) shallow cavity. In this study, transitions induced by six 3-D shallow cavities with the same length–depth ratios of 20 and different widths of 0.25, 0.5, 0.75, 1.0, and 1.5 times the baseline width (W), as well as infinite width, were investigated. A direct numerical simulation was conducted using the Roe scheme with 4th-order minimum dispersion and controllable dissipation, and weighted essentially non-oscillatory reconstruction, based on our in-house code: Unsteady NavIer–STokes equations solver. Cavity width was observed to have non-monotonic influences on transition. Both the 0.25 and 1.0 W cavities could induce transition constantly. Moreover, flow was maintained as laminar past the 1.5 W and InfW cavities. For the 0.5 and 0.75 W cavities, intermittent transition was observed with different intermittency factors. The intermittent transition phenomenon was determined to be caused by the periodic increase and decrease in the adverse pressure gradient (APG) in the front part of cavity. Notably, the recirculation with a synchronic size change was the origin of the APG oscillation.

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The large number of transitions involved in vibration–vibration–translation (VVT) transitions during molecule–molecule collisions in the state-to-state (StS) simulation significantly increases computational cost and limits the multidimensional application. Two methods, multiquantum restrictions and simplifying VVT transitions to vibration–translation (VTm) transitions, are proposed to reduce computational consumption. This study is a continuation of a previous work [Guo et al., “Investigation of high enthalpy thermochemical nonequilibrium flow over spheres,” Phys. Fluids 36, 016122 (2024)]. It aims to systematically analyze the effects of VVT and VTm transitions on the flowfield characteristics and heat flux distributions of high-enthalpy nitrogen flow over a sphere. The numerical results show that the StS VVT and VTm transitions predict nearly identical shock standoff distances and stagnation-point heat fluxes, which agree with experimental data. Incorporating VVT transitions yields minimal differences in the predictions of translational temperature, vibrational temperature, and N mass fraction compared to VTm processes. Near the wall, the VVT and VTm transitions yield almost identical state populations, with gradually pronounced non-Boltzmann distributions. Therefore, the VVT transitions can be effectively reduced to VTm transitions while maintaining essential accuracy for predicting aerodynamic heating.

A high-speed compressible solver capable of solving detailed chemical reaction mechanisms is developed by coupling the open-source computational fluid dynamic toolbox OpenFOAM® and Cantera 2.5.0. A sonic hydrogen jet discharging from a circular injector into a high enthalpy supersonic crossflow over a flat plate is selected as a test case for the developed solver. The incoming boundary layer is laminar, and an adverse pressure gradient-induced transition is expected due to transverse injection. The test case is selected to serve two purposes. First, to validate the developed solver. Second, to inspect the capability of Reynolds-Averaged Navier–Stokes (RANS) in predicting the flame characteristics in high-speed flows involving laminar to turbulent transition. The present study features three-dimensional RANS simulations with Shear Stress Transport (SST) k–ω and Langtry–Menter SST k–ω turbulence models, with three values of inlet turbulent intensity: I = 0.5, 1, and 2. Analysis showed that laminar to turbulent transition plays a significant role in the resulting flame structure. A fully turbulent SST k–ω model showed several discrepancies from the experiment, especially near the boundary layer. On the other hand, the Langtry–Menter SST k–ω model predicts transition onset and hence the flame structures accurately. Furthermore, the transition onset and the flame structure strongly depend on I. The low-velocity recirculation regions near the injector aid in flame stabilization upstream of the injector. At the same time, the horseshoe vortex dictates the flame spread in a spanwise direction. The reflected shock–boundary layer interaction helps in flame stabilization downstream of the injector.

Thermal protection is required for vehicles entering planetary atmospheres to protect against the severe heating loads experienced. Characterization of candidate materials is often done utilizing plasma or arc-jet facilities, which provide steady-state testing of the thermal environments experienced during hypersonic flight, but do not correctly simulate hypersonic flowfields. Conversely, impulse facilities can reproduce flight velocities and enthalpies but have extremely short test times, prohibiting testing of thermal response. To better understand how these materials interact with hypersonic flows, experiments were conducted at the X2 expansion tunnel at the University of Queensland. Preheated strips of carbon-carbon and silicon carbide-coated carbon-carbon were mounted in a two-dimensional compression wedge and tested in Earth entry flow, marking the first time silicon carbide has been investigated in this facility. Calibrated spectral measurements were obtained in the near-stagnation and expansion regions for surface temperatures from 1900 K to 2600 K. Cyanogen emissions dominated while atomic silicon and dicarbon were also observed. Emissions for both materials displayed a similar increase near the wall, while emissions for silicon carbide-coated samples displayed a distinct rise downstream of the shock, which suggests a higher concentration of ablative species resulting from a higher ablation rate.

During hypersonic flight, air reacts with a planetary re-entry vehicle's thermal protection system (TPS), creating reaction products that deplete the TPS. Reliable assessment of TPS performance depends on accurate ablation models. New finite-rate gas-surface chemistry models are advancing state-of-the-art in TPS ablation modeling, but model reductions that omit chemical species and reactions may be necessary in some cases for computational tractability. This work develops hybrid physics-based and data-driven enrichments to improve the predictive capability and quantify uncertainties in such low-fidelity models while maintaining computational tractability. We focus on discrepancies in predicted carbon monoxide production that arise because the low-fidelity model tracks only a subset of reactions. To address this, we embed targeted enrichments into the low-fidelity model to capture the influence of omitted reactions. Numerical results show that the hybrid enrichments significantly improve predictive accuracy while requiring the addition of only three reactions.

Surface ablation significantly affects the distribution of plasma in high-speed flow and the characteristics of their interaction with electromagnetic fields. Considering the mechanism of ablation and ejection on the surface of hypersonic vehicle, the participation of ablation products in the plasma generation process in the flow field, the conduction mechanism of mixed ionized gas containing alkali metal and the electromagnetic dynamics mechanism, the coupled calculation method of high-speed flow/plasma/electromagnetic field with alkali metal ablation is established by solving the three-dimensional thermochemical non-equilibrium flow governing equation with electromagnetic source term, the electric field Poisson equation and the magnetic vector Poisson equation. Combined with the common ablation and pyrolysis process of carbon-carbon materials and silicon-based phenolic resin materials, the mechanism and law of the interaction between surface ablation and electromagnetic field on the hypersonic plasma sheath under various conditions are systematically studied. The results show that the ablation effect affects the plasma distribution in the flow field, which is affected by the ablation mass ejection rate and the mass proportion of alkali metal. When the alkali metal content is high, the alkali metal ionization reaction is dominant, and the electron number density can increase by 1 ~ 2 orders of magnitude. The influence of different materials on plasma is different. The mass ejector ratio of silicon-based phenolic resin is larger, and the molar concentration of CO+ and C+ produced by ionization is close to that of NO+ and O2+, which can not be ignored. Alkali metals in ablative materials can significantly improve the control effect of magnetohydrodynamics. With the increase of the proportion of alkali metals, the coupling effect of electromagnetic fields increases, and the relationship between them is nonlinear. When the speed is low, the ionization degree of air itself is low and the coupling effect of electromagnetic field is weak. But the efficiency of "improving the electromagnetic effect by ablation of alkali metal" is higher.

Abstract In this paper, we study the receptivity of non-modal perturbations in hypersonic boundary layers over a blunt wedge subject to free stream vortical, entropy and acoustic perturbations. Due to the absence of the Mack-mode instability and the rather weak growth of the entropy-layer instability within the domain under consideration, the non-modal perturbation is considered as the dominant factor triggering laminar–turbulent transition. This is a highly intricate problem, given the complexities arising from the presence of the bow shock, the entropy layer and their interactions with oncoming disturbances. To tackle this challenge, we develop a highly efficient numerical tool, the shock-fitting harmonic linearised Navier–Stokes (SF-HLNS) approach, which offers a comprehensive investigation on the dependence of the receptivity efficiency on the nose bluntness and properties of the free stream forcing. The numerical findings suggest that the non-modal perturbations are more susceptible to free stream acoustic and entropy perturbations compared with the vortical perturbations, with the optimal spanwise length scale being comparable with the downstream boundary-layer thickness. Notably, as the nose bluntness increases, the receptivity to the acoustic and entropy perturbations intensifies, reflecting the transition reversal phenomenon observed experimentally in configurations with relatively large bluntness. In contrast, the receptivity to free stream vortical perturbations weakens with increasing bluntness. Additionally, through the SF-HLNS calculations, we examine the credibility of the optimal growth theory (OGT) on describing the evolution of non-modal perturbations. While the OGT is able to predict the overall streaky structure in the downstream region, its accuracy in predicting the early-stage evolution and the energy amplification proves to be unreliable. Given its high-efficiency and high-accuracy nature, the SF-HLNS approach shows great potential as a valuable tool for conducting future research on hypersonic blunt-body boundary-layer transition.

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Exploring the influence of material-coupled hypersonic plasma flow fields on radio-frequency (RF) communication environments holds significant engineering value. A three-dimensional hypersonic flow field is constructed based on the Navier–Stokes equations coupled with a 7-species, 18-reaction thermochemical model. The simulation considered multi-species transport, ionization reactions, and thermochemical nonequilibrium effects. The predicted axial distribution of peak electron density agrees well with NASA flight test data and previously published results, confirming the accuracy and applicability of the proposed model. The spatiotemporal evolution of the plasma sheath is systematically investigated, with a focus on three representative ablative materials: carbon–carbon composites (C–C), silicon-based thermal protection materials (Si-Phenolic Resin, Si-PR), and aluminum matrix composites (AMCs). Based on this, distributions of electron number density, collision frequency, and flow field temperature are obtained at two representative reentry altitudes, revealing the coupled effects of ablation composition, thermal radiation, and product transport on sheath structure. The results indicate that along the axial direction of the vehicle, the C–C material exhibits the highest ionization near the stagnation point, Si-PR maintains strong ionization activity in the midsection, and AMCs show favorable thermal diffusion capability in the rear region. Based on the quantitative distribution characteristics of electron density and collision frequency, a region-specific material adaptation strategy is proposed. This strategy provides theoretical guidance for thermal protection material design and enhances the understanding of material selection and layout optimization in the electromagnetic environment design of reentry vehicles.

Legacy and modern-day ablation codes typically assume equilibrium pyrolysis gas chemistry. Yet, experimental data suggest that speciation from resin decomposition is far from equilibrium. A thermal and chemical kinetic study was performed on pyrolysis gas advection through a porous char, using the Theoretical Ablative Composite for Open Testing (TACOT) as a demonstrator material. The finite-element tool SIERRA/Aria simulated the ablation of TACOT under various conditions. Temperature and phenolic decomposition rates generated from Aria were applied as inputs to a simulated network of perfectly stirred reactors (PSRs) in the chemical solver Cantera. A high-fidelity combustion mechanism computed the gas composition and thermal properties of the advecting pyrolyzate. The results indicate that pyrolysis gases do not rapidly achieve chemical equilibrium while traveling through the simulated material. Instead, a highly chemically reactive zone exists in the ablator between 1400 and 2500 K, wherein the modeled pyrolysis gases transition from a chemically frozen state to chemical equilibrium. These finite-rate results demonstrate a significant departure in computed pyrolysis gas properties from those derived from equilibrium solvers. Under the same conditions, finite-rate-derived gas is estimated to provide up to 50% less heat absorption than equilibrium-derived gas. This discrepancy suggests that nonequilibrium pyrolysis gas chemistry could substantially impact ablator material response models.

Hypersonic boundary layer transition is critical to the design of all hypersonic vehicles due to its effect on the heat transfer into the vehicle surface and potential drag enhancement or reduction during reentry. Boundary layer transition and boundary layer stability analysis under hypersonic conditions has been studied for decades, yet there is ample room for improved accuracy and further investigations into the relevant phenomena. In this work, we present a recent implementation of chemical equilibrium, finite-rate chemistry, and thermochemical nonequilibrium capabilities into LASTRAC, an existing well-established boundary-layer stability analysis code. Verification against existing numerical results in the literature are presented. LASTRAC was previously able to address calorically perfect flows. By using solutions of the Parabolized Stability Equations (PSE) with chemical and thermal nonequilibrium, we are able to investigate the effects of chemical and thermal nonequilibrium on a variety of phenomena including stationary crossflow instability on a swept wing and 2 mode instabilities over a wedge.

For increasing understanding of fundamental hypersonic phenomena, the flight test program, named MF-1, is to gather fundamental scientific and engineering data on the physics and technologies critical to future operational hypersonic flight with low-cost flight test platform, which is built on the retrofitted rockets. The MF-1 program is a hypersonic flight test program executed by China Aerodynamic Research and Development Center (CARDC). The MF-1 flight flew in December 2015. The flight focuses primarily on integration of instrumentation on the test vehicle, with application to boundary layer transition and shock interaction experiments. The MF-1 payload consists of a blunted 7°half angle cone, a cylinder and 33° flare configuration. The payload was boosted to Mach 5.32 utilizing a solid-rocket booster without control for the whole flight. The flight was fully successful, and measured transition under supersonic and hypersonic conditions. The heat flux data were given by the three-dimensional thermal identification method to discriminate transition zone. The preliminary analysis shows that the real-time flight data obtained by MF-1 are reliable and can be used to validate the transition predicting model and software. The results show that the existing model is able to predict the transition location of cone at a small angle-of-attack for supersonic or hypersonic flow. This paper describes the MF-1 mission and some general conclusions derived from the experiment.

The ablation product carbon dioxide (CO2) around the hypersonic vehicle may become an important source of radiation. Accurate simulation of the chemical reaction process between CO2 and air is of great significance for the design of thermal protection systems. A combination of ground experiments and numerical simulations is used in this study to investigate the reaction process of CO2–air mixtures with different proportions behind shock waves with velocities ranging from 3.6 to 4.1 km/s. Flow parameters are simulated based on the two-temperature (2-T) model and the Lee and Park chemical reaction mechanism, and the laser absorption spectroscopy measurements are carried out on the Φ800 mm shock tube at the Chinese Academy of Sciences, Institute of Mechanics, to obtain time-resolved post-shock gas temperatures and partial pressures of CO2 and carbon monoxide (CO). Measurement results indicate that the post-shock temperature is much higher than the equilibrium value when it is close to the shock front and then gradually develops towards equilibrium. The partial pressure of CO2 shows the same trend as temperature. Comparisons between experimental results and calculated results show that the used model can accurately predict the post-shock temperature. When the initial CO2 volume fraction is below 50%, the model underestimates the CO2 dissociation rate, meaning that the actual CO2 content is lower than the calculated value and the CO content is higher. Conversely, when the CO2 volume fraction is greater than or equal to 50%, the prediction of the CO2 dissociation rate tends to be overestimated.

The hypersonic flow over a compression ramp is investigated by utilizing direct numerical simulation (DNS) and various stability analysis tools. The free-stream Mach number and Reynolds number based on the length of the flat plate are 8.0 and 3.9×105, respectively. Global stability analysis is applied to confirm the weekly unstable nature of the current flow condition. As a result of the low growth rate, this case is believed to be more susceptible to convective instability than intrinsic instability. Subsequently, across a wide range of frequencies and a globally stable wavelength, resolvent analysis is utilized to investigate the response of two-dimensional base flow to external disturbances. It reveals that the optimal response to upstream disturbances located adjacent to the leading edge manifests in the form of streamwise streaks, which result from transient growth in the flat-plate boundary layer. Downstream of reattachment, the Mack second mode and low-frequency streaks as a manifestation of Görtler instability coexist within the boundary layers. Further downstream, the amalgamation of the amplification of Mack's second mode with the sinuous and varicose breakdown of streaks disrupts the boundary layers via the ejection-sweep motion, resulting in the creation of a strong localized vorticity region and contributing to the concentration of vorticity within the boundary layers. This kicks off the vortex roll-up process, which results in the formation of hairpin vortices, and eventually leads to the breakdown process.

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The influence of flow parameters and nose radius on the location of laminar–turbulent transition is investigated. The model is an ogive-conical body of revolution having half angle about 9°. Experiments were conducted in a shock tunnel at Mach number 5. The transition location was diagnosed by the heat transfer rate distribution determined with the aid of luminescent temperature converters. It is shown that transition reversal can occur either (a) in the absence of turbulent wedges or (b) at a constant level of freestream disturbances. Both increasing and decreasing branches of Re∞,t (Re∞,R) dependency were observed at constant nose radius while varying only the unit Reynolds number.

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An experimental investigation of distributed sand-grain surface roughness effects on boundary-layer transition and convective heating has been performed. Two representative entry vehicle geometries, a spherical-cap aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic wind tunnel models of each geometry were fabricated with various roughness heights to simulate an ablated thermal protection system. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flowfield simulations.

Abstract Instability measurements of an axisymmetric, laminar separation bubble were made over a sharp cone-cylinder-flare with a $12^{\circ }$ flare angle under hypersonic quiet flow. Two distinct instabilities were identified: Mack's second mode (which peaked between 190 and 290 kHz) and the shear-layer instability in the same frequency band as Mack's first mode (observed between 50 and 150 kHz). Both instabilities were measured with surface pressure sensors and were captured with high-speed schlieren. Linear stability analysis results agreed well with these measured instabilities in terms of both peak frequencies and amplification rates. Lower-frequency fluctuations were also noted in the schlieren data. Bicoherence analysis revealed nonlinear phase-locking between the shear-layer and second-mode instabilities. For the first time in axisymmetric, low-disturbance flow, naturally generated intermittent turbulent spots were observed in the reattached boundary layer. These spots appeared to evolve from shear-layer-instability wave packets convecting downstream. This work presents novel experimental evidence of the hypersonic shear-layer instability contributing directly to transition onset for an axisymmetric model.

This work deals with the numerical solution of hypersonic flow of viscous fluid over a compressible ramp. The solved case involves very important and complicated phenomena such as the interaction of the shock wave with the boundary layer or the transition from a laminar to a turbulent state. This type of problem is very important as it is often found on re-entry vehicles, engine intakes, system and sub-system junctions, etc. Turbulent flow is modeled by the system of averaged Navier–Stokes equations, which is completed by the explicit algebraic model of Reynolds stresses (EARSM model) and further enhanced by the algebraic model of bypass transition. The numerical solution is obtained by the finite volume method based on the rotated-hybrid Riemann solver and explicit multistage Runge–Kutta method. The numerical solution is then compared with the results of a direct numerical simulation.

Ogive cylinders are representative designs of hypersonic forebodies and harbor several flow complexities. This work evaluates the effects of nose bluntness of ogive-cylinder forebodies on their laminar, transitional, and turbulent boundary layers (BLs). Laminar simulations indicate that the thickness of BL and entropy layer (EL) increases with nose bluntness, with the latter displaying a stronger sensitivity to nose radius due to the increased strength and curvature of the leading-edge shock system. Linear perturbation analysis identifies the presence of the second mode, the first mode, and EL modes with increasing bluntness. While the former two are prevalent over a significant streamwise extent, EL modes are limited to the upstream region, well within the swallowing length. Using direct numerical simulations, it is found that transition over the sharpest nose design occurs through the modal fundamental resonance route, driven by second-mode instabilities. Due to weakened BL modes, the blunter geometries undergo transition further downstream in the low-amplitude freestream perturbation environment tested in this study. The transition mechanism in the blunter geometries is intermittent in nature and appears to be driven by turbulent spots generated from streamwise vortices reminiscent of Klebanoff modes.

Hypersonic flow over axisymmetric bodies with four evenly spaced fins was investigated computationally. The test conditions encompassed a freestream Mach number of 5.9 and Reynolds numbers based on body length of [Formula: see text] and [Formula: see text], resulting in shock-wave/boundary-layer interactions in the presence of laminar–turbulent transition. The test article was the High-Speed Army Reference Vehicle, a set of generic configurations for research on missilelike bodies. Steady-flow computations employing the Menter Shear Stress Transport turbulence model and one-equation transition model were performed for angles of attack between 0.0 and 7.0-deg. The predicted heat flux was compared with experimental measurements obtained at Texas A&M University. As the angle of attack was increased, complex vortex interactions and shock-wave/boundary-layer interactions were observed in the region between the fins.

Accurate prediction of laminar-to-turbulent flow transition is challenging because of its complex nature. In this study, we use a combination of linear stability theory (LST) and direct numerical simulations (DNSs) to examine perturbation growth on ogive-cylinder forebodies, relevant to modern hypersonic vehicles, focusing on the effect of nose bluntness. The spatial evolution of disturbances is first examined through the solution of a linear eigenvalue problem. As the ogive nose bluntness is increased, the first mode, which was unstable downstream for the sharp nose, moves upstream, and its oblique component displays higher instability. The LST predictions are used to inform DNS to examine the effect of amplitude on perturbation growth and modal characteristics. For small 2D perturbations, second-mode growth with DNS yields similar results as LST for all bluntness cases examined, suggesting that the streamwise gradients inherent to ogive forebodies have relatively little effect on this behavior. When the forcing is azimuthally localized, the oblique nature associated with low-frequency perturbations becomes apparent, analogous to the first mode. These are, however, inhibited when the forcing is azimuthally coherent. Furthermore, at higher amplitudes, nonlinear interactions appear among unstable modes, providing insights into the different interactions governing transition in such nose geometries.

Abstract The attachment-line boundary layer is critical in hypersonic flows because of its significant impact on heat transfer and aerodynamic performance. In this study, high-fidelity numerical simulations are conducted to analyse the subcritical roughness-induced laminar–turbulent transition at the leading-edge attachment-line boundary layer of a blunt swept body under hypersonic conditions. This simulation represents a significant advancement by successfully reproducing the complete leading-edge contamination process induced by a surface roughness element in a realistic configuration, thereby providing previously unattainable insights. Two roughness elements of different heights are examined. For the lower-height roughness element, additional unsteady perturbations are required to trigger a transition in the wake, suggesting that the flow field around the roughness element acts as a perturbation amplifier for upstream perturbations. Conversely, a higher roughness element can independently induce the transition. A low-frequency absolute instability is detected behind the roughness, leading to the formation of streaks. The secondary instabilities of these streaks are identified as the direct cause of the final transition.

Structural morphing is regarded as a potential solution for future space transportation systems, necessitating aircraft capable of extensive altitude ranges, wide speed ranges, and multi-regime operations. The investigation of laminar-turbulent transition incorporating morphing effects aims to compensate for the lack of physical mechanisms in aircraft designs. This paper particularly focuses on the impact of a continuously extending or retracting wall in the chordwise direction on the evolution of inviscid Mack modes in a Mach 5.92 hypersonic boundary layer. Using the high-Reynolds-number asymptotic technique, a model is developed to quantify the morphing effect by an amplification factor, which includes both the scattering effect at the rigid-morphing junction and the successive modification of the Mack growth rate by the morphing velocity in the downstream region. Such a model enables us to conduct a comprehensive investigation across a wide parameter space defined by morphing speed and flow parameters. A critical frequency is identified near the most unstable second-mode frequency. For an extending wall, the oncoming Mack modes are suppressed for frequencies above this critical frequency and enhanced for frequencies below it. Conversely, the retracting-wall effect exhibits an inverse impact on the Mack modes. To validate the accuracy of the asymptotic predictions, the harmonic linearized Navier–Stokes calculations are performed, resulting in favorable agreements. This research sheds light on the complex interplay between wall morphing and the Mack-mode evolution, offering informative contributions to the understanding of the aerodynamic behaviors in hypersonic boundary layers.

The assumption of axial symmetry for hypersonic flows over conically shaped geometries is ubiquitous in both experiments and numerical simulations. Yet depending on the free stream conditions, many of these flows are unsteady and their transition from laminar to turbulent is a three-dimensional phenomena. Combining triple deck theory/linear stability analysis with the kinetic direct simulation Monte Carlo method we analyze the azimuthal eigenmodes of flows over single and double-cone configurations. For Mach 16 flows we find that the strongest amplification rate occurs for the non-axisymmetric azimuthal wavenumber of n = 1. This occurs in regions quite close to the tip of the cone due to the proximity of the conical shock to the viscous shear layer where non-axisymmetric modes are amplified through linear mechanisms. Comparison of triple deck linear stability predictions shows that in addition to the azimuthal wavenumber, both the temporal content and amplification rate of these non-axisymmetric disturbances agree well with the time accurate DSMC flowfield. In addition to the loss of axial symmetry observed at the conical shock, the effect of axial symmetry assumptions on the more complex shock-shock and shock-boundary layer interactions of a flow over a double cone are considered. The results for the separation region show that axisymmetric and three dimensional simulations differ in almost all of the main flow structures. Three dimensional flowfields result in a smaller separation bubble with weaker shocks and three dimensional effects were manifest in the variation in surface parameters in the azimuthal direction as well. Interestingly, the DSMC simulations show that the loss of axial symmetry in the separation region, begins near the cone tip.

Abstract The conventional $\textrm{e}^N$ laminar-to-turbulent transition-prediction method focuses on the relative growth rate, called the $N$ factor, and neglects receptivity. To improve predictions, Mack (1977) proposed the amplitude method to incorporate receptivity, nonlinear effects and broadband characteristics. Currently, the lack of accurate receptivity coefficients, estimates of initial disturbance amplitudes at the lower-branch neutral position, referred to as branch I (where the imaginary part of the spatial wavenumber is zero), hinders the application of the amplitude method. Although experimental- and numerical-receptivity analyses have been conducted previously, they rely on correlations or indirect approaches. For the purpose of direct evaluation, this study applies bi-orthogonal decomposition to direct numerical simulation (DNS) data of a hypersonic boundary layer over a blunt cone, extracting initial amplitudes of instability modes. The decomposition framework incorporates both boundary-layer and entropy-layer modes, enabling direct evaluation of receptivity coefficients at branch I. The decomposed modal amplitudes show reduced multimode interference and the receptivity coefficients have been computed to have fewer oscillations. With an overall greater magnitude, the receptivity coefficients suggest a possible earlier transition location than the previous numerical study by He & Zhong (2023 J. Spacecr. Rockets, vol. 60, no. 6, pp. 1927–1938). Additionally, a discrete entropy-layer mode is recovered, contributing to instability development alongside modes F and S. These findings support the use of bi-orthogonal decomposition as a practical tool for receptivity analysis and enhancement of the amplitude method in transition prediction.

Laminar–turbulent transition in hypersonic boundary layers significantly influences heat transfer, skin friction, and flow separation. To examine the combined effects of shock wave–boundary layer interaction (SBLI) and adverse pressure gradients on boundary layer transition under hypersonic conditions, a flared cone–swept fin configuration was designed. High-fidelity transition data were obtained through wind tunnel experiments utilizing temperature-sensitive paint (TSP), temperature sensors, and high-frequency pressure sensors from PCB Piezotronics (PCB). TSP measurements revealed a triangular high-heat-flux region at the rear of the model, induced by the interaction between SBLI and the adverse pressure gradient. Temperature sensors provided precise wall heat flux, capturing a distinct “growth-decrease-regrowth” heat flux distribution pattern along the flared cone, highlighting the complex interplay between flow structures and transition phenomena. PCB sensors identified dominant instability modes, including low-frequency disturbances and second-mode instabilities. Furthermore, numerical simulations validated the wind tunnel results by reproducing the basic flow structures. This study presents the comprehensive wind tunnel dataset on transition for a complex configuration influenced by both SBLI and adverse pressure gradients, assessing the effects of Reynolds and Mach numbers on the flared cone–swept fin configuration. The findings offer valuable insights for transition prediction and thermal protection design in hypersonic vehicles with complicate fin–body interactions.

Accurate modeling of laminar-to-turbulent transition is crucial for the design of hypersonic flight systems. However, the current transition models used in production CFD codes are insufficient for high-speed flows. Many extensions to low-speed models have been suggested; however, a thorough verification and validation effort is needed before these models can be used in design settings. A meaningful assessment of the generalization capability of these models is also hindered by a lack of information regarding the specific flow configurations and grids employed for model calibration. In this work, we present an independent assessment of two models for high-speed transition, namely, a one-equation model within the shear-stress-transport [Formula: see text] framework and a two-equation model based on the [Formula: see text] equations. These models are implemented in the NASA OVERFLOW 2.3e solver, and a thorough description of the models, numerical parameters, and coupled turbulence modeling parameters is given. Additionally, a formal grid convergence study is conducted for six test cases, covering first-mode and second-mode transitions, such that grid-independent predictions for each model for each case are found. These contributions represent a critical step in the characterization of the models and provide a basis for future users to verify their predictions.