高超声速烧蚀转捩相关的综述文献

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.

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.

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.

No abstract available

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.

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.

No abstract available

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.

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.

No abstract available

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.

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 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.

No abstract available

No abstract available

The crossflow-induced transition for a Mach 6 flow over a swept plate with a 35 mm nose radius and a sweep angle of 45 deg is investigated using direct numerical simulations (DNSs). In order to shed light on the complete transition process, the evolution of a stationary crossflow wave is first simulated. Then, unsteady wall blowing/suction perturbations are introduced to trigger transition to turbulence. The results show that both the type I and III secondary modes are excited, and they subsequently undergo a linear stage of amplification before breakdown. The type II mode is undetected despite its amplification predicted by the two-dimensional eigenvalue stability approach. Overall, the type I mode achieves a dominant amplitude and plays a key role in the transition. Furthermore, a transition correlation method is proposed based on the dominant secondary instability and threshold amplitude concept, and in order to verify and calibrate it, wall perturbations with different amplitudes are introduced to mimic the varying intensities of the background noise. The transition locations predicted by the most amplified secondary mode agree well with those by DNSs, validating the secondary-instability-based criterion for hypersonic crossflow transition.

Air-blowing is one of the techniques for active flow control and thermal protection system of hypersonic vehicles. Introducing air into the hypersonic boundary layer alters the cross-sectional profile of the boundary layer, thereby influencing the boundary-layer transition. This study investigates the active air-blowing control effects on the hypersonic flat-plate boundary layer under various blowing mass flow rates and incoming Mach numbers by solving the Reynolds-averaged Navier–Stokes equations with the Langtry–Menter four-equation transitional shear stress transport model. The study examined alterations in the blowing boundary-layer profiles under two conditions: natural and bypass transition, induced by different blowing flow rates. Blowing significantly alters the sonic line and boundary-layer profile characteristics, triggering blowing oblique shock and causing alterations in the instability mechanisms of the two transition states. A higher Mach number intensifies compressibility effects, stabilizing the boundary layer and leading to an increase in the thickness of the blowing boundary layer and air film.

Görtler vortex-induced hypersonic boundary layer transition controlled by grooves is investigated using direct numerical simulations and spatial bi-global stability analysis. In the simulations, Görtler vortices are excited by wall steady blowing and suction with spanwise wavelengths of 3 mm. It is found that when the wall is covered with grooves, the Görtler streaks keep more regular even at the end of the model. In addition, the skin friction coefficient is reduced efficiently. Furthermore, the wall-normal and spanwise velocity shear are both reduced, suppressing growths of secondary instabilities. In conclusion, grooves can delay Görtler vortex-induced transition by modifying the Görtler streaks structure and instability, which would shed light on hypersonic boundary layer transition control.

Experiments were executed to investigate the effect of ablative outgassing on boundary-layer instabilities on a slender cone in Mach 6 quiet flow. A [Formula: see text] half-angle cone with solid and porous nose tips was tested. The porous nose tips were of varying bluntness: nominally sharp, [Formula: see text], and [Formula: see text]. The permeability of each nose tip was measured using a modified ISO standard, and the outflow was measured using miniature hot-wire anemometry. The model was instrumented with surface pressure transducers and an atomic layer thermopile, and high-speed schlieren was used to visualize the boundary layer. Experiments were done in the Boeing/AFOSR Mach 6 Quiet Tunnel at Purdue University at approximate Reynolds numbers of [Formula: see text] and [Formula: see text] and with varying blowing rates. Without blowing, bluntness suppressed the second-mode instability. In all cases, blowing advanced transition upstream. On the sharp cone and the [Formula: see text] geometry, second-mode instabilities were the dominant transition mechanism. With the [Formula: see text] nose tip installed, second-mode waves were not observed. However, blowing advanced transition, and a 40 kHz instability dominated laminar-to-turbulent transition on the [Formula: see text] geometry.

This study investigates the crossflow breakdown of a Mach 6 flow over a swept flat plate by direct numerical simulation (DNS) considering three cases with different spanwise wavenumbers of stationary vortices. Transition in these cases is initiated by the linear and nonlinear evolution of these vortices, followed by secondary instabilities and breakdown due to type-I, type-II modes, and wall blowing/suction perturbations, respectively. The results showed that amplified secondary instabilities significantly distort the mean flow, causing a steep rise in the wall friction coefficient. Fourier analysis shows that, in this fast-varying flow region, the low-frequency disturbances undergo significantly greater amplifications than high-frequency disturbances. Moreover, the stability characteristics of the time- and spanwise-averaged mean flow were examined to elucidate the breakdown mechanisms. It was found that the unstable region initially contracts to a lower frequency band and then expands significantly in the spanwise wavenumber range at low frequencies. This suggests the significant amplifications of low-frequency disturbances, consistent with the observations from DNS. These amplified low-frequency disturbances, in turn, modify the mean flow, leading to the final breakdown. The presented mechanisms, highlighting the critical role of low-frequency disturbances in the breakdown process, are likely to be universally relevant across various parameter regimes.

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.

No abstract available

No abstract available

No abstract available

Abstract Three-dimensional numerical simulations of hypersonic boundary layer transition delay due to porosity representative of carbon-fibre-reinforced carbon-matrix ceramics (C/C) were carried out on a 7$^{\circ {}}$ half-angle cone for unit Reynolds numbers $Re_m=2.43 \times 10^6$–$6.40\times 10^6\ \text {m}^{-1}$, at the free-stream Mach number $M_\infty =7.4$, for both sharp and 2.5 mm nose tip radii. A broadband time-domain impedance boundary condition was used to model the acoustic effects of the porous surface on the flow field. A quasi-spectral sub-filter-scale dynamic closure was adopted to stabilize the computations upon turbulent breakdown under extreme cooling conditions, with wall-to-adiabatic temperature ratio of $T_{w}/ T_{ad} \simeq 0.08 $, while accurately recovering the growth rates of the unstable modes present in the early transition stages. Good agreement is observed with the reference experimental data, both in terms of the predicted extent of the transition delay and the measured second-mode frequency spectrum. The latter is strongly modulated by the formation of near-wall low-temperature three-dimensional streaks. Pressure disturbances concentrate in corridors of locally thickened boundary layer, with frequencies lower than what predicted by linear theory. Here, trapped wavetrains are formed, which can persist long into the turbulent region. Finally, it is shown that the presence of a porous wall simply shifts the onset of turbulence downstream, without affecting its structure.

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.

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.

Hypersonic boundary-layer transition onset is commonly characterized in wind tunnel experiments by measuring the surface heat transfer rise above the laminar level. Techniques such as infrared thermography and thin film gauges are routinely used in the field. However, when an interfering cooling effect is present due to foreign gas transpiration, these methods are known to be inadequate. This study uses a 7° half-angle cone at Mach 7 with helium or nitrogen injection through a porous segment within the model frustum. The injector spans 60° in azimuth and is located 300 mm from the sharp nose tip, close to the onset of natural boundary-layer transition. Nitrogen and helium injection reduce the surface heat flux below the laminar level for up to 50 mm downstream of the injector. Comparisons to schlieren images and pressure measurements indicate an advance of transition. Optical diagnostics reveal how instabilities are pushed away from the model surface by the injected gas. This is found through spectral analysis of schlieren images and focused laser differential interferometry signals, which revealed further information about how inaccuracies of detecting transition with surface gauges under the influence of transpiration cooling originate.

By using a high-order accurate finite difference scheme, direct numerical simulation of hypersonic flow over an 8° half-wedge-angle blunt wedge under freestream single-frequency entropy disturbance is conducted; the generation and the temporal and spatial nonlinear evolution of boundary layer disturbance waves are investigated. Results show that, under the freestream single-frequency entropy disturbance, the entropy state of boundary layer is changed sharply and the disturbance waves within a certain frequency range are induced in the boundary layer. Furthermore, the amplitudes of disturbance waves in the period phase are larger than that in the response phase and ablation phase and the frequency range in the boundary layer in the period phase is narrower than that in these two phases. In addition, the mode competition, dominant mode transformation, and disturbance energy transfer exist among different modes both in temporal and in spatial evolution. The mode competition changes the characteristics of nonlinear evolution of the unstable waves in the boundary layer. The development of the most unstable mode along streamwise relies more on the motivation of disturbance waves in the upstream than that of other modes on this motivation.

Hypersonic flow conditions pose exceptional challenges for Reynolds-averaged Navier–Stokes (RANS) turbulence modeling. Critical phenomena include compressibility effects, shock/turbulent boundary layer interactions, turbulence–chemistry interaction in thermo-chemical non-equilibrium, and ablation-induced surface roughness and blowing effects. This comprehensive review synthesizes recent developments in adapting turbulence models to hypersonic applications, examining approaches ranging from empirical modifications to physics-based reformulations and novel data-driven methodologies. We provide a systematic evaluation of current RANS-based turbulence modeling capabilities, comparing eddy viscosity and Reynolds stress transport formulations in their ability to predict engineering quantities of interest such as separation characteristics and wall heat transfer. Our analysis encompasses the latest experimental and direct numerical simulation datasets for validation, specifically addressing two- and three-dimensional equilibrium turbulent boundary layers and shock/turbulent boundary layer interactions across both smooth and rough surfaces. Key multi-physics considerations including catalysis and ablation phenomena along with the integration of conjugate heat transfer into a RANS solver for efficient design of a thermal protection system are also discussed. We conclude by identifying the critical gaps in the available validation databases and limitations of the existing turbulence models and suggest potential areas for future research to improve the fidelity of turbulence modeling in the hypersonic regime.

Measurements in high-enthalpy flows are important to understand hypersonic flight and reentry environments. In this study, we use nanosecond coherent anti-Stokes Raman scattering (CARS) to simultaneously probe CO and [Formula: see text] molecules within the reaction layer of a graphite sample exposed to an atmospheric pressure plasma plume. The plasma plume is generated by an inductively coupled plasma torch, with temperatures in the plasma freestream ranging from 5000 to 7000 K. The CARS measurement volume, with a diameter of approximately [Formula: see text], can be positioned as close as 0.2 mm from the graphite surface. The acquired CARS spectra are fitted to theoretical results to determine rotational-vibrational equilibrium temperatures and relative CO to [Formula: see text] mole fractions. We present highly spatially resolved profiles of these quantities in the boundary layer of the graphite ablator, report uncertainties, and discuss the presence of chemical nonequilibrium in the boundary layer.

There are represented research results related to the study of phase transitions in the wall boundary layer occurring during the flow of an ablating surface in a hypersonic stream. The influence of the catalytic wall on the heat flux is considered. The main attention is paid to the analysis of the melting bodies ablation from the surface of high-speed aircraft, based on a detailed recordkeeping of heterogeneous catalytic reactions flow mechanism in surface mass transfer conditions. The distribution of temperature factors over the thickness of the boundary layer at the critical point of the blunted body for a particular flight path segment is given. The ablation from the surface of crystalline bodies is designated. It is shown that the flow of non-Newtonian fluids escapes Newton’s law. Liquids are considered are highly inhomogeneous and consists of large molecules forming complex spatial structures. Moreover, the faster the external impact occurs on the binder macromolecules suspended in a liquid, the higher will be its viscosity. A large temperature gradient is researched within the layer of liquid that forms on the surface of an ablating material. The necessity of simultaneous integration of the entire system of the equation of a liquid layer which consists of equations of continuity, motion and energy together with the dependence of the viscosity coefficient on temperature and tangential stress, is explained.There are represented research results related to the study of phase transitions in the wall boundary layer occurring during the flow of an ablating surface in a hypersonic stream. The influence of the catalytic wall on the heat flux is considered. The main attention is paid to the analysis of the melting bodies ablation from the surface of high-speed aircraft, based on a detailed recordkeeping of heterogeneous catalytic reactions flow mechanism in surface mass transfer conditions. The distribution of temperature factors over the thickness of the boundary layer at the critical point of the blunted body for a particular flight path segment is given. The ablation from the surface of crystalline bodies is designated. It is shown that the flow of non-Newtonian fluids escapes Newton’s law. Liquids are considered are highly inhomogeneous and consists of large molecules forming complex spatial structures. Moreover, the faster the external impact occurs on the binder macromolecules suspended in a liquid, the high...

Measurements of carbon monoxide (CO) during air–carbon ablation of graphite are performed in the Sandia Hypersonic Shock Tunnel. Resistive heating is used to bring graphite samples to wall temperatures ([Formula: see text]) as measured by an imaging pyrometer. The heated models are subjected to a flow condition of [Formula: see text], [Formula: see text], [Formula: see text], and [Formula: see text]. Tunable diode laser absorption spectroscopy measures CO temperature and concentration within the boundary layer at 50 kHz. Ablation product concentrations from experiments are presented as a function of time, including the transient startup of the tunnel. The measured CO concentration increases by approximately a factor of 2 as surface temperatures increase from 1250 to 1630 K. Finally, the CO concentration is observed to be lower than that predicted by several air–carbon ablation models, with the best agreement occurring at the higher surface temperature.

No abstract available

Abstract The stability characteristics of a Mach $5.35$ boundary-layer flow over a flat plate with parametrised two-dimensional sinusoidal surface roughness are investigated. The investigation involves varying the roughness height from $10\,\%$ to $44\,\%$ of the boundary-layer thickness and exploring wavelengths ranging between $0.44$ and $3.56$ times the dominant second-mode wavelength in the region. The introduction of surface roughness leads to notable variations in the mean flow, resulting in separation behind the roughness elements and the propagation of local compression and expansion waves into the free stream. Stability investigations involved the utilisation of wave packet tracking in a linear disturbance simulation (LDS) framework and linear stability theory. The findings revealed significant effects on Mack modes including a reduction in frequency corresponding to maximum amplification with increased roughness height. Proper scaling of the dominant wavelength facilitates a collapse of the growth rate data. In contrast to the commonly reported stabilisation effects for roughness wavelengths significantly larger than the instability mode’s wavelength, the findings primarily revealed destabilisation compared with the smooth-wall case, except for cases with very small roughness wavelengths and large amplitudes approaching the threshold of being classified as porous media. The LDS findings depicted lobed wall pressure amplitude plots, indicating potential undiscovered instability mechanisms or differences compared with the smooth wall. A detailed stability analysis elucidates these LDS findings, establishing a connection between the lobed amplitude structures and substantial changes in local stability characteristics, along with the emergence of Mack’s first, second and third modes.

Abstract Boundary-layer instability and transition control have drawn extensive attention from the hypersonic community. The acoustic metasurface has become a promising passive control method, owing to its straightforward implementation and lack of requirement for external energy input. Currently, the effects of the acoustic metasurface on the early and late transitional stages remain evidently less understood than the linear instability stage. In this study, the transitional stage of a flat-plate boundary layer at Mach 6 is investigated, with a particular emphasis on the nonlinear mode–mode interaction. The acoustic metasurface is modelled by the well-validated time-domain impedance boundary condition. First, the resolvent analysis is performed to obtain the optimal disturbances, which reports two peaks corresponding to the oblique first mode and the planar Mack second mode. These two most amplified responses are regarded as the dominant primary instabilities that trigger the transition. Subsequently, both optimal forcings are introduced upstream in the direct numerical simulation, which leads to pronounced detuned modes before breakdown. The takeaway is that the location of the acoustic metasurface is significant in minimising skin friction and delaying transition onset simultaneously. The bispectral mode decomposition results reveal the dominant energy-transfer routine along the streamwise direction – from primary modes to low-frequency detuned modes. By employing the acoustic metasurface, the nonlinear triadic interaction between high- and low-frequency primary modes is effectively suppressed, ultimately delaying transition onset, whereas the late interaction related to lower-frequency detuned modes is reinforced, promoting the late skin friction. The placement of the metasurface in the linearly unstable region of the second mode delays the transition, which is due to the suppressed streak in the oblique breakdown scenario. However, in the late stage of the transition, the acoustic metasurface induces an undesirable increment of skin friction overshoot due to the augmented shear-induced dissipation work, which mainly arises from reinforced detuned modes related to the combination resonance. Meanwhile, by restricting the location of the metasurface upstream of the overshoot region, this undesirable augmentation of skin friction can be eliminated. As a result, the reasonable placement of the metasurface is crucial to damping the early instability while causing less negative impacts on the late transitional stage.

Abstract Many hypersonic flows of interest feature high free-stream stagnation enthalpies, which lead to high flow-field temperatures and thermochemical non-equilibrium (TCNE) effects, such as finite-rate chemistry and vibrational excitation. However, very few studies have considered receptivity for high-enthalpy flows. In this paper, we investigate the receptivity of a high-enthalpy Mach 5 straight-cone boundary layer to slow and fast acoustic free-stream waves using direct numerical simulation alongside linear stability theory and the linear parabolised stability equations. In addition, we investigate the TCNE effect on receptivity by comparing results between the TCNE gas model and a thermochemically frozen gas model. The dominant instability mechanism for this flow configuration is found to be Mack’s second mode, with the unstable mode being the fast mode. Second-mode receptivity coefficients are obtained for a number of frequencies. For free-stream slow acoustic waves, these receptivity coefficients are found to generally increase with frequency. For a small subset of the considered frequency range, the receptivity coefficients corresponding to free-stream fast acoustic waves are found to be several times larger than for free-stream slow acoustic waves. The TCNE effects are found to lead to higher peak $N$ -factors while also reducing second-mode receptivity coefficients, indicating that TCNE effects have competing impacts on receptivity versus stability for the considered frequencies.

The influences of the forward-facing step (FFS) and backward-facing step (BFS) on the 7° half-angle cone's boundary layer instability at different meridians are investigated at Mach 6, with step heights of 0.6 mm, and a 1° angle of attack (AoA) in a hypersonic quiet wind tunnel using the Nano-tracer-based Planar Laser Scattering techniques, and high-frequency pressure sensors. The results show that at AoA = 1°, the instability that dominates conical boundary layer transition is the second-mode wave. At a given streamwise location, as the position approaches the centerline of the leeward side, the frequency of the second-mode wave decreases while its amplitude increases, resulting in a more unstable boundary layer. Moreover, the development of the second-mode waves in the BFS model progresses faster than that in the FFS model, at different meridians. Notably, on the leeward side, the FFS stabilizes the second-mode waves, whereas the BFS model destabilizes them.

Seepage through microstructure on the surface, as a novel approach for boundary layer flow control, holds considerable promise for drag and thermal reduction in hypersonic vehicles. This paper presents an experimental study conducted in a Mach 6 hypersonic high-temperature wind tunnel, utilizing a cone with seepage through microstructure on the surface. The study employs schlieren, high-frequency fluctuating pressure measurement techniques, and infrared thermometry. It focuses on the influence of the amplitude-frequency characteristics of the second-mode waves within the hypersonic boundary layer on aerodynamic heating on the wall. According to the research, the following findings were observed: For smooth surfaces, the characteristic frequency of the second-mode wave decreases from 109 to 87 kHz along the flow direction. However, microstructures cause an average increase in the characteristic frequency by approximately 20 kHz, while seepage causes the characteristic frequency along the flow direction to change little. Under the same seepage flow rate, the amplitude of the second-mode waves within the boundary layer exhibits two peaks along the flow direction. When the second-mode wave reaches the first peak, the wall temperature rise also peaks at the same location, indicating a synchronous increase and decrease between the second-mode waves and wall temperature rise during this stage. Under different seepage flow rates, the downstream wall with microstructure-seepage experiences the lowest temperature rise when F = 0.019 19% which may be attributed to the formation of small vortices within the microstructures by the seepage.

Nonmodal linear stability analysis results are presented for hypersonic flow over a cone at 6° angle of attack complementing earlier modal stability analysis. Based on the parallel flow assumption, singular value decomposition is applied to obtain the optimal linear combination of global crossflow modes. The optimal disturbance exhibits significant transient growth in the initial short distance and progressively follows the path of the most unstable mode downstream. The largest transient energy gain is observed for disturbances at around 40 kHz close to the most amplified modal frequency and tends to increase with the Reynolds number. The optimal disturbance initially exhibits two amplitude peaks in the azimuthal direction, one lying in the leeward region where the unstable crossflow modes reside and the other in the windward region where the adjoint modes exist. As the optimal disturbance travels downstream, the second amplitude peak rapidly shifts toward the leeward side and reaches the optimal energy gain when it eventually merges with the first amplitude peak. The evolution process of the optimal disturbance indicates that the optimal disturbance might have exploited the locally crossflow instability through traveling from the windward side to the leeward side.

No abstract available

No abstract available

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.

No abstract available

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.

Numerical results are presented for the stability analysis of the wake induced by a cuboidal roughness element mounted on a flat plate inside a Mach 6 freestream. Linear BiGlobal stability calculations are carried out for a single frequency on a spanwise plane located behind the roughness, using base flows obtained from laminar Navier–Stokes simulations. The results show that the Mack mode is the most unstable perturbation growing in the boundary layer, followed by varicose and sinuous deformations of the low-velocity streak that characterizes the wake flow structure. The shock wave induced at the leading edge of the flat plate is found to have a significant stabilizing effect on the flow field. The use of a higher wall temperature stabilizes the Mack mode but increases the growth rate of the varicose perturbation.

No abstract available

No abstract available

The thermal protection materials enveloping hypersonic vehicles ablate. This paper develops a numerical methodology of thermochemical nonequilibrium flow and structure temperature field to study the key flow and structure parameters of a hypersonic hemispherical nose under the condition of carbon-based thermal protection material ablation. The multi-species chemical reaction model of flow field and the corresponding numerical algorithm are constructed. Oxidation and sublimation ablation process are considered in ablation model coupled with flow field solver by gas-solid interaction method. The computational results of a hemispherical nose indicate that carbon-based thermal protection material ablation has significant effects on thermochemical nonequilibrium flow. Details of the ablation flow characteristics are reported to highlight the influences of ablation on microcosmic reaction species and macroscopic flow parameters including shock wave position, pressure and temperature.

No abstract available

Research on hypersonic crossflow transition holds significant engineering and scientific importance. This paper investigates the impact of distributed roughness elements (DREs) on crossflow transition for a cone set at a 6° angle of attack, using experimental methods. The research was conducted in the Mach 6 wind tunnel, employing temperature-sensitive paint (TSP) as the measurement technology. Two types of nosetip were examined: a sharp nosetip with a radius of 0.1 mm and a blunt nosetip with a radius of 2.5 mm. The circumferential wavenumbers of the DREs on the nosetip included k = 35, k = 50, and k = 70. The results indicate that the nosetip with DREs at k = 50 has a more pronounced effect in promoting boundary layer transition to turbulence on the leeward side of the cone compared to the nosetips with DREs at k = 35 and k = 70. However, all three types of DREs exhibit similar effects on transition on the windward side. Additionally, the bluntness of the nosetip, at R = 2.5 mm, diminishes the effectiveness of DREs in promoting transition; however, the degree of diminished effectiveness varies with the circumferential azimuth.

It is difficult to diagnose the electron density of a high-temperature ablation plasma flow field because a traditional cylindrical Langmuir probe (CP) is easily damaged under these conditions. In this work, a new type of embedded Langmuir probe, referred to as a double flush-mounted probe (DFP), was developed to measure the electron density of a high-temperature ablation plasma flow field. It was verified that the DFP can work stably in different types of wind tunnels. In addition, the results from the new probe were compared with those from a CP. The results suggest that the DFP can be used to accurately determine the plasma density over long time periods. Therefore, this work provides a feasible method for solving the problem of online diagnostics in a high-temperature ablation plasma flow field.

No abstract available

No abstract available

No abstract available

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.

The impact of Görtler vortices on roughness-induced transition in a Mach 6 concave hypersonic boundary layer is examined through implicit large eddy simulation. Our prior numerical investigations indicated that the evolution of the most amplified Görtler vortices appeared to have minimal influence on the process of roughness-induced transition within a hypersonic boundary layer. In this study, we conduct a comprehensive analysis of the effects of Görtler vortices on roughness-induced transition by simulating various roughness locations, dimensions, and multiple Görtler vortices. In hypersonic flows, the trapped layer Görtler mode is predominant; however, the most amplified Görtler vortices exhibit negligible influence on roughness-induced vortices, regardless of whether the roughness location or dimension is altered. This behavior contrasts sharply with that observed in other Görtler vortices. Amplified Görtler vortices with a wavelength twice that of the most amplified Görtler vortices significantly affect roughness-induced vortices and delay the transition process. It appears that distinct Görtler vortices exert varying degrees of influence on roughness-induced transition.

Temperatures within the boundary layers of high-enthalpy hypersonic flows can soar to thousands or even tens of thousands of degrees, leading to significant real gas phenomena. Although there has been significant research on real gas effects on hypersonic boundary layer stability, their impact on the boundary layer’s receptive stage is still poorly understood. Most aerodynamic boundary layers in flight vehicles are three-dimensional. Because of complex geometry and significant crossflow effects, the crossflow mode in three-dimensional boundary layers is crucial in hypersonic vehicle design. In this study, a linear stability analysis (LST) accounting for chemical nonequilibrium effects (CNE) and its adjoint form (ALST) is developed to investigate the real gas effects on the stability and receptivity of stationary crossflow modes. The results indicate that real gas effects significantly influence the receptivity of stationary crossflow modes. Specifically, chemical nonequilibrium effects destabilize the crossflow modes but reduce the receptivity coefficients of the stationary crossflow modes. The Mach number effect was also investigated. It was found that increasing the Mach number stabilizes the stationary crossflow modes, but the receptivity coefficients increase. As the Mach number progressively rises, these effects alternately dominate, leading to a non-monotonic shift in the transition position.

Wall roughness has significant effects on hypersonic boundary layer transition. In this paper, the propagation process of linear perturbations passing two-dimensional single or double hump-type roughness elements is simulated under Mach number 5.92 using the harmonic linearized Navier–Stokes equation. For the first time, the kinetic energy budget analysis is adopted to study the influencing patterns and mechanisms of the humps on the perturbations. For cases with a single hump, the hump has little effect on low-frequency perturbations, since perturbations are first promoted by the non-parallelism mechanism in the region upstream of the hump and then suppressed above the hump due to the negative production term. At higher frequencies, the hump can significantly promote/suppress the perturbation below/above the synchronization frequency, which is mainly caused by the production and non-parallelism terms in the upstream region of the hump. For cases with double humps, when they are close, it is found that their effects on perturbations cannot be obtained by the linear superposition of their individual effect. The interaction between them can make the promotion or suppression effects on perturbations significantly weaker than their linear superposition. This is due to the fact that the first hump occupies the upstream region of the second hump and that the contribution of the second hump's upstream region is limited. Under the condition in this paper, until the distance between the two humps is greater than 14δ*, where δ* is the local boundary layer thickness, the magnitude of their interaction is less than 5%.

A supersonic (M∞=5.95) flow around a blunted cone with a cylindrically shaped single roughness element located in the bluntness region is considered. Numerical simulations are performed by means of solving the unsteady three-dimensional Navier-Stokes equations with a computational module that takes into account the action of external acoustic waves on the flow. Numerical data are obtained for the mean flow characteristics and evolution of disturbances formed in the wake behind the roughness element. It is shown that intense interaction of steady streamwise vortices occurs behind the single roughness element (Rekk=4880, k/δ=3.95) on the cone with a bluntness radius of 5mm. This interaction is accompanied by formation of three-dimensional finger-shaped structures in the azimuthal direction, which testifies to the initial stage of the laminar-turbulent transition.

Surface roughness is known to have a substantial impact on the aerothermodynamic loading of hypersonic vehicles, particularly via its influence on the laminar-turbulent transition process within the boundary layer. Numerical simulations are performed to investigate the effects of a distributed region of densely packed, smooth-shaped roughness elements on the laminar boundary layer over a 7-degree half-angle, circular cone for flow conditions corresponding to a selected trajectory point from the ascent phase of the HIFiRE-1 flight experiment. For peak-to-valley roughness heights of 50 percent or less in comparison with the thickness of the unperturbed boundary layer, the computations converge to a stationary flow, suggesting that the flow is globally stable. Analysis of convective instabilities in the wake of the roughness patch indicates two dominant families of unstable disturbances, namely, a high frequency mode that corresponds to Mack mode waves modified by the wake and a lower frequency mode that corresponds to shear layer instabilities associated with the streaks in the roughness wake. Even though the peak growth rate of the later mode is more than 35 percent greater than the peak growth rates of the Mack modes, the latter modes achieve higher amplification ratios, and hence, are likely to dominate the onset of transition, which is estimated to occur slightly later than that in the unperturbed, i.e., smooth surface boundary layer. Additional computations are performed to investigate the effects of various roughness patch configurations on a Mach 3.5 flat plate boundary layer, to help guide an upcoming experiment in the Mach 3.5 Supersonic Low Disturbance Tunnel at NASA Langley Research Center. In this case, the cumulative reinforcement of basic state distortion over the length of the roughness patch is predicted to yield a significantly earlier transition than that over a smooth plate or a plate with a shorter length roughness patch.

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.

No abstract available

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.

No abstract available

Laminar-turbulent transition caused by modal disturbance growth in the wake flow of isolated roughness elements on blunt re-entry capsules is studied numerically at typical cold hypersonic wind-tunnel conditions. Two fundamentally different heat shield shapes are considered. On the sphere-cone forebody the wake flow of the roughness is exposed to an adverse pressure gradient, whereas the spherical heat shield exhibits a strongly favorable pressure gradient. The pressure gradient effects on the development of the stationary wake flow and its modal instability characteristics are discussed for various heights and diameters of the cylindrical roughness element. Regions of increased shear develop in its wake, which persist longer in the adverse pressure gradient case. Consequently, the results of spatial two-dimensional eigenvalue analyses reveal that the unstable wake-flow region extends much further downstream and the wake-mode instabilities are considerably more amplified. The disturbance kinetic energy production terms are used to assess the contributions of the different shear-layer regions to the mode growth and its dependence on the pressure gradient.

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.

Experiments and computational analysis have been conducted to study instabilities in the boundary layer over a swept fin–cone geometry in hypersonic flow. Experiments were carried out at the United States Air Force Academy’s Mach 6 Ludwieg Tube. Infrared thermography was employed to assess the surface temperature distribution, particularly the thermal striations on the fin that indicated a crossflow-dominated transition. Increasing the unit Reynolds number increased heating and moved the crossflow-transition front closer to the fin leading edge. Stability analysis based on the linear parabolized stability equations identified unstable stationary crossflow modes across the range of tested unit Reynolds numbers, and the most amplified disturbances were found to have wavelengths of 2–4 mm. The experimentally observed thermal striations were analyzed and found to have initial wavelengths of 2–5 mm but eventually formed higher-amplitude structures with a wavelength of approximately 10 mm. Discrete roughness elements of different diameters and spacings were implemented near the fin leading edge. Boundary-layer transition onset was found to be delayed with a subsequent lowering of the fin surface heating, with discrete roughness elements having a wavelength of 13.33 mm.

Direct numerical simulation with up to 10×109 scale grid points based on graphics processing unit computation is carried out to investigate the bluntness effect on the hypersonic boundary-layer transition over a slender cone with zero angle of attack at Mach 6. Four cases with the nose radii of 1, 10, 20, and 40 mm are conducted, and the corresponding Reynolds number based on the nose radius varies from 1.0×104 to 4.0×105. Random disturbances with a broad spectrum of frequencies and a wide range of azimuthal wavenumbers were applied to the wall to simulate disturbances caused by wall roughness. The numerical results show that as the nose tip radius increases, the transition position gradually moves downstream with increased transition region. For the case with a nose radius of 1 mm, the flow transition and entropy swallowing occur almost simultaneously, while for other cases, the transition takes place earlier than the entropy swallowing. In consequence, the disturbance amplitude upstream of the transition in the 1 mm case is much larger than that of other cases. To further study the mechanism of the transition, the frequency spectrum analysis is carried out. It is found that all cases exhibit two characteristic frequencies within the transition region, i.e., the high frequency and extremely low frequency. Owing to the influence of the entropy layer, the characteristic high frequency of the 1 mm case is significantly higher than that of other cases. With the increase in the nose radius, the characteristic frequency of the high frequency decreases gradually.

This study investigates a novel method to control hypersonic boundary layer transition using a combined local cooling and local metasurface treatment. The method’s effectiveness was investigated on a 5-degree half-angle blunt wedge with a nose radius of 0.0254 mm at a freestream Mach number of 6.0 using direct numerical simulations and linear stability theory. We explored four cases: (i) adiabatic baseline case, (ii) locally cooled case, (iii) local metasurface case, and (iv) combined local cooling-local metasurface case. Results showed that the combined local cooling-local metasurface treatment significantly reduced both wall pressure disturbance amplitude and the density perturbation amplitude around the sonic line, indicating a potential for controlling hypersonic boundary layer transition. In the local cooling-local metasurface case, the disturbance amplitude at the end of the computational domain was 270 times lower than the baseline case. The study also examined the impact of Reynolds numbers, ranging from 25.59 million per meter to 32.80 million per meter. Unsteady simulations revealed that the Reynolds number had a negligible effect on the local cooling-local metasurface performance, indicating that the proposed method applies to a wide range of flight conditions.

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.

No abstract available

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.

Accurate prediction of hypersonic transitional boundary layer (HTrBL) transition is critical for high-speed vehicle design but remains challenging due to complex, multi-factor characteristics of the flow. This study develops a data-driven framework integrating machine learning (ML) with Structural Ensemble Dynamics (SED) theory to enhance the SED-based stress-length transition model for sharp cones, addressing the key challenge of accurately capturing transition dynamics across diverse flow conditions. The framework uses an Ensemble Kalman Filter to infer key model parameters, such as near-wall vortex scale and transition center, from limited experimental heat flux data, achieving prediction errors less than 6% for transition onset and peak heat flux. A hybrid feature selection strategy combines global flow parameters and local flow features to train a Hyper-Deep Neural Network, enabling accurate generalization across various scenarios. Furthermore, scaling laws for those model parameters are derived, converting the machine-learning model from the “black-box” to a “white-box” system with clear physical interpretability. The model provides unified predictions of the entire transition process, including transition onset location, peak heat flux, and fully developed turbulent state, outperforming traditional approaches like the C–γ–Reθ model. This work establishes a data-driven turbulence modeling paradigm—physics-guided ML—to reveal universal laws and facilitate the advancement of reliable HTrBL applications.

No abstract available

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.

Previous investigations in the High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) show that carbon fiber reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities resulting in a delay of boundary layer transition onset. Numerical stability analyses confirmed these experimental results. However, C/C has some disadvantages, especially the limited oxidation resistance and its low mechanical strength, which could be critical during hypersonic flights. Thus, an ultrasonically absorptive fiber reinforced ceramic material based on a silicon carbide (C/C-SiC) was developed in the past years to fulfill this need. The present paper addresses the numerical rebuilding of the C/C-SiC absorber properties using impedance boundary conditions together with linear stability analysis. The focus of this paper is on the numerical comparison of the original C/C material and the improved C/C-SiC material, referred to as OCTRA in the literature. The influence on the second modes and the transition itself is investigated. The numerical results are compared with HEG wind tunnel tests. The wind tunnel model tested in HEG is a 7∘\documentclass[12pt]{minimal} \usepackage{amsmath} \usepackage{wasysym} \usepackage{amsfonts} \usepackage{amssymb} \usepackage{amsbsy} \usepackage{mathrsfs} \usepackage{upgreek} \setlength{\oddsidemargin}{-69pt} \begin{document}$$7^\circ$$\end{document} half-angle blunted cone with an overall model length of about 1.1m\documentclass[12pt]{minimal} \usepackage{amsmath} \usepackage{wasysym} \usepackage{amsfonts} \usepackage{amssymb} \usepackage{amsbsy} \usepackage{mathrsfs} \usepackage{upgreek} \setlength{\oddsidemargin}{-69pt} \begin{document}$$1.1 \,\textrm{m}$$\end{document} and a nose tip radius of 2.5 mm. These experiments were performed at Mach 7.5 and at different freestream unit Reynolds numbers.

Wall temperature significantly affects stability and receptivity of the boundary layer. Changing the wall temperature locally may therefore be an effective laminar flow control technique. However, the situation is complicated when the wall temperature distribution is nonuniform, and researchers have experimentally found that local wall cooling may delay the onset of transition. We attempt to clarify the physical mechanisms whereby the local wall temperature affects the transition and the stability of a hypersonic boundary layer. A numerical investigation of the disturbance evolution in a Mach-6 sharp cone boundary layer with local wall heating or cooling is conducted. Direct numerical simulation (DNS) is performed for the single-frequency and broadband disturbance evolution caused by random forcing. We vary the local wall temperature and the location of heating/cooling, and then use the e N method to estimate the transition onset. Our results show that local wall cooling amplifies high-frequency unstable waves while stabilizing low-frequency unstable waves, with local heating amplifying all unstable waves locally. The disturbance amplitude and second-mode peak frequency obtained by DNS agree well with the previous experimental results. Local cooling/heating has a dual effect on the stability of the hypersonic boundary layer. For local cooling, while it effectively inhibits the growth of the low-frequency unstable waves that dominate the transition downstream, it also further destabilizes the downstream flow. In addition, while upstream cooling can delay the transition, excessive cooling may promote it; local heating always slightly promotes the transition. Finally, recommendations are given for practical engineering applications based on the present results.

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.