Hypersonic Boundary Layer Research Articles

Measurement of hypersonic turbulent boundary layer on a flat plate using cylindrical focused laser differential interferometer

A modified Cylindrical-Focused Laser Differential Interferometer (CFLDI) system was used to measure the density fluctuations generated by a Mach 6 turbulent boundary layer on a flat plate. The amplitude, spectral statistics, and correlation scale of the density fluctuations were analyzed at different wall-normal heights throughout the boundary layer. Direct numerical simulation was performed under the similar condition to verify the experimental results. The results show that the CFLDI system herewith can accurately represent the statistical characteristics of density fluctuations with a bandwidth of 3–200 kHz in the hypersonic turbulent boundary layer. As the distance from the wall increases, the amplitude of density fluctuations keeps increasing until y/δ ≈ 0.8 and levels off after y/δ ≈ 1.5. As the measurement position moves into the freestream, the characteristic frequency of the density fluctuation decreases rapidly, while its integral timescale gradually increases. Similar phenomena were also observed for pressure fluctuations.

Roughness effect on hypersonic second mode instability and transition on a cone

Recent studies by Fong et al. [“Finite roughness effect on modal growth of a hypersonic boundary layer,” AIAA Paper No. 2012-1086, 2012; “Stabilization of hypersonic boundary layer waves using 2-D surface roughness,” AIAA Paper No. 2013-2985, 2013; “Numerical simulation of roughness effect on the stability of a hypersonic boundary layer,” Comput. Fluids 96, 350–367 (2014); and “Second mode suppression in hypersonic boundary layer by roughness: Design and experiments,” AIAA J. 53, 1–6 (2015)] have shown that finite roughness can attenuate Mack's second mode instability when placed at the discrete mode synchronization location for two-dimensional planar flow over a flat plate. However, more practical hypersonic flows are non-planer conical flows, and the roughness effect phenomenon in conical flows has not been extensively investigated. For that reason, this investigation research the ability of finite roughness strips to attenuate the second mode instability on a Mach 8 straight blunt cone with a freestream unit Reynolds number of 9 585 000. Two roughness configurations are studied: a single roughness strip and an array of six sequential strips. N-factor calculations determine the second mode frequency most likely to lead to turbulent transition, and the linear stability theory is used to determine the mode's synchronization location. In the unsteady simulations of the roughness configurations, a blowing-suction actuator introduces an upstream broadband Gaussian pulse. Fourier decomposition of the pulse's history shows that the single roughness strip attenuates frequencies higher than 208 kHz while lower frequencies are amplified. Likewise, the roughness array exhibits similar results, attenuating frequencies higher than 164 kHz and amplifying lower frequencies downstream. The results show that both configurations can delay second mode instability growth on a hypersonic blunt cone and possibly delay turbulent transition. However, investigations of the roughness effect's behavior downstream of the roughness configurations show that disturbance growth resumes and becomes more destabilizing to the boundary layer.

Direct numerical simulation of hypersonic wall-bounded turbulent flows: An improved inflow boundary condition and applications

In this paper, the method of generating inflow turbulence based on turbulence fluctuation library (TFL) in direct numerical simulation (DNS) of the hypersonic turbulent boundary layer (TBL) is investigated. The application of the TFL method to the DNS of a supersonic TBL shows that, although there are significant differences in freestream between the TFL and the target TBL, the flow could successfully develop to the target TBL downstream as the fluctuations of TFL are suitably scaled and added to the DNS inflow. However, there is a “transition”-like recovery process from the inflow to the target turbulence. To deal with the defects of the thermodynamic fluctuations scaling laws in the current TFL method under the hypersonic TBL, new thermodynamic fluctuations scaling laws are theoretically derived by introducing the generalized Reynolds analogy. The application in the DNS of Mach 7.25 TBL shows that the new scaling laws for thermodynamic fluctuations are more rational and accurate than the previous ones. Furthermore, the study on the recovery process shows that the matching degree between the TFL and the target TBL on the friction Reynolds number (Reτ) is the dominant factor in determining the length of recovery distance. Guaranteeing the similar Reτ of the TFL and the target TBL can make the two possess similar coherence structures, which can avoid the distortion of the coherence structures at the inflow after spanwise and normal interpolation, prevent the process of Reynolds stress decay and readjustment downstream the inflow, and finally effectively shorten the recovery distance.

Receptivity and its influence on transition prediction of a hypersonic boundary layer over a small bluntness cone

Hypersonic boundary-layer receptivity is investigated using direct numerical simulation for Mach 6 flow over a 0.79 mm nose radius, 5° half-angle cone to slow acoustic waves. Two ways of exciting the second mode are discovered: For a low-frequency forcing (region I where f < 370 kHz), the first mode is initially excited and subsequently evolves into the second mode, whereas for a high-frequency forcing (region II where f≥ 370 kHz), the second mode is excited through the disturbances in the entropy layer. The receptivity mechanisms corresponding to the first and second modes are consistent with those of a previously investigated large bluntness cone. The receptivity coefficient shows a dramatical decrease with the frequency for disturbances in region I while a slight increase for disturbances in region II. The results show that the correlation between the receptivity coefficient of the first mode and the frequency proposed for large bluntness cones also applies to the current case. Furthermore, with receptivity and the measured freestream noise spectra taken into account, it is found that at the measured transition location, the disturbance predicted by the traditional eN method is no longer the most amplified disturbance, since the disturbances with lower frequencies can acquire larger amplitudes due to their large receptivity coefficients despite their smaller growth rates.

Effect of two-dimensional micro-cavity surface on hypersonic boundary layer

For hypersonic boundary layer, the porous coating comprising small blind holes is an effective passive method that can delay the laminar-turbulent transition by damping the second mode wave. In this paper, surface with two-dimensional rectangular cavities with larger size comparable to the thickness of boundary layer is numerically investigated. The micro-cavity surface affects not only the second mode wave, but also the base flow due to the roughness effects. The numerical results showed that the total drag of base flow can be reduced by the micro-cavity surface compared to the smooth surface, more significant with smaller cavity width and higher porosity, while the cavity depth has little influence on the drag performance. The total drag can be further reduced when the micro-cavity surface suffers moderate static deformation under the pressure difference in the base flow. The micro-cavity surface is able to damp the second mode wave like the porous coating, and thus also has the potential to delay the transition. The effect on the second mode wave is stronger with higher porosity and smaller cavity width, while not monotonic with the cavity depth. The static deformation of structure causes local change of cavity characteristics, resulting in the second mode more damped over the deformed micro-cavity surface. When the structure dynamically responses to the fluctuating pressure of the second mode wave, the intensity of wave is slightly further decreased.

On the drag reduction mechanism of hypersonic turbulent boundary layers subject to heated wall blowing

Turbulence drag reduction is of great significance for the range increase of hypersonic flight vehicles. The proposed velocity-temperature coupling control method (Liu et al, Phys Rev Fluids 6:044603, 2021) is further extended to the hypersonic turbulent boundary layer. Direct numerical simulation results of four comparative cases show that the heated wall blowing achieves a drag reduction rate of 10.58%, which is about the sum of wall blowing (5.27%) and wall heating (6.35%). By evaluating the control efficiency, however, it is found that heated wall blowing is not as good as wall blowing and cannot obtain net energy saving rate. The modified FIK decompositions of skin friction coefficient indicate that the cliffy decrease of the mean convection term is the primary contribution for the drag reduction. Effects of the proposed control measure on turbulence statistics and coherent structures are also analyzed. Streamwise vortex is found to be away from the wall, thus leading to a lower friction drag.

Data-driven framework for input/output lookup tables reduction: Application to hypersonic flows in chemical nonequilibrium

In this paper, we present a novel model-agnostic machine learning technique to extract a reduced thermochemical model for reacting hypersonic flows simulation. A first simulation gathers all relevant thermodynamic states and the corresponding gas properties via a given model. The states are embedded in a low-dimensional space and clustered to identify regions with different levels of thermochemical (non)-equilibrium. Then, a surrogate surface from the reduced cluster-space to the output space is generated using radial-basis-function networks. The method is validated and benchmarked on a simulation of a hypersonic flat-plate boundary layer with finite-rate chemistry. The gas properties of the reactive air mixture are initially modeled using the open-source Mutation++ library. Substituting Mutation++ with the light-weight, machine-learned alternative improves the performance of the solver by 50% while maintaining overall accuracy.

Investigation of hypersonic cone boundary layer stability regulation with plasma actuation

Boundary layer transition has always been a frontier in the field of fluid mechanics, and hypersonic boundary layer stability experiments can help reveal the physical mechanisms behind such transitions. In particular, the regulation of unstable waves in boundary layers is critical for transition prediction and control. Plasma actuation is a popular flow control method that has made progress in moderating the stability of supersonic boundary layer. However, there have been few studies on regulating the stability of hypersonic boundary layers with plasma actuation. In this paper, wind tunnel experiments are carried out under Mach 6 flow to study the stability regulation of a hypersonic sharp cone boundary layer with nanosecond pulsed plasma actuation. First, the typical characteristic structure of the rope-like structure is captured by the high-speed schlieren method. Then, combining the sensor results and theoretical analysis, the rope-like structure and the dominant instability wave resolved by the schlieren power spectrum density method are determined to correspond to the second-mode wave. The characteristic unstable quasi-ordered structure of the boundary layer under actuation is then extracted, and the impact effect and modulation effect of this structure on the second-mode wave are analyzed. Finally, the mechanism by which actuation influences the boundary layer instability is studied using proper orthogonal decomposition. The results show that actuation can enhance boundary layer pulsation, and the coupling effect between the actuation and boundary layer can produce regular unstable quasi-ordered structures. The intrinsic mechanism works by impinging on and modulating the second-mode waves, and there are characteristic modes of the rope-like structure and the unstable structures distributed over the whole flow-direction range. This verifies the ability of plasma actuation to stimulate the instability of hypersonic cone boundary layers and provides technical support for the further development of transition control methods.

An experimental study on the stability of hypersonic plate boundary layer regulated by a plasma actuation array

Laminar–turbulent transition control is one of the key problems in the design of hypersonic vehicles. In particular, for air-breathing hypersonic vehicles, early onset turbulence in the boundary layer of a scramjet inlet precursor is desirable. Plasma actuation has become a popular flow control method in recent years and has made progress in regulating the stability of supersonic boundary layers. However, there have been few studies on the regulation of the stability of hypersonic boundary layers. In this paper, an experimental study on the stability of the hypersonic plate boundary layer is carried out using a spanwise plasma actuation array. The characteristics and evolution of different kinds of unstable waves in the hypersonic plate boundary layer (especially the dominant first mode wave) are analyzed based on the results of linear stability theory and high-frequency pulsation sensors. The typical morphological characteristics of the boundary layer and the macro-control effect of the plasma actuation array are explored through a high-speed schlieren method. Finally, based on grayscale mode extraction and proper orthogonal decomposition, the influences of three different actuation frequencies on the instability waves and characteristic structure of the boundary layer are studied, including the dominant frequency of the first mode wave, the half-frequency of the dominant frequency of the first mode wave, and the dominant frequency of the second mode wave, the change of characteristic structures under the regulation of plasma actuation is further discussed. The corresponding regulation rules and mechanisms are summarized. The results show that the plasma actuation array can advance the starting position of laminar discontinuities and that the induced coherent structure can excite instabilities at an earlier flow-direction position. The actuation can be used to control the stability of the boundary layer by acting on the first mode wave to break the original unstable wave spectrum characteristics. This verifies the ability of extensional array plasma actuation to regulate the stability of the hypersonic plate boundary layer and suggests it has great potential in the promotion of hypersonic boundary layer transition.

Improvement of transition prediction model in hypersonic boundary layer based on field inversion and machine learning framework

The classical four-equation γ−Reθ transition model has presented excellent accuracy in low-speed boundary layer transition prediction. However, once the incoming flow reaches hypersonic speed, the original model is no longer applicable due to the compressibility problem and the appearance of multiple instability modes. Recently, there has been widespread interest in data-driven modeling for quantifying uncertainty or improving model prediction accuracy. In this paper, a data-driven framework based on field inversion and machine learning is performed to extend the prediction capability of the original γ−Reθ transition model for the hypersonic boundary layer transition. First, the iterative regularized ensemble Kalman filter method is applied to obtain the spatial distribution of the perturbation correction term β for the switching function Fonset1, and the effectiveness of this method is initially verified in the hypersonic flat plate case. Then, the random forest algorithm is adopted to construct a mapping from the average flow features to β. The generalizability of the well-trained learning model is fully validated in the blunt cone cases with different unit Reynolds numbers, free-stream flow temperature, and bluntness. The simulation results indicate that the performance of the original γ−Reθ transition model in the hypersonic boundary layer transition prediction is significantly improved, and the boundary layer transition onset location and the length of transition zone can be correctly obtained. In addition, the machine learning model investigates the importance of the input features and confirms that the effective length scale plays a significant role in the numerical simulation of the hypersonic boundary layer transition.

Large-eddy simulation of a hypersonic turbulent boundary layer over a compression corner

In this paper, large-eddy simulation of the interaction between a shock wave and the hypersonic turbulent boundary layer in a compression corner with a fixed 34° deflection angle at Ma = 6 for different Reynolds number cases is conducted. For investigating the effects of the Reynolds number for hypersonic cases, three cases where the free-stream Reynolds numbers are 14000, 20000, and 30000/mm are selected. The averaged statistics, such as the mean velocity, the skin friction, the heat flux, and the wall pressure, are used in this paper. The flow structures in the compression ramp including the shock wave and interaction region are discussed. The decomposition of the mean skin-friction drag for the flat flow is extended to be used in the compression corner. In addition, the turbulent kinetic energy is studied through the decomposition of the mean skin-friction drag for the flat-plate region and the corner region. It is found that higher Reynolds numbers would increase the turbulent kinetic energy by turbulent dissipation at the interaction region, while higher Reynolds numbers would decrease the turbulent kinetic energy by turbulent dissipation after reattachment. In addition, it is also found that the turbulent kinetic energy is larger with a higher Reynolds number and higher turbulent kinetic energy inhibits the movement from the separation point to the inflection point (x = 0 mm), which deduces larger separation bubbles.

Wall temperature effects on the hypersonic boundary-layer transition over an inclined, blunt cone

Wall temperature is a crucial factor for the transition of hypersonic boundary layers. This paper investigates the wall temperature effects on the hypersonic boundary-layer transition over an inclined blunt cone with two wall temperatures by direct numerical simulations and instability analysis. The breakdown processes of leeward streamwise vortices and crossflow vortices are both studied. Inner modes with lower phase velocities and outer modes with higher phase velocities are responsible for the breakdown of leeward streamwise vortices, respectively, in cooled- and heated-wall cases. A higher wall temperature will promote the boundary layer transition in the vicinity of the leeward centerline due to the collision of nearby crossflow vortices with the leeward mushroom structure, instead of by destabilizing inner and outer modes. In the crossflow-dominated region, the transition is triggered by the interaction between second modes and crossflow vortices for the cooled-wall case. Increasing the wall temperature will promote the transition in this region even though second modes are suppressed.

Numerical Investigation of Wall-Cooling Effect on Aero-Optical Distortions for Hypersonic Boundary Layer

Compressible wall-modeled large-eddy simulations of Mach 8 turbulent boundary-layer flows over a flat plate were carried out for the conditions of the hypersonic wind tunnel at Sandia National Laboratories. The simulations provide new insight into the effect of wall cooling on the aero-optical path distortions for hypersonic turbulent boundary-layer flows. Four different wall-to-recovery temperature ratios, 0.3, 0.48, 0.71, and 0.89, are considered. Despite the much lower grid resolution, the mean velocity, temperature, and resolved Reynolds stress profiles from the simulation for a temperature ratio of 0.48 are in good agreement with those from a reference direct numerical simulation. The normalized root-mean-square optical path difference obtained from the present simulations is compared with that from reference direct numerical simulations, Sandia experiments, as well as predictions obtained with a semi-analytical model by Notre Dame University. The present analysis focuses on the effect of wall cooling on the wall-normal density correlations, on key underlying assumptions of the aforementioned model such as the strong Reynolds analogy, and on the elevation angle effect on the optical path difference. Wall cooling is found to increase the velocity fluctuations and decrease the density fluctuations, resulting in an overall reduction of the normalized optical path distortion. Compared to the simulations, the basic strong Reynolds analogy overpredicts the temperature fluctuations for cooled walls. Also different from the strong Reynolds analogy, the velocity and temperature fluctuations are not perfectly anticorrelated. Finally, as the wall temperature is raised, the density correlation length, away from the wall but inside the boundary layer, increases significantly for beam paths tilted in the downstream direction.

Hypersonic Boundary-Layer Instability Suppression by Transverse Microgrooves with Machining Flaw

Transverse rectangular microgrooves with local machining flaw are investigated using direct numerical simulations from the viewpoint of suppressing the second-mode instability in a Mach 6 hypersonic boundary layer. Two types of machining flaw are modeled for the local deformation of microgrooves, i.e., chamfered and protruded ones. The effect of locally deformed microgrooves on the boundary-layer stability is studied by superposing the second-mode disturbance to the base flow of the boundary layer. The investigated unstable second mode is 400 kHz, and the microgrooves are installed in its synchronous region. The results show that compression and expansion waves are generated on the surface of the deformed microgrooves, and microscale separation bubbles are formed at the rear edge of the deformed microgrooves. With the increase of local deformation height of microgrooves, the suppression effect of the second-mode disturbance gradually increases. As the deformation height reaches 0.09 mm, the amplitude inhibition rate of locally deformed microgrooves on the second-mode disturbance reaches 44.7%, compared to that of the regular one.

On the identification of cross-flow mode in three-dimensional boundary layers

Parametric zones are obtained where the cross-flow instability can be identified as a mode in the three-dimensional boundary layers with the Mach number ranging from 0 to 10. Although the term cross-flow mode is widely used in the investigations on boundary-layer instability, the previous work [Z. Liu, Phys. Fluids 34, 094110 (2022)] has demonstrated that a cross-flow mode cannot be designated in certain circumstances. The identification of the cross-flow mode is significant not only in the justifiable use of the term but also in judging whether a disturbance is more inclined to a cross-flow type or not. In this work, a criterion is built to identify the cross-flow mode based on the growth-rate peak and disturbance shape. By means of extensive calculations and identifications, parametric zones are presented for the unstable cross-flow mode. It is found that the cross-flow mode cannot be identified at large local sweep or at Mach numbers larger than 1.6. In parametric zones, the cross-flow mode can be distinguished from the Tollmien–Schlichting mode, or identified as the solely dominant mode. Based on the identifications, the maximum growth rates of the cross-flow mode, the Tollmien–Schlichting mode, and the Mack modes in the three-dimensional boundary layers are provided. The dominant mode can be determined at different Mach numbers. The cross-flow mode is revealed to dominate the boundary-layer instability at low Mach numbers. Under large cross-flow strengths, the second mode could not be the dominant mode in the hypersonic boundary layers, which contrasts sharply with the two-dimensional case.

Experimental Investigation of Hypersonic Boundary Layer Transition on a 5° Smooth Straight Cone

Experimental Investigation of Hypersonic Boundary Layer Transition on a 5° Smooth Straight Cone

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

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

Hypersonic Boundary-Layer Disturbances on a Cooled, Slender Cone at Mach 6

This paper presents an experimental study of the effect of wall cooling on the boundary-layer structures of a sharp-nosed, 5° half-angle cone in Mach 6 flow. An experimental methodology was developed that allowed quantitative measurements of second-mode disturbance growth over a range of wall-temperature ratios from without modifying the freestream disturbance environment. The cooling system comprised probe cooling, liquid nitrogen injection, and thermal insulation. The model temperature was monitored before each run, with thermocouples mounted at the base and nosetip piece agreeing within 3.3 K. Calibrated schlieren provided global and time-resolved measurements of the boundary-layer features at a freestream unit Reynolds number of . An increase in wall-cooling level was associated with a shrinking of the boundary layer and a lowering of wavepacket structure angle at the wall. Despite some outlying results, an overall trend of increasing disturbance frequency with decreasing surface temperature was observed; additionally, cooling increased the spectral amplitude and growth rate of second-mode disturbances, pushing the point of saturation upstream.

Experimental study of bluntness effects on hypersonic boundary-layer transition over a slender cone using surface mounted pressure sensors

In this work, we studied the bluntness effect on the hypersonic boundary-layer transition over a slender cone at Mach 6 with interchangeable tips in a noisy Ludwieg tube tunnel before the so-called “transition reversal” phenomenon occurs. The evolution of instability waves is characterized using surface flush-mounted pressure sensors deployed along the streamwise direction within unit Reynolds number from 4E+ 6/m ≤ Reunit ≤ 10E+ 6/m, and the bluntness of the cone nose ranges from 0.1 mm to 5 mm. Power spectral density (PSD) of pressure fluctuation indicates that small nose bluntness (ReR ≤ 2000) has little influence on the evolution of instability waves along the hypersonic boundary-layer, whereas with a moderate nose size (2000 ≤ ReR ≤ 5000), the hypersonic boundary layer transition is delayed monotonically as the nose radius increases before the boundary-layer turns into fully laminar without instability waves. The delaying effect can be attributed to the increased entropy-layer swallowing distance with a large tip radius. Instability wave characterization reveals that the second mode instability wave plays a dominant role before the transition reversal happens. The quadratic phase locking of second mode instabilities can be identified by bispectral analysis, and it attenuates as the nose tip radius increases.

Velocity measurements in a hypersonic flow using acetone molecular tagging velocimetry.

In the present work, a non-intrusive diagnostic technique known as molecular tagging velocimetry was used to collect quantitative freestream velocity measurements in the Mach 7 Ludwieg Tube Wind Tunnel located at The University of Texas at San Antonio. This laser-based diagnostic technique used a single Nd:YAG 4th harmonic laserline to excite acetone molecules seeded in the flow field. From the resulting emitted light, mean and instantaneous velocity profiles in the hypersonic freestream flow and facility boundary layer were measured. Uncertainty in the velocity measurements for individual test runs is estimated at ≤ 8% while overall 1D freestream mean velocity measurements were recorded with ±2.4% (± 21 m/s) accuracy. The effect of acetone seeding on the speed of sound was also quantified.