Hypersonic Boundary Layer Transition Research Articles

Effect of surface cooling on second-mode dominated hypersonic boundary layer transition

Effect of surface cooling on second-mode dominated hypersonic boundary layer transition

Early stage evolution of naturally developing Görtler streaks

This is a companion paper to a previous study of hypersonic boundary layer transition over a concave wall [Huang et al., “Inner structures of Görtler streaks,” Phys. Fluids 33, 034116 (2021)]. Experiments are performed in a Mach 6.5 quiet wind tunnel using CO2-enhanced filtered Rayleigh scattering flow visualization, PCB fast-response pressure sensors, and a high-frequency Schlieren technique to investigate naturally developing Görtler streaks. This is the first time that naturally developing Görtler streaks have been visualized using the Rayleigh scattering technique in a hypersonic flow. The results show that three-dimensional waves can be clearly visualized inside the Görtler streaks. More importantly, these waves are amplified owing to Görtler instability, contributing to the formation of Görtler streaks and mushroom-like structures. Thus, the three-dimensional waves can be considered to be basic structures inside Görtler streaks and to play a crucial role in boundary layer transition.

Wall Cooling Effect on High-Enthalpy Supersonic Modes over a Cone

High-enthalpy hypersonic boundary-layer transition plays an important role in many entry/descent vehicles. The aerothermodynamic performance of these vehicles strongly depends on the transition location on the surface. However, detailed transition flow physics in these chemically reacting boundary layers are poorly understood and transition estimates during the design phase rely heavily on empirically derived transition criteria such as . One of the most intriguing characteristics in hypersonic boundary layers is the presence of unstable supersonic modes, first identified in the 1990s. Due to a recent surge in hypersonic applications, there has been renewed interest in studying the flow physics of supersonic modes using either theory or direct numerical simulations. This paper investigates the rise of supersonic modes in a high-enthalpy hypersonic flow over a 5 deg half-angle cone at various wall temperatures using both quasi-parallel linear stability theory and linear parabolized stability equations (PSEs). It was found that supersonic modes exist in all wall temperature conditions, including the adiabatic wall case. A cooler wall temperature causes the second Mack mode to become an unstable supersonic mode naturally downstream of the upper-branch neutral location when the wall is sufficiently cooled. In terms of the integrated growth, the second mode is still the dominant mode. Nonetheless, supersonic modes can cause additional series of relatively weaker growth than that of the second mode beyond the peak amplitude location. According to the present linear PSE results, the formation of unstable supersonic modes is mainly associated with the synchronization of phase speed between the instability and acoustic waves in a nonparallel boundary layer, and not due to nonlinear modal interaction as suggested in the literature.

Theoretical, numerical and experimental study of hypersonic boundary layer transition: Blunt circular cone

• Hypersonic boundary layer transition is studied using various approaches. • Leading edge receptivity is crucial for understanding “transition reversal”. • Centreline and crossflow vortices significantly promote transition on a yawed cone. • The newly proposed C - γ - Re θ transition model works well at various flow conditions. • Valuable transition data is acquired in real flight test for a cone model. This article reviews the hypersonic laminar-turbulent boundary layer transition over a blunt circular cone under support of the National Key Research and Development Program of China (contract No. 2016YFA0401200). Various approaches, including theoretical, numerical, experimental and flight test, are combined to obtain a complete perspective of the transition process as well as some fundamental issues for this kind of flow. The interaction between freestream disturbances and the front shock wave is analysed theoretically, and multiple receptivity routes are unveiled using a sub-zone analysis technique. The entire transition process is captured using high-resolution nano-tracer-based planar laser scattering technique in a quiet windtunnel as well as large-scale high-fidelity direct numerical simulations. Using dynamic mode decomposition, the relationship between the second Mach mode and the low-frequency mode in the late stage of the transition is revealed. The “transition reversal” phenomenon is recovered in two of our conventional windtunnels. The numerical results show that leading-edge receptivity varies significantly for different cone bluntness, which might be responsible for the “transition reversal” phenomenon. For the cone at angles of attack, new instabilities including the leeward centerline vortex instability and the crossflow instability, occur and can significantly promote the transition process. A newly developed C - γ - Re θ transition model can give a good transition front prediction compared to the experimental data. The flight test for a cone model conducted within this project is reviewed, and the correlation between the ground experimental data and the flight data is established.

Inner structures of Görtler streaks

The hypersonic boundary layer transition over a concave wall is investigated in a Mach 6.5 quiet wind tunnel using temperature sensitive paint (TSP), CO2-enhanced filtered Rayleigh scattering flow visualization, PCB fast-response pressure sensors, and a high-frequency schlieren technique. The TSP shows that low- and high-temperature streaks are distributed in the spanwise direction. The wavelengths of naturally developing Görtler streaks are randomly distributed, with an average of approximately 7 mm, and change little as the unit Reynolds number increases. More importantly, three-dimensional waves are clearly visualized and quantitatively measured inside the Görtler streaks. This is the first time that the entire evolution of the Görtler instability has been visualized using the Rayleigh-scattering flow visualization in hypersonic flow. The results demonstrate that three-dimensional waves are amplified as a result of the Görtler instability, resulting in a localized high-shear layer around the interface of the three-dimensional waves, which contributes to the formation of hairpin vortices and mushroom-like structures. The three-dimensional waves grow and play an important role in Görtler instability-induced boundary layer transitions.

Secondary instability of the hypersonic high-enthalpy boundary layers with thermal–chemical nonequilibrium effects

Compared to the widely studied linear instabilities, the nonlinear analyses of the high-enthalpy and hypersonic boundary-layer transition had received much less attention. In this work, secondary instabilities of high-enthalpy boundary layers are studied to highlight the impacts of thermal–chemical nonequilibrium (TCNE). The numerical tools adopted here are the nonlinear parabolized stability equation (NPSE) and the Floquet analysis (or the secondary instability theory, SIT). The flow over a blunt cone is computed at a free-stream Mach number of 20. The fundamental resonance is found to dominate over the subharmonic resonance. The development of secondary waves from SIT agrees well with that from NPSE, suggesting satisfactory cross-validation of both solvers. It is also illustrated that the disturbances of the species mass and vibrational temperature are mainly generated by the product term of the wall-normal velocity disturbance and the mean-flow gradient. The typical rope-like structures are observed near the boundary-layer edge. The effects of TCNE on the secondary instability are to cause larger maximum secondary growth rate and the corresponding azimuthal wavenumber.

Hypersonic boundary layer receptivity on flat plate with blunt leading edge due to acoustic disturbance

In the process of hypersonic boundary layer transition, the leading edge receptivity due to free-stream acoustic disturbance provides unstable perturbations in boundary layer with initial amplitudes, which is one of hot issues in the study about transition mechanism. The receptivity of leading edge is influenced by many factors, involving geometry parameters, disturbance types, etc., among which the bluntness plays a key role for generating disturbances. In order to reveal the mechanism of bluntness influence on the boundary layer transition, investigations into the acoustic receptivity on flat plates with sharp and blunt leading edge are carried out. The flow conditions are as follows: Mach number 6, Reynolds number 1e7/m, adiabatic wall. By introducing two-dimensional plane fast and slow acoustic waves into the free-stream, the whole receptivity process of free-stream disturbances passing through shock wave into boundary layer is simulated with high-order accuracy DNS method. With sharp leading edge as a reference, the generation mechanism of the perturbations near blunt leading edge and its influence on the boundary layer downstream are analyzed. Some conclusions are obtained: for acoustic disturbances in hypersonic free-streams, the receptivity mechanism of blunt leading edge is different from that of sharp leading edge; both the fast and the slow mode can be produced respectively near the sharp leading edge by the fast and the slow acoustic wave, while only fast mode appears near the blunt leading edge; the receptivity of blunt leading edge due to the fast acoustic is stronger than that due to slow acoustic, and is also stronger than that of the sharp leading edge in a certain region of boundary layer downstream.

Application of focused laser differential interferometer to hypersonic boundary-layer instability study

Accurate prediction of hypersonic boundary-layer transition plays an important role in thermal protection system design of hypersonic vehicles. Restricted by the capability of spatial diagnostics for hypersonic boundary-layer study, quite a lot of problems of hypersonic boundary-layer transition, such as nonlinearity and receptivity, remain outstanding. This work reports the application of focused laser differential interferometer to instability wave development across hypersonic boundary-layer on a flared cone model. To begin with, the focused laser differential interferometer is designed and set up in a Mach number 6 hypersonic quiet wind tunnel with the focal point in the laminar boundary-layer of a 5 degree half-angle flared cone model. Afterwards, instability experiments are carried out by traversing the focal point throughout the hypersonic boundary-layer and the density fluctuation along the boundary-layer profile is measured and analyzed. The results show that three types of instability waves ranging from 10 kHz to over 1 MHz are co-existing in the hypersonic boundary-layer, indicating the powerful capability of focused laser differential interferometer in dynamic response resolution for instability wave study in hypersonic flow regime; furthermore, quantitative analyses including spectra and bicoherence analysis of instability waves throughout the hypersonic boundary-layer for both cold and heated cone models are performed.

DNS of Hypersonic Boundary Layer Transition

極超音速流における境界層の乱流遷移では一次モード（TS波）よりも一ケタ高い周波数（数百kHz）の二次モード変動が支配的な役割を果たすと考えられている．近年，二次モード変動が低周波の変動と干渉しながら成長する過程が極超音速風洞において観測されているが，従来の線形安定性解析ではこれを説明することができず，非線形効果を考慮した新たな理論の確立が必要である．筆者の研究グループでは，低周波モードと干渉する二次モードの成長過程を直接数値シミュレーション（DNS）により捉え，非線形理論を構築するために必要なデータを取得することを目指している．本稿では，エネルギー生成項を介して境界層内部に10kHzの低周波温度擾乱を付加して乱流遷移を発生させることを試みた結果，付加した周波数とは異なる高周波の二次モードが励起されて乱流遷移する過程がDNSによって捉えられた．

Boundary Layer Transit Flight Experiment: Mission Overview, Launch Vehicle and Payload Subsystems

The objective of the Boundary Layer Transition (BOLT) flight experiment is to investigate the hypersonic boundary-layer transition mechanisms on a low-curvature concave surface with a swept leading edge at Mach numbers between five and seven. This shall be achieved during a captive-carry flight experiment on an S31/Improved Orion sounding rocket. The BOLT project is coordinated by the U.S. Air Force Research Laboratory U.S. Air Force Office of Scientific Research and will be carried out in collaboration with the U.S. Air Force Research Laboratory Aerospace Systems Directorate, the Johns Hopkins University Applied Physics Laboratory, and a multiorganization team of researchers. DLR, German Aerospace Center’s Mobile Rocket Base (DLR MORABA) is responsible for the development of the required mission and flight profile, layout and adaption of the launch vehicle, as well as the provision of payload subsystems for attitude control, time synchronization, data telemetry, despin, and separation. DLR MORABA will complete ground testing, coordinates, and conducts launch operations; and it will support postflight analysis. This paper presents these contributions in detail.

Influences of Steps on the Hypersonic Boundary-Layer Transition on a Cone

The influences of steps on the hypersonic boundary-layer transition on a cone are investigated in a hypersonic quiet wind tunnel, using high-frequency pressure sensors and nano-tracer-based planar laser scattering (NPLS) techniques. The high-frequency surface pressure fluctuations are analyzed using power spectral density analysis, band-pass filtering, and cross-correlation calculation. Quantitative analysis of the NPLS images is carried out based on the grayscale information and time interval. The results demonstrate that the second-mode wave amplitude increases initially and then drops along the streamwise, whereas the characteristic frequency decreases monotonously. However, the second-mode wave amplitude in the 0.4 mm () and 0.6 mm () forward-facing step models is constantly rising, and the obvious second-mode wave first appears closer to the downstream than the smooth cone, which suggests that the forward-facing steps could suppress the second-mode disturbances and shifts the transition downstream. Nonetheless, there is only a clear promotion effect when the height of the forward-facing step is 1.6 mm (). In addition, the backward-facing step could destabilize the boundary layer and shift the transition upstream, and this effect is enhanced by increasing the step height.

Aerothermodynamic Ground Tests in Support of the Boundary Layer Transition Flight Experiment

The U.S. Air Force Research Laboratory/U.S. Air Force Office of Scientific Research has sponsored the Boundary Layer Transition (BOLT) experiment to investigate hypersonic boundary-layer transition on a low-curvature, concave surface with swept leading edges. This paper provides a review of aerothermodynamic ground-test contributions by NASA Langley Research Center to the design of the BOLT flight experiment. Several test entries into the NASA Langley Aerothermodynamics Laboratory 20 Inch Mach 6 Air Tunnel are discussed. Global surface heating distributions on subscale BOLT models were measured using either phosphor thermography or infrared thermography at a range of model attitudes and freestream Reynolds numbers. An initial test entry provided the first experimental measurements of transition on the BOLT geometry. A second test entry investigated the effects of distributed surface roughness along the swept leading edges. A third test entry provided an evaluation of the aeroheating environment on the aerodynamic fairings and a portion of the flight vehicle downstream of the BOLT experiment. These entries were intended to support the development and design of flight hardware and instrumentation for the BOLT flight experiment.

Final Design of the Boundary Layer Transition (BOLT) Flight Experiment

To improve scientific knowledge of boundary-layer transition physics on increasingly complex geometries, the Air Force Research Laboratory/Air Force Office of Scientific Research has sponsored the Boundary Layer Transition (BOLT) Experiment to obtain measurements of hypersonic boundary-layer transition in flight on a low-curvature concave surface with highly swept leading edges. This paper will summarize the completed design of the BOLT flight experiment. The final selection and layout of the flight scientific instrumentation was driven by extensive ground testing and numerical analyses. The complex flight experiment hardware was designed with over 344 scientific measurements to characterize transition in flight, and to provide a new validation dataset for the development of next-generation transition prediction methods.

Convergence acceleration for high-order shock-fitting methods in hypersonic flow applications with efficient implicit time-stepping schemes

The research on hypersonic boundary layer transition is a recent focal point in fluid dynamics. One of the effective numerical methods for transition study is the high-order shock-fitting method, which is especially suitable for dealing with shock-related interactions and removing post-shock oscillations as compared with shock-capturing methods. Currently, explicit time-stepping schemes are mainly used for shock-fitting methods. In comparison, implicit schemes have received less attention due to the extra complexity arising from dynamic shock and moving mesh, even though they can accelerate the convergence speeds for steady flows by up to an order of magnitude. In this paper, two implicit time-stepping methods, GMRES (generalized minimum residual) and LR (line-relaxation), are realized for the efficient calculation of steady flow fields using the 5th-order shock-fitting method. The gas models include both calorically perfect gas and 5-species thermal-chemical non-equilibrium gas required for high-speed and high-enthalpy flows. The results of test cases showed that the convergence rates of GMRES+LR and LR are 4 to 20 times faster than the explicit/semi-implicit Runge-Kutta schemes. Therefore, the usage of implicit GMRES and LR as time-stepping schemes raises the computational efficiency for obtaining steady boundary layer flows, which helps the important high-order shock-fitting methods play a greater role in the study of hypersonic boundary layer stability and transition.

The effect of concave roughness on the hypersonic flat-plate boundary-layer transition

In order to reveal the effect of aircraft surface concave roughness on the transition, several hypersonic quiet wind tunnel tests were carried out on the flat-plate model with independent concave roughness on the surface. The experimental Mach number was 6.5 and the total temperature was 430 K. Using infrared thermal imaging technology, a general rule of the effect of concave roughness on the boundary layer transition was obtained. The flow visualization results and the PCB pressure measurement results showed that the concave roughness promotes the transition by increasing the growth of the first mode wave rather than the growth of the second mode wave.

Mechanism for frustum transition over blunt cones at hypersonic speeds

Numerical and experimental studies have demonstrated laminar–turbulent transition in hypersonic boundary layers over sharp cones via the modal growth of planar Mack-mode instabilities. However, due to the strong reduction in Mack-mode growth at higher nose bluntness values, the mechanisms underlying the observed onset of transition over the cone frustum are currently unknown. Linear non-modal growth analysis has shown that both planar and oblique travelling disturbances that peak within the entropy layer experience appreciable energy amplification for moderate to large nose bluntness. However, due to their weak signature within the boundary-layer region, the route to transition onset via non-modal growth of travelling disturbances remains unclear. Nonlinear parabolized stability equations (NPSE) and direct numerical simulations (DNS) are used to identify a potential mechanism for transition over a 7-degree blunt cone that was tested in the AFRL Mach-6 high-Reynolds-number facility. Specifically, computations are conducted to study the nonlinear development of a pair of oblique, unsteady non-modal disturbances in the regime of moderately blunt nose tips. Excellent agreement was demonstrated between the NPSE and DNS predictions. Results reveal that, even though the linear non-modal disturbances are primarily concentrated outside the boundary layer, their nonlinear interaction can generate stationary streaks that penetrate and amplify within the boundary layer, eventually inducing the onset of transition via the breakdown of these streaks. The results indicate that a pair of oblique, controlled non-modal disturbances can produce transition at the location measured in the experiment when their initial amplitude is chosen to be approximately 0.15 % of the free-stream velocity.

Review of progress in some issues and engineering application of hypersonic boundary layer transition

A variety of complicated flow phenomena such as shock wave, boundary layer, flow separation, rarefied gas effects and real gas effects exist in the flows over hypersonic flight vehicles. The hypersonic boundary layer transition is not only a basic aerodynamics issue, but also a hot topic and challenging subject in the research of hypersonic flows. The accurate prediction and the effective control of boundary layer transition can achieve meticulous design of aerodynamic force and heating, thus improving the vehicle performance and enhancing the mission capability. In this paper, the research progress in typical instability characteristics commonly existed in the flight vehicles with complex configurations in engineering is reviewed, and the complicated issues of boundary layer transition in engineering practice which are urgent to be solved are proposed. Meanwhile, the engineering application direction of the research on hypersonic boundary layer transition is specified. Moreover, the investigations on flow control of hypersonic boundary layer are reviewed in order to realize the boundary layer control in the design of hypersonic vehicle in future to enhance the flight performance.

Injection-gas-composition effects on hypersonic boundary-layer transition

The thermal protection system in atmospheric-entry and hypersonic-cruise vehicles are oftentimes designed to ablate during their operation, thus injecting a mixture of various gases with distinct properties into the boundary layer. Such outgassing affects the propagation of instabilities within the boundary layer that ultimately originates the transition to turbulence. This work uses linear stability theory, in combination with the e method, to establish the underlying reason for the experimentally observed advancement/delay of transition in sharp slender hypersonic cones, when injecting lighter/heavier gases. Contrary to the current understanding and experimental correlations, this numerical analysis suggests that such a behaviour is not linked to the isolated effect of the injected gas’ molar weight, but to its combination with the blowing discontinuity, porosity and the appearance of a shocklet, consequence of the injected gas composition. The shocklet constitutes a density gradient that acts on second-mode instabilities like a thermoacoustic impedance.

Experimental study of hypersonic boundary layer transition on a permeable wall of a flared cone

The effects of an alternative permeable material on the hypersonic boundary layer transition are investigated. The new permeable material is shown to be effective in delaying the transition, although the second mode grows faster on the material surface. Experiments are conducted on a flared cone using Rayleigh-scattering flow visualization, fast-response pressure sensors, and infrared thermography. On the permeable wall, the second mode appears earlier and persists over a longer distance along the flow direction. By applying bicoherence analysis, it is determined that the second mode decays more slowly on the permeable wall due to weaker nonlinear interactions.

Experimental study on effect of transverse groove with/without discharge hole on hypersonic blunt flat-plate boundary layer transition

Experiments are carried out on the <i>Φ</i>2 m Shock Tunnel (FD-14A) at the China Aerodynamics Research and Development Center to study the effect of the transverse groove with/without discharge hole on the hypersonic blunt flat-plate boundary layer transition, and the preliminary computational and theoretical research are carried out. The inflow Mach number of the test is 6, the unit Reynolds number is 3.3 × 10<sup>7</sup>/m, the leading edge radius of the flat-plate is 1 mm, and the angle of attack is –4°. Three different sets of two-dimensional transverse grooves are arranged at 110 mm away from the leading edge of the flat-plate. The width and depth of the grooves are, respectively, 2.5 mm and 1 mm for groove 1, 3.75 mm and 1.5 mm for groove 2, and 5 mm and 2 mm for groove 3, at the same time, both ends of the groove 1 can open the discharge hole (the discharge hole has a size of 2.5 mm × 5.0 mm and a width the same as that of groove 1). The discharge hole is denoted as the groove 4, and the smooth flat-plate when the groove is not included is denoted as groove 5 or the flat. The <i>Φ</i>2-mm-diameter cylindrical heat flux sensor is used to measure the heat flux distributions of the center line of the flat-plate under different conditions, and thus we can judge the transition of the boundary layer. The measurement results show that the smooth plate starts to transit at <i>x</i> ≈ 340 mm, and the transition is nearly completed at <i>x</i> ≈ 425 mm. The groove causes the transition position of the boundary layer of the plate to advance, and as the width and depth of the groove increase, the promoting effect on the transition is enhanced, and the transition position moves upstream. After the groove 1 is added to the discharge hole (groove 4), the heat flux distribution and the transition position are substantially the same as those of the smooth plate. After the boundary layer flow completely transits into turbulent flow, the difference in heat flux for each case is small, which indicates that the grooves of different specifications affect only the heat flux distribution in the transition process, but have little effect on the heat flux of the turbulent wall after the transition. The computational fluid dynamic results show that the discharge holes cause passive suction, and the test results show that the suction effect of the discharge holes at both ends of the groove counteracts the effect of the groove on the transition of the center line boundary layer, but it may be just a coincidence, and further research is needed. The linear stability theory (LST) and the optimal perturbation method are used to analyze the flow instability mechanism of the smooth blunt plate. The LST results show that there is no first mode instability nor second mode instability in the blunt plate flow. The modal instability mechanism cannot explain the observed transition in the test. The optimal disturbance calculation shows that the blunt plate flow suffers strong non-modal instability, which can qualitatively explain the observed transition phenomenon.