Hypersonic Boundary-layer Instability Research Articles

Interaction of second-mode instability and distributed roughness elements in the hypersonic boundary layer over a sharp cone

The ablation of the thermal protection system (TPS) on the aircraft surface generates distributed roughness elements (DRE), which is expected to affect the hypersonic boundary-layer instability seriously. This work presents the experimental investigation of hypersonic boundary-layer instability waves on a 7° half-angle sharp cone with DRE allocated downstream of synchronization point at 0° angle of attack using surface flush-mounted pressure sensors, focused laser differential interferometer (FLDI), infrared thermography and schlieren technique. The experimental results showed that when the DRE (lower than the local boundary-layer height) was located downstream of the synchronization point, the Mack second mode still dominate the boundary-layer instability, and the transition onset was not significantly affected by DRE at various Reynolds numbers. Based on further analysis in the region downstream of roughness, the DRE didn't influence the nonlinear interaction among instability waves materially, although it promoted the growth of instability waves compared with the smooth case. However, the nonlinear interactions were decaying adjacent to the DRE, as the Mack second-mode instability waves cannot penetrate into the separation bubble because of the flow separation induced by the DRE. The FLDI measurement along the wall-normal direction revealed the second-mode instability waves were suppressed close to the wall by the complex roughness surface at the DRE region, and also confirmed the existence of second-mode instability waves at these streamwise locations. The evolution of instability waves followed the linear stability theory (LST), despite the existence of complex wall surfaces. This work enriches the understanding of the effect of DRE on hypersonic boundary-layer transition and demonstrates the spatial characteristics of instability waves above the complex surface.

Displacement of hypersonic boundary layer instability and turbulence through transpiration cooling

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.

Experimental investigation of hypersonic boundary layer instability with wavy wall located downstream of synchronization point

The wavy wall surface has been reported to be effective in suppressing the growth of second mode instability waves in hypersonic boundary layer, while the allocation of wavy wall relative to the synchronization point requires further study. Upon this work experimental investigation of hypersonic boundary layer instability with wavy wall located downstream of synchronization point based on 7° half-angle sharp cone at zero angle of attack was carried out in a hypersonic noisy wind tunnel. The evolution of instability waves with wavy wall surface is measured using both surface-mounted pressure fluctuation sensors and focused laser differential interferometer (FLDI). The results reveal that the growth of instability waves has been enhanced significantly when the wavy wall region is allocated downstream of the synchronization point when the wavy wall height is slightly smaller than the local boundary layer thickness. Additionally, second mode instability waves are observed at both crests and troughs within the wavy wall region based on the FLDI measurement and moreover, the high-frequency instability waves measured in the separation bubbles of troughs are speculated to be the transmission of second mode instability waves from the shear layer, rather than the local generation of second mode instability.

Experimental study of the effects of isolated roughness elements on the stability and transition of a hypersonic boundary layer on a flat plate

Surface roughness elements on hypersonic vehicles can cause early boundary layer transition, increasing wall skin friction and heat flux and affecting aircraft range and thermal protection systems. Accurate prediction of the transition caused by these roughness elements is crucial for the design of hypersonic vehicles. In this work, wind tunnel experiments on isolated roughness-induced boundary layer transition at Ma = 6 are conducted. Infrared thermography and high-frequency pressure sensors are utilized to investigate the effects of different roughness element configurations (cylindrical, diamond, ramp) on the hypersonic boundary layer instability and transition. The experimental results show that all three roughness elements can effectively enhance the generation of second mode waves and promote boundary layer transition. Compared to smooth surfaces, they exhibit similar frequency band range, faster growth, and earlier saturation. Among them, the ramp roughness element most effectively triggers the boundary layer transition, with a relatively small heat flux increase. Furthermore, bispectral analysis illustrates that all three roughness elements undergo self-interactions that lead to spectral broadening, ultimately resulting in boundary layer transitions.

Hypersonic boundary-layer instability characterization and transition downstream of distributed roughness

Hypersonic boundary-layer instability characterization and transition downstream of distributed roughness

Stabilization of hypersonic boundary-layer instability using porous coatings under thermochemical nonequilibrium

In this study, the suppression of the instability of the hypersonic boundary-layer transition under thermochemical nonequilibrium (TCNE) using a porous coating configuration is examined through linear stability theory (LST) and direct numerical simulation (DNS). Fedorov's admittance model is applied in LST to account for the porous coating effect. The model is modified accordingly such that the flow condition simulated under the TCNE assumption is considered. The results are compared with the case under the calorically perfect gas (CPG) assumption. Conditions of Mach 20-6 degree wedge flow and Mach 45-10 degree wedge flow are chosen for this study. In general, the second mode instability is destabilized and shifts upstream under the TCNE effect. Through the comparison between the DNS and LST results in the presence of a porous coating, it is concluded that the effect of suppressing the instability is very similar under both the CPG and TCNE assumptions. The TCNE effect has minimal impact on the acoustic attenuation of the instability within the cavity. Furthermore, the use of the admittance models in LST can fairly predict the trend of disturbance growth in DNS if the porous coating region is sufficiently long, validating the feasibility of extending the application of the admittance model to a different thermochemical state in addition to the CPG assumption. Examination of the perturbation contours under different state assumptions reveals the absence of second-mode wave structure in the region of microcavities, confirming the stabilization of hypersonic boundary-layer instability.

Effect of Local Thermal Strips on Hypersonic Boundary-Layer Instability

An effect of localized thermal strips on hypersonic boundary-layer instability is investigated using linear stability theory (LST). The linear evolution of Mack's second mode is analyzed for a Mach 4.5 flat-plate boundary-layer with and without the strips. The validity of the LST analysis is assessed through the comparison of the results with implicit large eddy simulations (ILES) for selected cases. Parametric studies on the temperature intensity, length, and location of a single strip are performed to examine the influence of the strip on the second-mode instability. A reversal in the stabilizing and destabilizing effect according to the local streamwise surface pressure gradient induced by the strip is identified. The reversal point of the stabilizing/destabilizing effect is located in the vicinity of the location of the maximum growth rate of the second mode. The influence of three strips in a series arrangement is also analyzed for several combinations. The proper combination of strips, considering the reversal phenomenon, results in more significant stabilizing/destabilizing effects for a disturbance with a specific frequency. The overall effects are also evaluated and analyzed in terms of the envelope of the N-factor curves by carrying out the LST calculations for disturbances with various frequencies.

Impact of Wavy Wall Surface on Hypersonic Boundary-Layer Instability of Sharp Cone Model

A comprehensive investigation of the impact of wavy wall on hypersonic boundary layers is carried out in a Mach 6 Ludwieg tube based on a 7 deg half-angle sharp cone. Linear stability theory is applied on the cone base flow with both smooth and wavy surfaces first. The evaluation of amplification rate as well as the frequency of instability waves reveals that the second mode instability is suppressed downstream of the wavy wall region. The development of instabilities is then characterized using surface flush-mounted pressure sensors and focused laser differential interferometer (FLDI). The FLDI measurement is conducted within the wavy wall zone both at troughs and at crests. Good agreement is found between the surface pressure measurement and computation, except that instability waves are invisible shortly downstream of the wavy wall. Additionally, bispectral analysis indicates that the nonlinear interaction of instability waves downstream of the wavy wall region is suppressed. FLDI measurement along the streamwise direction shows that the second mode instability waves disappear in the wavy wall region, and the boundary-layer transition process is delayed. Spatial measurement with FLDI of boundary layer identifies various types of disturbance at the troughs and crests; moreover, the co-existence of low-frequency and high-frequency disturbances is observed off the wall surface downstream of the wavy wall.

Double-layer acoustic metasurface for the suppression of the Mack second mode in hypersonic boundary-layer flow

An acoustic metasurface consisting of two layers of perforated plates is proposed for suppression of the Mack second mode in hypersonic boundary-layer flow. The upper layer with very tiny holes is permeable to acoustic waves and hardly alters the background hypersonic boundary-layer flow, offering rather low resistive and inductive components of surface acoustic impedance. The bottom layer with large and sparse square holes is attached to a rigid wall surface and forms a periodic array of chambers, each covering multiple holes of the upper layer, which can adjust the impedance phase by working as a capacitive component. Based on a linear stability analysis of hypersonic boundary-layer instability, such an acoustic metasurface satisfying the required surface impedance is designed and numerically investigated. The results show that the metasurface can efficiently suppress the Mack second mode over a relatively wide bandwidth. This work provides an alternative strategy for the design of porous walls for hypersonic boundary-layer stabilization.

Effect of mass injection on secondary instability of hypersonic boundary layer over a blunt cone

In low environmental disturbances, secondary mechanisms play a crucial role in flow instability and transition. Over the years, researchers have shown the existence of secondary waves prior to nonlinear breakdown and turbulence. In high-speed flows, the vehicles are subjected to extreme thermal loads, and usually, an ablative heat shield is used for protection. The ablative materials eject gaseous products upon heating, significantly affecting flow stability. Mass injection through the surface in the boundary layer loosely replicates the ablation process. Its effect on stability and transition has been explored earlier using experiments, numerical simulation, and Linear Stability Theory (LST). However, the effect of the surface injection on secondary instability, the most viable path to transition, remains uncertain. The present work studies the secondary instability of a hypersonic boundary over a 7° half-angle blunt cone in the presence of mass injection through the surface of the boundary layer. Mean flow over the cone is solved using a high-order shock fitting direct numerical simulation code. Primary instability is studied using the LST, and secondary instability is studied using the secondary instability theory based on the Floquet model. Computations are carried out for different injection rates, and it is found that instability increases with mass injection rate. The mass injection has increased the primary and secondary growth rates. Fundamental modes are more dominant than subharmonic and detuned modes at higher primary wave amplitudes. The mass injection has increased the primary and secondary N-factor, and transitioning behavior is observed at the maximum injection rate.

Experimental Measurements of Hypersonic Instabilities over Ogive-Cylinders at Mach 6

Hypersonic boundary-layer instability disturbance and transition experiments were performed in the Air Force Research Laboratory Mach-6 Ludwieg Tube on a 1-m-long ogive-cylinder model with interchangeable nose-tip geometries of varying ogive radii and nose bluntnesses. Boundary-layer disturbances were measured via focused laser differential interferometry, surface-mounted pressure sensors, and high-speed schlieren videography at unit Reynolds numbers ranging from to . Three primary disturbance structures were observed that were dependent on ogive radius and nose bluntness. Each disturbance structure exhibits unique characteristics that are closely tied to the leading-edge geometry of the model. The addition of nose bluntness is shown to have an overall delay in transition onset, and a possible blunt nose transition reversal is observed as bluntness is further increased. It is also shown that the effects of the ogive radius in addition to nose bluntness alter the flow instability disturbance and transition behaviors and are likely tied to a modification of the entropy layer structure. A novel curvature-based Reynolds number is proposed that successfully correlates the effects of multiple leading-edge radii of curvature to dominant downstream boundary-layer disturbances and transition onset locations.

Effect of three-dimensional smooth humps on hypersonic boundary layer instability of streamwise vortices over a yawed cone

This paper describes the design of a class of three-dimensional (3D) smooth humps mounted on a yawed cone model for a high-precision compressible Navier–Stokes solver. The influence of isolated 3D humps with different heights on the linear stability of streamwise vortices over a yawed cone with a 7° half-angle at a 6° angle-of-attack, free-stream Mach number of 6, and unit Reynolds number of 1.0×107/m is investigated. The results show that smooth humps induce stationary streaks inside the streamwise vortices over a yawed cone, effectively increasing the energy of the inward vortex and inhibiting the evolution and development of the outward vortex. Stability analysis shows that the induced streaks enhance the inner mode instability, suppress the outer mode instability, and reduce the dominant frequency of the outer mode. The stability properties of the hump-induced streaks are studied, and the physical process of the transformation between the new streak instability mode and the original boundary layer mode is revealed. The eN method based on global stability theory is used to predict the transition location along the centerline on the leeward side of the cone. It is found that configurations with hump heights of 0.153 and 0.267 times the local boundary layer thickness result in a transition delay of ∼3% and ∼11%, respectively.

Interaction of second-mode disturbances with an incipiently separated compression-corner flow

Abstract

Secondary subharmonic instability of hypersonic boundary layer in thermochemical equilibrium over a flat plate

The transition from laminar to turbulent flow is a complex phenomenon and difficult to predict. There have been significant contributions over the years, enhancing the understanding of mechanisms that cause transition. The effect of chemical reactions on boundary-layer instability becomes significant for high-temperature flow. The existence of the secondary instability before the transition in a low-disturbance environment has been proved in earlier studies, although none of these studies considered high-temperature effects. The real gas (identified as a gas undergoing high-temperature phenomena) effects on the secondary subharmonic instability for a two-dimensional hypersonic boundary layer over a flat plate are studied using the Floquet model. The real gas is assumed to be in thermochemical equilibrium, where air is considered as a mixture of perfect gases in local thermal and chemical equilibrium. A five-species air model is used to compute the air mixture's thermal and transport properties. The stability of the primary disturbance is analyzed using linear stability theory, whereas a mean flow superposed with a second mode primary disturbance has been considered for the secondary instability analysis. The local and global methods have been used to solve the eigenvalue problem based on the finite difference method. The real gas effect on the secondary instability is destabilizing since the maximum secondary amplification rate is higher and appears at a lower Reynolds number for the real gas than that of the perfect gas. The increase in the primary wave amplitude increases the secondary amplification rate significantly. Furthermore, secondary disturbances decay faster into the freestream compared to primary disturbances. The eN method is employed to compute the amplification factor for the secondary amplification rate for the real and perfect gas. Based on the growth rate and N-factor for the secondary disturbances, it is found that the real gas flow is more unstable and susceptible to transitioning earlier than the perfect gas.

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.

Growth mechanisms of second-mode instability in hypersonic boundary layers

Abstract

Linear, nonlinear and transitional regimes of second-mode instability

Abstract

The influence of wavy wall on hypersonic boundary layer instability over a flared cone

The influence of wavy wall on the hypersonic boundary layer instability and the related aerodynamic heating are investigated over a flared cone. Experiments are conducted in a Mach 6 wind tunnel using Rayleigh-scattering flow visualization, fast-response pressure sensors and infrared thermography. Three sizes of wavy wall have been placed on the model in comparison with the smooth wall cone. The results show that compared to the smooth-wall cone, all the wavy-wall cones can suppress the second-mode instability to a certain degree and eliminate the local heating spot before transition is completed. Also, it is found that with the decrease of wavy wall wavelength, the transition location is shifted downstream.

Toward robust prediction of crossflow-wave instability in hypersonic boundary layers

Abstract Traveling and stationary crossflow-wave instabilities in a laminar Mach 6 boundary layer are investigated on a 38.1% scale model of the Fifth Hypersonic International Flight Research Experiment (HIFiRE-5) elliptic cone at zero angle of attack and yaw. The Langley Stability and Transition Analysis Code (LASTRAC) was used to analyze the crossflow dominated boundary layer in the mid-span region near the downstream end of the model. Disturbance growth rates, wave angles, and phase speeds are computed with LASTRAC using quasi-parallel Linear Stability Theory (LST), Linear Parabolized StabilityEquations (LPSE), and two-plane or surface marching LPSE (2pLPSE). The predicted wave angles and phase speeds are validated using experimental data, and are found to be in better agreement than previous computations. Further numerical analysis is conducted using the Spatial BiGlobal technique (SBG), which simultaneously accounts for wall-normal and spanwise gradients in the mean boundary layer at a particular axial station. For the first time in the literature, a comparison is made between crossflow wave growth rates computed using LST, LPSE, and 2pLPSE and those computed using SBG, accounting for curvature and geometric divergence of the elliptic cone. The agreement between LST, LPSE, and SBG is fair at best, but excellent agreement is realized between 2pLPSE and SBG. This result constitutes a co-verification of the LASTRAC and SBG stability codes, and provides evidence that 2pLPSE accurately models the physics of the traveling-crossflow instability.

Improved algorithm for solving nonlinear parabolized stability equations**Project supported by the National Natural Science Foundation of China (Grant Nos. 11332007 and 11402167).

Due to its high computational efficiency and ability to consider nonparallel and nonlinear effects, nonlinear parabolized stability equations (NPSE) approach has been widely used to study the stability and transition mechanisms. However, it often diverges in hypersonic boundary layers when the amplitude of disturbance reaches a certain level. In this study, an improved algorithm for solving NPSE is developed. In this algorithm, the mean flow distortion is included into the linear operator instead of into the nonlinear forcing terms in NPSE. An under-relaxation factor for computing the nonlinear terms is introduced during the iteration process to guarantee the robustness of the algorithm. Two case studies, the nonlinear development of stationary crossflow vortices and the fundamental resonance of the second mode disturbance in hypersonic boundary layers, are presented to validate the proposed algorithm for NPSE. Results from direct numerical simulation (DNS) are regarded as the baseline for comparison. Good agreement can be found between the proposed algorithm and DNS, which indicates the great potential of the proposed method on studying the crossflow and streamwise instability in hypersonic boundary layers.