Wave Boundary Layer Research Articles

Drag Reduction and Energy Saving by Spanwise Traveling Transversal Surface Waves for Flat Plate Flow

Wall-resolved large-eddy simulations are performed to study the impact of spanwise traveling transversal surface waves in zero-pressure gradient turbulent boundary layer flow.Eighty variations of wavelength, period, and amplitude of the space- and time-dependent sinusoidal wall motion are considered for a boundary layer at a momentum thickness based Reynolds number of {rm Re}_theta = 1000. The results show a strong decrease of friction drag of up to 26% and considerable net power saving of up to 10%. However, the highest net power saving does not occur at the maximum drag reduction. The drag reduction is modeled as a function of the actuation parameters by support vector regression using the LES data.A dependence of the spanwise pressure drag on the wavelength is found. A substantial attenuation of the near-wall turbulence intensity and especially a weakening of the near-wall velocity streaks are observed. Similarities between the current actuation technique and the method of a spanwise oscillating wall without any normal surface deflection are reported. In particular, the generation of a directional spanwise oscillating Stokes layer is found to be related to skin-friction reduction.

Improvement of drainage structure and numerical investigation of droplets trajectories and separation efficiency for supersonic separators

As a new type of separation device, supersonic separator has received more and more attention in many areas. However, the interaction between boundary layer and shock wave in supersonic section results in the reverse pressure gradient and makes the boundary layer thicken, which form a so-called "second throat" that reduces the actual flow area and causes the fluid to be chocked. A novel cylinder drainage structure with larger drainage area was developed. When shock wave occurs, the high pressure will make more fluid enter into the openings on cylinder drainage structure, which is equivalent to increase the flow capacity. Therefore, the shock wave is weakened and the total pressure loss is reduced. Moreover, the prediction model of discrete phase coupling verified by experiments was adopted in order to predict the trajectory of droplets and separation efficiency for supersonic separator. The results show that with the increase of droplet size, the separation efficiency increases gradually and the particle diameter of 2–4 μm is sensitive to the separation. The droplets trajectory is more clear and regular in the nozzle with reflux channel and cylindrical drainage structure.

Explicit approximation for velocity and sediment flux above mobile sediment bed beneath current and asymmetric wave

An explicit approximation is proposed for velocity distribution and sediment flux above mobile sediment bed beneath current and asymmetric wave. The predicted velocity is decomposed into an oscillatory part and a current part such that the interaction between current and asymmetric wave can be performed directly. The oscillatory part consists of a free stream velocity and a defect function which includes phase lead, mobile sediment bed level and wave boundary layer thickness. The oscillatory part is influenced by the current due to the changes of both mobile sediment bed level and wave boundary layer thickness. The current part is influenced by a wave eddy viscosity in the wave boundary layer and by an apparent wave roughness of the combined wave-current flow outside the wave boundary layer. Eight free variables are required for the approximation, and iterative bottom shear stress and roughness height are applied for the wave boundary layer thickness and mobile sediment bed level. The mobile sediment bed level takes account of the effects of phase lag (phase shift and residual), acceleration and flow asymmetry. Therefore, the mobile bed effect, instantaneous velocity, net velocity and sediment flux in the wave boundary layer can be reasonably presented by the explicit approximation beneath current and asymmetric wave.

Visualization of hypersonic incident shock wave boundary layer interaction

The incident shock interactions with the hypersonic laminar and forced turbulent boundary layer are visualized by the planar laser scattering technique. The effect of interaction strength on flow structures is also investigated. The results show that the boundary layer shape has been greatly altered by the incident shock. In both the laminar and turbulent inflows, the boundary layer thickness has an abrupt decrease at the interaction region. In laminar inflow, the boundary layer transition rapidly takes place due to the incident shock. The greater the incident shock angle, the bigger the boundary layer thickness downstream the incident shock. In turbulent inflow, the thickness downstream the shock is less than that of the inflow. Fractal analysis is firstly carried on hypersonic shock boundary layer interaction. The effect of the shock angle and inflow condition on fractal dimension is studied. The results indicate that with the stronger shock, the fractal dimension value is bigger for both laminar and forced turbulent inflows.

Numerical modelling of shock wave-boundary layer interaction control by passive wall ventilation

The normal shock wave-boundary layer interaction (SBLI) phenomenon is known to constitute a main factor limiting the aerodynamic performance in many aeronautical applications (transonic wings, helicopter rotor blades, compressor and turbine cascades). The interaction process highly disturbs the boundary layer, often causing flow separation and onset of large scale unsteadiness (e.g. airfoil buffet or supersonic inlet buzz). In certain conditions it may also initiate a dramatic increase of acoustic emission levels (e.g. high-speed impulsive noise). To limit the negative impact of the phenomenon various flow control strategies are implemented, here in a form of a passive control system realised by placing a shallow cavity covered by a perforated plate just beneath the shock. Details of the flow structure obtained by this method are studied numerically. Three distinctive experimental set-ups are considered with the interaction taking place: on a flat wall (transonic nozzle, ONERA), on a convex wall (curved duct, University of Karlsruhe), and on an airfoil (NACA 0012, NASA Langley). Depending on the relative cavity length the ventilation process leads to a transformation of the normal shock topology into: a large λ-foot structure (classical, short cavity), a system of oblique waves (extended cavity), or a gradual compression (full-chord perforation). The reference and flow control cases are simulated with the SPARC code (RANS) with Spalart–Allmaras turbulence and Bohning–Doerffer transpiration models. The results are compared with the measurements, emphasizing the streamwise evolution of the boundary layer profiles and integral parameters during the interaction. The prediction capabilities of the solver in terms of the shock wave-boundary layer interaction control by wall ventilation are assessed and presented in details for the investigated range of flow configurations and conditions.

Impact of different sea surface roughness on surface gravity waves using a coupled atmosphere–wave model: a case of Hurricane Isaac (2012)

The complicated pattern of the chaotic ocean surface depends strongly on the interaction between wind and waves. An accurate representation of momentum and energy exchange at air–sea interface is very important for ocean modeling and climate studies. The exchange of momentum, heat, and moisture at the air–sea interface considered a fundamental process in the development of mesoscale atmospheric phenomena such as hurricanes and precipitating systems. Since these exchanges take place in the wave boundary layer, one cannot neglect the importance of ocean surface waves in modifying air–sea interaction processes. In this paper, the sensitivity of a regional coupled atmosphere–wave model to sea surface roughness parameterizations is discussed. In order to see the impact of sea spray in modifying the surface gravity waves, we have introduced a sea-spray parameterization in surface layer scheme of the atmospheric model. We have taken the case of Hurricane Isaac, formed in the Gulf of Mexico in 2012 for this study. It is observed that the newly added sea-spray scheme shows a better performance compared with three existing, well-known formulations. Thus, during extreme events, sea spray can play an important role in modifying the wind and wave height.

Nonlinear stability of the boundary layer and rarefaction wave for the inflow problem governed by the heat-conductive ideal gas without viscosity

Nonlinear stability of the boundary layer and rarefaction wave for the inflow problem governed by the heat-conductive ideal gas without viscosity

Theoretical Analysis of Supersonic Low Reynolds Number Flow Control Technique Based on Flow Mixing

The thickness of boundary layer increases rapidly in supersonic low Reynolds number flows. As a result, there will be shock wave-boundary layer interference. The disturbances of downstream can spread upstream through in boundary layer because the flow in boundary layer is subsonic, which may decrease flow field uniformity and stability. The pressure of this type of flow is always very low. For the sake of flow continuity, the increase of flow pressure is necessary. A kind of supersonic low Reynolds number flow control technique was proposed. The high pressure supersonic flow jet-ted from plates which located in the flow channel blew away boundary layer, cut off the disturbances spread from downstream to upstream, and eliminated the shock wave-boundary layer interference. The pressure of primary flow was also increased by mixing with the high pressure jetting flow. The calculation model of this flow control technique was established. A group of standard inputs were selected to calculate the parameters of flow and the preliminary law of flow was also obtained.

Control of transitional shock wave boundary layer interaction using structurally constrained surface morphing

The potential of surface morphing techniques, including passive shock control bumps (SCB) and active surface morphing, is explored to control transitional shock wave boundary layer interactions (SWBLI). In addition to reducing the size of the separation bubble, a key objective is to mitigate the low-frequency unsteadiness that can cause detrimental structural response. To this end, three-dimensional flow simulations are performed using direct numerical simulations (DNS) at Mach 2 and Reynolds number based on inflow boundary layer thickness Reδin=996. An incident oblique shock of angle σ=35deg and strength p2/p1=1.4 impinges on a laminar boundary layer that evolves from a Blasius profile. The resulting boundary layer separation leads to transition and a Görtler-like instability is observed; the nominally two-dimensional and steady flow becomes three-dimensional and unsteady. The goal of this work is to develop a surface modification method to mitigate separation and low-frequency unsteadiness, which can trigger structural response and flow distortion. An aero-structural solver framework is developed and employed to examine both passive SCB and active surface morphing. To avoid unrealistic structural deformation in the transient or final states, the structural integrity is concurrently monitored so that the intermediate morphing solutions are restricted to achievable elastic deformations. The results indicate that transitional SWBLI can be controlled in this manner, to essentially inhibit transition and thus eliminate the separation and unsteadiness associated with the Görtler-like vortices. The mechanism modulates sharp increases in surface pressure at separation and shock-impingement locations encountered in uncontrolled SWBLI and results a lower specific entropy rise.

Incident shock wave and supersonic turbulent boundarylayer interactions near an expansion corner

Direct numerical simulations of incident shock wave and supersonic turbulent boundary layer interactions near an expansion corner are performed at Mach number M∞ = 2.9 and Reynolds number Re∞ = 5581 to investigate the expansion effect on the characteristic features of this phenomenon. Four expansion angles, i.e. α = 00 (flat-plate), 20, 50 and 100 are considered. The nominal impingement point of the oblique shock wave with a flow deflection angle of 120 is fixed at the onset of the expansion corner, and flow conditions are kept the same for all cases. The numerical results are in good agreement with previous experimental and numerical data. Various flow phenomena, including the flow separation, the post-shock turbulent boundary layer and the flow unsteadiness in the interaction region, have been systematically studied. Analysis of the instantaneous and mean flow fields indicates that the main effect of the expansion corner is to significantly decrease the size and three-dimensionality of the separation bubble. A modified scaling analysis is proposed for the expansion effect on the interaction length scale, and a satisfactory result is obtained. Distributions of the mean velocity, the Reynolds shear stress and the turbulent kinetic energy show that the post-shock turbulent boundary layer in the downstream region experiences a faster recovery to the equilibrium state as the expansion angle is increased. The flow unsteadiness is studied using spectral analysis and dynamic mode decomposition, and dynamically relevant modes associated with flow structures originated from the incoming turbulent boundary layer are clearly identified. At large expansion angle (α=100), the unsteadiness of the separated shock is dominated by medium frequencies motions, and no low frequency unsteadiness is observed. The present study confirms that the driving mechanism of the low frequency unsteadiness is strongly related to the separated shock and the detached shear layer.

Parameterization of Wave Boundary Layer

It is known that drag coefficient varies in broad limits depending on wind velocity and wave age as well as on wave spectrum and some other parameters. All those effects produce large scatter of the drag coefficient, so, the data is plotted as a function of wind velocity forming a cloud of points with no distinct regularities. Such uncertainty can be overcome by the implementation of the WBL model instead of the calculations of drag with different formulas. The paper is devoted to the formulation of the Wave Boundary Layer (WBL) model for the parameterization of the ocean-atmosphere interactions in coupled ocean-atmosphere models and wave prediction models. The equations explicitly take into account the vertical flux of momentum generated by the wave-produced fluctuations of pressure, velocity and stresses (WPMF). Their surface values are calculated with the use of the spectral beta-functions whose expression was obtained by means of the 2-D simulation of the WBL. Hence, the model directly connects the properties of the WBL with an arbitrary wave spectrum. The spectral and direct wave modeling should also take into account the momentum flux to a subgrid part of the spectrum. The parameterization of this effect in the present paper is formulated in terms of wind and cut-off frequency of the spectrum.

On the use of thermally perfect gas model for heating prediction of laminar and turbulent SWBLI

The model of thermally perfect gas has been investigated in the heating prediction of laminar and turbulent shock wave boundary layer interaction (SWBLI) flows. The method of the polynomial curve fit coupling to thermodynamics equation is examined for higher temperature range. The characteristic weighted essentially non-oscillatory (WENO) scheme is utilized to capture the complex flow details and meantime stabilize the strong shocks in the SWBLI flows. Both the laminar and turbulent, steady and unsteady test cases are carried out to assess the performance of the designed method. Numerical results have substantiated the thermally perfect gas model as an effective model in the challenging SWBLI heating predictions in comparison with the existing model of calorically perfect gas.

Wave Boundary Layer Model based wind drag estimation for tropical storm surge modelling in the Bay of Bengal

Modelling of storm surge numerically has always been strenuous as it is associated with various influencing factors. The major governing component being the wind forcing or the wind stress - that signifies the computational accuracy of simulated surge and wave parameters. The present study is aimed at analysing the most reliable wind drag evaluation method for real-time predictions of storm surge. In an attempt to examine the same, two tropical cyclones (Thane and Vardah) eventuated in the Northern Indian Ocean were modelled. Four most widely used linear empirical and quadratic relations along with the enhanced Wave Boundary Layer Model (WBLM) were executed for the estimations of wind drag coefficient. The surge was subsequently simulated (using the coupled hydrodynamic circulation and wave model: ADCIRC and SWAN, respectively), individually for each of the above wind stress methods to obtain the corresponding water levels and wave features. The modelled values were further validated with the observed water level and wave data. It was quite intuitively observed that, WBLM outperformed (i.e. significantly correlated with the observed values) its linear counterparts since, the former pragmatically includes the effects of air-sea interface at high wind speeds in the model. However, due to insufficiency of the observed parameters, water levels obtained at locations of maximum wind speeds could not be satisfactorily validated. The WBLM-based computation of significant wave heights in deep as well as shallow water nevertheless enabled efficient and reasonably-reliable estimations of the peak incidents (along with relative wave directions and energy densities). Although this study essentially includes cyclonic wind velocities, it was identified that - with given wind field that includes the ‘pre-existing’ wind-wave climate in the relevant domain (i.e. before the cyclone genesis) the modelled estimates could be more precise.

Influences of wall vibration on shock train structures and performance of two-dimensional rectangular isolators in scramjet engine

Shock trains are formed by shock wave boundary layer interaction in isolator to match the pressure between inlet and combustor. However, the complexity of this problem has primarily restricted to rigid isolators without deformation. A little is known regarding the impact of wall vibration, due to structural elasticity, on the flow in isolator. The current study focuses on this problem by comparing changes of 13 parameters regarding separation zone, shock structure, flow asymmetry and isolator performance. Results examine the differences due to one-wall vibration, two-wall vibration and short panel vibration. Analyses indicate that one-wall vibration will cause the upstream movement and length decrease of separation zone, upstream movement and length growth of shock trains, increase of flow asymmetry with large transient side loads and decrease of performance. Moreover, compared to one-wall vibration, two walls vibrating in opposite direction leads to a larger influence on flow structures and isolator performance, while the effect of two-wall vibration in the same direction is relative minor on flow structures and significant on flow asymmetry as well as isolator performance. Furthermore, by placing vibrating panels at different locations, analyses identify that the influence of upstream and midstream vibrations is stronger than downstream vibration in wall vibration. The upstream vibration has a greater impact on flow structures while the midstream vibration has a greater effect on flow asymmetry and isolator performance. The current study emphasizes the importance of future investigations on aeroelastic problem in isolator and provides primary knowledge for flow control by considering flexible or vibrating panel.

Effect of Varying Ramp Angle and Leading-Edge Bluntness on the Behavior of Ramp Induced Shock Wave over Triple Ramped Cone Flare Configuration at Hypersonic Speed

Numerical simulation results are presented to show the effect of ramp angle variations and leading-edge bluntness on the flow around triple ramped cone flare in hypersonic flow. This study investigates the changes associated with shock wave boundary layer interaction due to ramp induced flow breakdown and the fluctuation in flow in the presence of blunted leading edge. This type of ramp junctions typically features in re-entry vehicles, engine intakes, system and sub-system junctions, control surfaces, etc. Ramp junctions usually are associated with strong separation bubble that has significant upstream influence impacting the effectiveness of aerodynamic surfaces, engine performance, thermal behavior and stability. Computation studies are carried out using finite volume-based RANS solver, accuracy of second order and considering compressible laminar flow characteristics, with solver settings provided similar to experimental conditions as per literature. Comprehensive double ramp studies with suggestions on reducing the separation bubble size are invariantly considered in literature, however there has been no study in understanding the inclusion of additional ramps in such flow scenarios, hence efforts are taken to understand the benefits and implications of including a third ramp along with varying bluntness on the bubble size and its upstream intensity.

Hypersonic Flows Over Multi-Ramp Configurations

The effects of attaching multiple ramps to the standard double ramp configuration along with variations in ramp angle, free-stream Mach number and surface temperature are discussed in this investigation. This study investigates the changes associated with shock wave boundary layer interaction (SWBLI) due to ramp induced flow breakdown and the flow field fluctuation with changes in flow characteristics and design. This type of ramp junctions typically features in re-entry vehicles, engine intakes, system and sub-system junctions, control surfaces, etc. Ramp junctions usually are associated with strong separation bubble that has significant upstream influence impacting the effectiveness of aerodynamic surfaces, engine performance, thermal behavior and stability. Computation studies are carried out using Second order accurate, finite volume RANS solver considering compressible laminar flow characteristics, with solver settings provided like experimental conditions as per literature. Comprehensive double ramp studies with suggestions on reducing the separation bubble size are invariantly considered in literature, however there has been no study in understanding the inclusion of additional ramps in such flow scenarios. At the end of this study it was evident that such complex junction needs detailed understanding on how they benefit or impact the overall design of the system. It also gave a very good insight on the nature of flow around such complex junctions and instills motivation for detailed experimental understanding.

Design and performance study of a parametric diverterless supersonic inlet

Diverterless supersonic inlet integration for a flight vehicle requires a three-dimensional compression surface (bump) design with an acceptable shock structure and boundary layer diversion; this results in a low drag induction system with acceptable propulsive efficiency. In this investigation, a computational fluid dynamics-based-generated bump is used to design an integrated diverterless supersonic inlet without any bleed mechanism on a forebody with a large wetted area. Numerical solution of the Navier–Stokes equations simulates the flow pattern of the configuration. The forebody design analysis includes simulating the effects of angle of attack and sideslip by dependent computational domains. Results demonstrate the ability of the bump surface to keep the shock structures in an operational mode even at high supersonic angles of attack. Analysis of shock structures and shock wave boundary layer interactions at supersonic maneuver conditions indicate that the aerodynamic efficiency of the diverterless supersonic inlet in conditions with a thick boundary layer and high angles of attack is sufficient to ensure operation throughout the supersonic flight envelope.

Transport due to Transient Progressive Waves

AbstractMaking use of a Lagrangian description, we interpret the kinematics and analyze the mean transport due to numerically generated transient progressive waves, including breaking waves. The waves are packets and are generated with a boundary-forced, air–water, two-phase Navier–Stokes solver. These transient waves produce transient transport, which can sometimes be larger than what would be estimated using estimates developed for translationally invariant progressive waves. We identify the critical assumption that makes our standard notion of the steady Stokes drift inapplicable to the data and explain how and in what sense the transport due to transient waves can be larger than the steady counterpart. A comprehensive analysis of the data in the Lagrangian framework leads us to conclude that much of the transport can be understood using an irrotational approximation of the velocity, even though the simulations use Navier–Stokes fluid simulations with moderately high Reynolds numbers. Armed with this understanding, it is possible to formulate a simple Lagrangian model that captures the mean transport and variance of transport for a large range of wave amplitudes. For large-amplitude waves, the parcel paths in the neighborhood of the free surface exhibit increased dispersion and lingering transport due to the generation of vorticity. We examined the wave-breaking case. For this case, it is possible to characterize the transport very well, away from the wave boundary layer, and approximately using a simple model that captures the unresolved breaking dynamics via a stochastic parameterization.

Shock wave-Boundary Layer Interactions in Wedge-induced Oblique Detonations

ABSTRACT The flow structure and stability of oblique detonation waves (ODWs) affected by shock wave-boundary layer interactions (SBLIs) are investigated based on Reynolds averaging method. ODWs with smooth and abrupt transitions are studied separately. The results show that there are no shock waves behind detonation wave surface for ODWs with smooth transitions, so the flow structures are only affected by the ramp-induced SBLI. Under the circumstances, the compression effects of the focused shock instead of the separation shock is the main cause of the initiation of ODW, which leads to an obvious increase in the initiation length. In ODWs with abrupt transitions, the primary transverse wave is formed and reflects on the wedge surface. Besides the ramp-induced SBLI, post-wave SBLI also occurs. The two kinds of SBLIs are influenced by the thickness of the inflow boundary layer and the activity of the inflow mixture. When the inflow boundary layer is thin and the activity is low, the separation zone is small and the distance between the ramp-induced separation zone and the post-wave separation zone is large, which makes the ramp-induced separation separated from the post-wave subsonic area. The post-wave SBLI makes the shock configuration at the end of the induction zone change from the λ-shaped to the Y-shaped, which weakens the stability of the ODW. When the inflow boundary layer is thick or the activity is high, the separation zone is large and the distance between the two separation areas is large, which makes the ramp-induced separation merge with the post-wave subsonic area and an extended separation is formed which covers the wedge surface. As the flowfield develops, the extended separation becomes larger and larger, leading to further increase of the initiation length. Finally the ODW propagates out of the calculation domain and fails to be stabilized on the wedge surface.

Probing Real Gas and Leading-Edge Bluntness Effects on Shock Wave Boundary-Layer Interaction at Hypersonic Speeds

AbstractThe present investigations are centered on understanding the discrepancies in shock wave boundary-layer interaction (SWBLI) for perfect and real gas laminar flows. In view of this, the in-h...