Shock Wave Turbulent Boundary Layer Research Articles

Reconstruction of velocity fields from wall pressure measurements in a shock wave/turbulent boundary layer interaction

We present here experimental results in a shock wave/turbulent boundary layer interaction at Mach number of 2.3 impinged by an oblique shock wave, with a deflection angle of 9.5°, as installed in the supersonic wind tunnel of the IUSTI laboratory, France. For such a shock intensity, strong unsteadiness are developing inside the separated zone involving very low frequencies associated with reflected shock motions.The present work consists in simultaneous PIV velocity fields and unsteady wall pressure measurements. The wall pressure and PIV measurements were used to characterize the pressure distribution at the wall in an axial direction, and the flow field associated. These results give access for the first time to the spatial-time correlation between wall pressure and velocity in a shock wave turbulent boundary layer interaction and show the feasibility of such coupling techniques in compressible flows. Linear Stochastic Estimation (LSE) coupled with Proper Orthogonal Decomposition (POD) has been applied to these measurements, and first results are presented here, showing the ability of these techniques to reproduce both the unsteady breathing of the recirculating bubble at low frequency and the Kelvin–Helmholtz instabilities developing at moderate frequency.

Time-Frequency Analysis of Rocket Nozzle Wall Pressures during Start-up Transients

Surveys of the fluctuating wall pressure were conducted on a sub-scale, thrust-optimized parabolic nozzle in order to develop a physical intuition for its Fourier-azimuthal mode behavior during fixed and transient start-up conditions. These unsteady signatures are driven by shock wave turbulent boundary layer interactions which depend on the nozzle pressure ratio and nozzle geometry. The focus however, is on the degree of similarity between the spectral footprints of these modes obtained from transient start-ups as opposed to a sequence of fixed nozzle pressure ratio conditions. For the latter, statistically converged spectra are computed using conventional Fourier analyses techniques, whereas the former are investigated by way of time-frequency analysis. The findings suggest that at low nozzle pressure ratios –where the flow resides in a Free Shock Separation state– strong spectral similarities occur between fixed and transient conditions. Conversely, at higher nozzle pressure ratios –where the flow resides in Restricted Shock Separation– stark differences are observed between the fixed and transient conditions and depends greatly on the ramping rate of the transient period. And so, it appears that an understanding of the dynamics during transient start-up conditions cannot be furnished by a way of fixed flow analysis.

Correlation study in shock wave–turbulent boundary layer interaction

Shock wave–turbulent boundary layer interaction is a critical problem in aircraft design. Therefore, a thorough understanding of the processes occurring in such flows is necessary. The most important task is to study the unsteady phenomena, in particular, the low-frequency ones, for this interaction. An experimental study of separated flow has been performed in the zone of interaction of the incident oblique shock wave with a turbulent boundary layer at Mach 2. Two-point correlation data in the separation zone and the upstream flow were obtained and showed that low-frequency oscillations of the reflected shock waves are related to pulsations in the inflow turbulent boundary layer.

Effect of air jet vortex generators on a shock wave boundary layer interaction

The effect of upstream injection by means of continuous air jet vortex generators (AJVGs) on a shock wave turbulent boundary layer interaction is experimentally investigated. The baseline interaction is of the impinging type, with a flow deflection angle of 9.5° and a Mach number M e = 2.3. Considered are the effects of the AJVGs on the upstream boundary layer flow topology and on the spatial and dynamical characteristics of the interaction. To this aim, Stereoscopic Particle Image Velocimetry has been employed, in addition to hot-wire anemometry (HWA) for the investigation of the unsteady characteristics of the reflected shock. The AJVGs cause a reduction of the separation bubble length and height. In addition, the energetic frequency range of the reflected shock is increased by approximately 50%, which is in qualitative agreement with the smaller separation bubble size.

Application of a dual-plane particle image velocimetry (dual-PIV) technique for the unsteadiness characterization of a shock wave turbulent boundary layer interaction

The unsteady organization and temporal dynamics of the interaction between a planar shock wave impinging on a turbulent boundary layer at a free-stream Mach number of Me = 1.69 is investigated experimentally by means of dual-plane particle image velocimetry (dual-PIV). Two independent PIV systems were combined in a two-component mode to obtain instantaneous velocity fields separated by a prescribed small time delay. This enables us to obtain, in addition to mean and statistical flow properties, also instantaneously time-resolved data to characterize the temporal dynamics of the flow phenomenon in terms of time scales, temporal correlations and convective velocities. The characteristic time scales for the incoming boundary layer, the separation region and the reflected shock are determined by means of the temporal auto-correlation coefficient in the complete flow field for a range of time delays from 5 µs to 2000 µs. These auto-correlation fields are used to quantify the time scales in selected regions of the flow, with special interest for the vortex shedding and the low frequency reflected shock dynamics. This permits resolving the dominant time scales within the boundary layer and the interaction region.

Effect of dimples on glancing shock wave turbulent boundary layer interactions

An experimental study has been conducted to examine the control effectiveness of dimples on the glancing shock wave turbulent boundary layer interaction produced by a series of hemi-cylindrically blunted fins at Mach numbers 0.8 and 1.4, and at angles of sweep 0°, 15°, 30° and 45°. Schlieren photography, oil flow, pressure sensitive paints, and pressure tappings were employed to examine the characteristics of the induced flow field. The passive control technique used a series of 2 mm diameter, 1 mm deep indents drilled across the hemi-cylindrical leading edge at angles 0°, 45° and 90°. The effects of dimples were highly dependent on their orientation relative to the leading edge apex, and the local boundary layer properties.

A General Strategy to Extend Turbulence Models to Rough Surfaces: Application to Smith’s k-L Model

Abstract A general procedure to extend turbulence models to account for wall roughness, in the framework of the equivalent sand grain approach, is proposed. It is based on the prescription of the turbulent quantities at the wall to reproduce the shift of the logarithmic profile and hence provide the right increase in wall friction. This approach was previously applied to Spalart and Allmaras one equation (1992, “A One-Equation Turbulence Model for Aerodynamic. Flows,” 30th Aerospace Sciences Meeting and Exhibit, Reno, NV, AIAA paper No. 92-0439;1994, ibid, Rech. Aerosp. 1, pp. 5–21). Here, the strategy is detailed and applied to Smith’s two-equation k-L model (1995, “Prediction of Hypersonic Shock Wave Turbulent Boundary Layer Interactions With The k-l Two Equaton Turbulence Model,” 33rd Aerospace Sciences Meeting and Exhibit, Reno, NV, Paper No. 95-0232). The final model form is given. The so-modified Spalart and Allmaras and Smith models were tested on a large variety of test cases, covering a wide range of roughness and boundary layer Reynolds numbers and compared with other models. These tests confirm the validity of the approach to extend any turbulence model to account for wall roughness. They also point out the deficiency of some models to cope with small roughness levels as well as the drawbacks of present correlations to estimate the equivalent sand grain roughness.

Large Eddy Simulation of Transonic Flow with Shock Wave/Turbulent Boundary Layer Interaction

Large eddy simulation was made of transonic flow over a two-dimensional bump where shock wave turbulent boundary layer interaction takes place. Grid refinement and the effect of the domain width were investigated. Special care was taken to ensure physically correct inlet boundary conditions. The shock wave turbulent boundary layer interaction induces strong separation of the boundary layer and events such as bursting events in the incoming boundary layer, and creation of large flow scale structures behind the shock are detected. However, the shock features no large scale movement even though, according to several sources, the prerequisites for such movement are fulfilled.

Incipient separation in shock wave/boundary layer interactions as induced by sharp fin

The incipient separation induced by the shock wave turbulent boundary layer interaction at the sharp fin is the subject of present study. Existing theories for the prediction of incipient separation, such as those put forward by McCabe (1966) and Dou and Deng (1992), can have thus far only predicting the direction of surface streamline and tend to over-predict the incipient separation condition based on the Stanbrook's criterion. In this paper, the incipient separation is firstly predicted with Dou and Deng (1992)'s theory and then compared with Lu and Settles (1990)' experimental data. The physical mechanism of the incipient separation as induced by the shock wave/turbulent boundary layer interactions at sharp fin is explained via the surface flow pattern analysis. Furthermore, the reason for the observed discrepancy between the predicted and experimental incipient separation conditions is clarified. It is found that when the wall limiting streamlines behind the shock wave aligning with one ray from the virtual origin as the strength of shock wave increases, the incipient separation line is formed at which the wall limiting streamline becomes perpendicular to the local pressure gradient. The formation of this incipient separation line is the beginning of the separation process. The effects of Reynolds number and the Mach number on incipient separation are also discussed. Finally, a correlation for the correction of the incipient separation angle as predicted by the theory is also given.

Numerical Investigation of Control of Three-Dimensional Shock Wave/Turbulent Boundary Layer Interactions by Vortex Generator.

Suppression of shock wave/turbulent boundary layer(SW/TBL)interaction occurred inside a supersonic air-intake is important for ensuring and maintaining the aerodynamic performance. It is well known that boundary layer bleed is effective for the SW/TBL interactions and has been employed practically. However, in respect to bleed, there are some shortomings, for example, loss of mass flow rate. In this study, we focused on vortex generators (VGs) that are used as a control device for separation. We calculated a three dimensional SW/TBL interaction field by VGs and investigated the effect of VGs in such a flow field numerically. We researched several configurations of VGs, and found that the largest VGs, which is almost as high as boundary layer thickness, can suppress the SW/TBL interaction in the light of the flow distortion.

Heating characteristics of blunt swept fin-induced shock wave turbulent boundary layer interaction

An experimental study was conducted on shock wave turbulent boundary layer interactions caused by a blunt swept fin-plate configuration at Mach numbers of 5.0, 7.8, 9.9 for a Reynolds number range of (1.0.similar to 4.7) x 10(7)/m. Detailed heat transfer and pressure distributions were measured at fin deflection angles of up to 30 degrees for a sweepback angle of 67.6 degrees. Surface oil flow patterns and liquid crystal thermograms as well as schlieren pictures of fin shock shape were taken. The study shows that the flow was separated at deflection of 10 degrees and secondary separation were detected at deflection of theta greater than or equal to 20 degrees. The heat transfer and pressure distributions on flat plate showed an extensive plateau region followed by a distinct dip and local peak close to the fin foot. Measurements of the plateau pressure and heat transfer were in good agreement with existing prediction methods, but pressure and heating peak measurements at M greater than or equal to 6 were significantly lower than predicted by the simple prediction techniques at lower Mach numbers.

Flowfield-dependent mixed explicit-implicit (FDMEI) methods for high and low speed and compressible and incompressible flows

Despite significant achievements in computational fluid dynamics, there still remain many fluid flow phenomena not well understood. For example, the prediction of temperature distributions is inaccurate when temperature gradients are high, particularly in shock wave turbulent boundary layer interactions close to the wall. Complexities of fluid flow phenomena include transition to turbulence, relaminarization, separated flows, transition between viscous and inviscid, incompressible and compressible flows, among others, in all speed regimes. The purpose of this paper is to introduce a new approach, called the Flowfield-Dependent Mixed Explicit-Implicit (FDMEI) method, in an attempt to resolve these difficult issues in CFD. In this process, a total of six implicitness parameters characteristic of the current flowfield are introduced. They are calculated from the current flowfield or changes of Mach numbers, Reynolds numbers, Peclet numbers, and Damköhler numbers (if reacting) at each nodal point and time step. This implies that every nodal point or element is provided with different or unique numerical scheme according to their current flowfield situations, whether compressible, incompressible, viscous, inviscid, laminar, turbulent, reacting, or nonreacting. In this procedure, discontinuities or fluctuations of all variables between adjacent nodal points are determined accurately. If these implicitness parameters are fixed to certain numbers instead of being calculated from the flowfield information, then practically all currently available schemes of finite differences or finite elements arise as special cases. Some benchmark problems to be presented in this paper will show the validity, accuracy, and efficiency of the proposed methodology.

Three-dimensional mixed explicit-implicit generalized Galerkin spectral element methods for high-speed turbulent compressible flows

In high speed flows the interactions of shock waves with turbulent boundary layers are important design considerations because of the complex flowfields resulting in increased adverse pressure gradients, skin friction and temperatures. Unsteadiness and three-dimensional flowfield structure are also characteristic of shock wave turbulent boundary layer interactions. Such physical phenomena require sophisticated numerical schemes in the solution of governing equations. The purpose of this paper, therefore, is to introduce an accurate and efficient approach—the Mixed Explicit-Implicit Generalized Galerkin Spectral Element Method (MEI-GG-SEM) with Legendre polynomial spectral elements in which flowfield dependent implicitness parameters provide automatically adequate computational requirements for compressible and incompressible flows or high speed and low speed flows. This is in contrast to the traditional approach in which all-speed-regime analysis requires a separate hyperbolic-elliptic pressure equation for pressure correction if the flow becomes incompressible. In the MEI-GG-SEM scheme, mesh refinements are carried out adaptively until shock waves are resolved, followed then by the adaptive increase of Legendre polynomial degrees until turbulence microscales are resolved, in which the traditional turbulence modeling is no longer required, aimed toward direct numerical simulation. In order to demonstrate the validity of the theory and numerical procedure, two-dimensional flat plate and compression corner high speed flows are investigated, followed by a three-dimensional sharp leading edged fin for swept shock wave turbulent boundary layer interactions. Comparisons of the present study with experimental measurements and other numerical studies show favorable agreement.

NUMERICAL SIMULATION OF HEAT TRANSFER IN CHEMICALLY REACTING SHOCK WAVE-TURBULENT BOUNDARY LAYER INTERACTIONS

Abstract The flow field of a transverse jet in a supersonic airstream subjected to shock wave-turbulent boundary layer interactions is simulated numerically by adaptive mixed explicit-implicit generalized-Galerkin finite element methods. In this scheme, convection and diffusion implicitness parameters are introduced to resolve shock wave discontinuities and widely disparate time and length scales of turbulence and finite rate chemistry. These parameters are flow field dependent, calculated from local Mach, Reynolds, and Damkohler numbers for each element. Effects of turbulence are taken into account with a two-equation (k-ε) model with a compressibility correction. Various cases of mixing, slot widths, and total pressure ratios with and without chemical reactions are examined. Favorable comparisons with experimental measurements are demonstrated.

Measurements of the mean heat transfer in a shock wave-turbulent boundary layer interaction

Using thin-film resistance technology characterized by a fast response time and a high sensitivity to heat transfer, a method is developed to measure mean Stanton number in a supersonic blowdown facility under near-adiabatic conditions. The successful application of thin-film gauges required corrections for the changes in total temperature with time and for departures from one-dimensional heat conduction in the gauge itself. This method was then demonstrated in a shock wave-turbulent boundary layer interaction generated by a 16° compression corner in a Mach 2.84 flow. The results show a strong increase in Reynolds analogy factor across the shock, and it appears to level off at a value of 2.15, which is approximately 1.75 times the usually quoted value for an unperturbed boundary layer flow.

Flow Separation in Shock Wave Boundary Layer Interactions

This work presents an assessment of the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists of selected configurations where the only characteristic length in the iteration is the incoming boundary layer thickness. It consists of two-dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the empirical correlations for incipient separation are reviewed. The effects of Mach number, Reynolds number, surface cooling, and the methods of detecting separation are discussed.

NUMERICAL INVESTIGATION OF SEPARATED TRANSONIC TURBULENT FLOWS WITH A MULTIPLE-TIME-SCALE TURBULENCE MODEL

Numerical investigation of transonic flows separated by streamline curvature and shock wave-boundary layer interaction is presented. The free-stream Much numbers considered are 0.4, 0.5, 0.6, 0.7, 0.8, 0.825, 0.85, 0.875, 0.90, and 0.925. In the numerical method, the conservation of mass equation is replaced by a pressure correction equation for compressible flows and thus the equation is solved for incremental pressure rather than density. The turbulence is described by a multiple-time-scale turbulence model supplemented with a near-wall turbulence model. The present numerical results show that there exists a reversed flow region at all free-stream Mach numbers considered whereas various k-e turbulence models fail to predict such a reversed flow region at low free-Stream Mach numbers. The numerical results also show that the size of the reversed flow region grows extensively due to the shock wave-turbulent boundary layer interaction as the free-stream Mach number is increased. These numerical results sho...

Numerical simulation of a three-dimensional shock wave-turbulent boundary layer interaction generated by a sharp fin at Mach 4

The three-dimensional inviscid-viscous interaction generated by the intersection of an oblique shock with a turbulent boundary layer is examined theoretically. An incoming Mach 4 equilibrium turbulent boundary layer at Reynolds number Re δx = 2 × 10 5 interacts with an oblique shock formed by a 16° sharp fin mounted normal to the flat plate. The theoretical model is the three-dimensional compressible Reynolds-averaged Navier-Stokes equations with turbulence incorporated through the Baldwin-Lomax algebraic eddy viscosity model. The governing equations are solved using the hybrid explicit-implicit algorithm of the author. Computed results for the surface skin friction, surface pressure and surface streamline angles are compared with experimental measurements and previous numerical results. The present results display good agreement with experiment, except in specific isolated regions, and show the closest agreement with experiment of all computations.

An experimental investigation of the Reynolds number effect on a normal shock wave-turbulent boundary layer interaction on a curved wall

The presented experiment concerned an interaction between a normal shock wave, terminating a local supersonic area in a curved duct, and a turbulent boundary layer developed along the convex wall. This paper deals with the Reynolds number effect upon the interaction structure. The measurements included flow parameters distribution determination, boundary layer development through the interaction area and the shock wave topography visualization. In order to gain more information about separation the wall oil tracing has been applied. The comparison of our results with other published data is presented.

Numerical investigation of the impingement of an oblique shock on a flat plate

The unsteady, compressible Navier-Stokes equations in Reynolds-averaged variables are solved for a shock wave turbulent boundary layer interaction. For the free stream Mach number 3.7 and Reynolds number 8.202×106 computation is performed using MacCormack's explicit-implicit finite difference scheme with 42×42 grid points. It is found that the peak heating amplification correlation agrees well with computational results.