Shock Wave-turbulent Boundary Layer Interaction Research Articles

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.

An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade

Experiments have been performed in a supersonic cascade facility to elucidate the fluid dynamic phenomena and loss mechanism of a strong shock-wave turbulent boundary layer interaction in a compressor cascade. The cascade geometry is typical for a transonic fan tip section that operates with a relative inlet Mach number of 1.5, a flow turning of about 3 deg, and a static pressure ratio of 2.15. The strong oblique and partly normal blade passage shock-wave with a preshock Mach number level of 1.42 to 1.52 induces a turbulent boundary layer separation on the blade suction surface. The free-stream Reynolds number based on chord length was about 2.7 × 106. Cascade overall performance, blade surface pressure distributions, Schlieren photographs, and surface visualizations are presented. Detailed Mach number and flow direction profiles of the interaction region (lambda shock) and the corresponding boundary layer have been determined using a Laser-2-Focus anemometer. The obtained results indicated that the axial blade passage stream sheet contraction (axial velocity density ratio) has a significant influence on the mechanism of strong interaction and the resulting total pressure losses.

Inception length to a fully developed, fin-generated, shock-wave, boundary-layer interaction

An experimental study of fin-generated shock wave turbulent boundary-layer interactions confirmed previous observations that, sufficiently far from the fin apex, such interactions become conical. The inception length to conical symmetry was found to increase weakly with Mach number for Mach numbers from 2.5 to 4 and fin angles from 4 to 22 deg. For the range of interactions examined, the inception length was found to depend primarily upon the inviscid shock angle, this angle ranging from 21 to 40 deg. The behavior of the inception length with shock angle can be broadly divided into two categories. For 'weak' interactions with shock angles less than about 35 deg, the inception length decreased as the shock angle increased. For 'strong' interactions with shock angles greater than about 35 deg, the inception region was small and was approximately constant at three boundary-layer thicknesses in length. In the latter, strong interaction case, the inception length was an order of magnitude smaller than that found in the weakest interactions examined, to the extent that strong interactions were practically fully-developed from the apex.

Flow Unsteadiness by Weak Normal Shock Wave/Turbulent Boundary Layer Interaction in Internal Flow

The interaction of a weak normal shock wave with a turbulent boundary layer developing along the wall surface of a supersonic diffuser was investigated experimentally, where the flow Mach number was 1.14∼1.53 and the Reynolds number based upon boundary layer thickness was 1.35∼2×104. Detailed streamwise surface pressures were measured, and a number of data sets obtained by wall pressure transducers were statistically analyzed. The power spectral density function, the intermittency and the higher-order moments of the surface pressure fluctuations were taken. The objective of the present experiment was to investigate the behaviors of the oscillations caused by the normal shock wave/turbulent boundary layer interaction (STBI). The results show that the intermittent feature of the unsteady shock motion is similar to those in external flows and that the maximum standard deviations correspond to points where the intermittency factor approaches 0.5 and are significantly influenced by flow separation. Also the present study indicates that far downstream from the STBI region the amplitude of pressure fluctuations is nearly independent of the flow Mach number.

Calculation of a three-dimensional shock wave-turbulent boundary-layer interaction

We deal with the numerical simulation of a shock wave-turbulent boundary-layer interaction in a three-dimensional channnel by solution of the Reynolds average Navier-Stokes equations with a mixing-lengh turbulence model.The numerical method is characterized by an explicit centered finite-difference scheme associated with a multigrid convergence acceleration.

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.

Mach number effects on conical surface features of swept shock-wave/boundary-layer interactions

A joint experimental and computational study is made of the shock-wave turbulent boundary-layer interaction generated by sharp fins, with emphasis on Mach-number effects. The Mach number range is from 2 to 4 and the unit Reynolds number is from 50 to 80 million per meter. Fin angles are varied from 4 to 22 deg. Surface-flow patterns are obtained using a color surface-flow-visualization technique. The results show that the upstream-influence response in the conical far-field region is a function of the freestream Mach number and the shock strength. A new interpretation of the behavior of the upstream influence with changes of the inviscid shock angle is given. Agreement between the experimental and the computed upstream-influence lines becomes poorer for stronger interactions, with the computations underpredicting the upstream-influence line.

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.

Upstream influence and separation scales in fin-induced shock turbulent boundary-layer interaction

An experimental study has been made of the effects of leading-edge geometry on the upstream influence and separated flow length scales in fin-induced shock wave turbulent boundary-layer interaction. The fins used had wedge-shaped, hemicylindrical and elliptic leading edges and were tested under adiabatic conditions at Mach 4.9. The primary objective was to determine if the length scales generated by different leading edges could be understood within a common framework. To a first approximation the normalized centerline length scales correlate with fin leading-edge drag coefficient. Attempts to correlate additional data at Mach 1.98 and 2.48 from other studies resulted in a reasonable correlation for upstream influence, but that for the separation length showed a clear influence of Mach number. The fact that the upstream influence can be reasonably well correlated in this form lends further support to the view that large-scale three-dimensional vortical separated flows of this type may be largely inviscid dominated.

Interpretation of separation lines from surface tracers in a shock-induced turbulent flow

IN experimental studies of shock wave turbulent boundarylayer interactions, surface-tracer methods, which have essentially zero frequency response, are widely used to determine the locations of separation lines. However, recent research has shown that in many cases the separation shockwave structure undergoes large-scale motion, raising questions about the physical meaning of separation lines deduced from such techniques. In the present study at Mach 5, the separation induced by an uriswept semi-infinite circular cylinder has been studied using platinum thin films. Conditional sampling analysis of the signals suggests that this line is the downstream boundary of a region of intermittent separation. Contents In experimental studies of shock wave turbulent boundarylayer interactions, surface tracer techniques are widely used to locate separation lines.1 Such methods are easy to use and generate sharply defined, repeatable separation lines (S). Quantitative measurements of flow length scales and angles are made using these methods and are used for evaluating the predictive capabilities of numerical methods and turbulence models. It is, therefore, important to understand what these lines physically represent. In many of these flows, dynamic wall-pressure measurements show that the separation shock is unsteady and moves from the upstream boundary of the flowfield, where the wall

Oblique shock wave—turbulent boundary layer interaction with suction

AbstractThe interaction of an incident oblique shock wave with the turbulent boundary layer with suction is investigated. Wall streamline and schlieren photograph visualisations as well as the pressure measurements on the wall are made. Boundary layer surveys with pressure probes and hot-wire probes are also done. The suction downstream of the interaction almost nullifies the shock reflection while the higher suction up and downstream of the interaction results in a reflection similar to that in inviscid flow.

Numerical Computations of Turbulence Amplification in Shock-Wave Interactions

Numerical computations are presented which illustrate and test various effects pertinent to the amplification and generation of turbulence in shock wave-turbulent boundary layer interactions. Several fundamental physical mechanisms are identified. Idealizations of these processes are examined by nonlinear numerical calculations. The results enable some limits to be placed on the range of validity of existing linear theories. Additional results are given which are of a fundamentally nonlinear nature.

Unsteadiness of the Separation Shock Wave Structure in a Supersonic Compression Ramp Flowfield

Wall pressure fluctuations have been measured in a two-dimensional separated compression ramp-induced shock wave turbulent boundary-layer interaction. The tests were made at a nominal freestream Mach number of 3 and at Reynolds numbers based on boundary-layer thickness of 7.8 X 10 and 1.4x 10. The wall temperature condition was approximately adiabatic. Large-amplitude pressure fluctuations exist throughout the interaction, particularly near separation and reattachment. In the upstream region of the flowfield, the unsteadiness of the separation shock wave structure generates an intermittent wall pressure signal. Mean wall pressures in this region result from the superposition of the relatively low-frequency, large-amplitude, shock wave-induced fluctuations on the pressure signal of the undisturbed boundary layer. This behavior is qualitatively similar to that observed in three-dimensional blunt fin-induced flows. In these two flowfields, the length scale of the shock motion is a significant fraction of the distance from the interaction start to separation.

Upstream influence in sharp fin-induced shock wave turbulent boundary-layer interaction

Upstream influence in sharp fin-induced shock wave turbulent boundary-layer interaction

Integral method improvement for computation of transonic shock wave-turbulent boundary-layer interactions

Integral method improvement for computation of transonic shock wave-turbulent boundary-layer interactions

Reynolds Number Effects on Shock-Wave Turbulent Boundary-Layer Interactions

An experiment is described that tests and guides computations of a shock-wave turbulent boundary-layer interaction flow over a 20° compression corner at Mach 2.85. Numerical solutions of the time-averaged NavierStokes equations for the entire flow field, employing various turbulence models, are compared with the data. Each model is evaluated critically by comparisons with the details of the experimental data. Experimental results for the extent of upstream pressure influence and separation location are compared with numerical predictions for a wide range of Reynolds numbers and shock-wave strengths.

Supersonic Shock Wave Turbulent Boundary-Layer Interactions

Results are presented of an experimental investigation of shock wave turbulent boundary-layer interactions in supersonic flow. The experiments were conducted at a freestream Mach number of 2.96 and over a boundarylayer Reynolds number range of 10 5 to 10 6. Surface static pressure measurements, oil flow photographs, and interferograms were obtained to define the length of separation and the incipient separation angles for 1) twodimensional compression corner and 2) planar shock wave interactions with a turbulent boundary layer. The tests were conducted in a high unit Reynolds number freestream on a long flat plate with a turbulent boundarylayer thickness in the region of from 0.12 to 0.18 in. Direct comparisons were made between the compression corner and incident shock wave interactions to determine the effects of configuration on turbulent boundary-layer separation. For both configurations the length of the separated region was found to decrease and the incipient separation angle to increase with increasing Reynolds number. For constant Reynolds number, the overall pressure rise for incipient separation was approximately the same for the compression corner and the incident shock wave interaction. Turbulent boundary-layer separation was found to be of the free interaction type whereby the separation angle and pressure distribution through separation were independent of Reynolds number, overall pressure rise, and configuration. fo Nomenclature skin friction coefficient at beginning of freestream Mach number pressurepressure freestream pressure Reynolds number based on freestream condition and boundary-layer thickness at beginning of . span °f shock generator stagnation temperature wall temperature axial distance from flat plate leading edge axial location of center of axial distance from flat plate/ramp hinge line to shock generator leading edge axial location of separation point axial offset distance vertical distance between flat plate and shock generator leading edge effective incipient separation corner angle compression ramp angle shock generator angle boundary-layer thickness at beginning of

Conical shock-wave turbulent boundary-layer interaction including suction effects

Conical shock wave-turbulent boundary layer interaction data obtained for adiabatic wall conditions at supersonic free stream Mach numbers, including suction effects