Some new developments of explicit algebraic Reynolds stress turbulence models (EARSM) are presented. The new developments include a new near-wall treatment ensuring realizability for the individual stress components, a formulation for compressible flows, and a suggestion for a possible approximation of diffusion terms in the anisotropy transport equation. Recent developments in this area are assessed and collected into a model for both incompressible and compressible three-dimensional wall-bounded turbulent flows. This model represents a solution of the implicit ARSM equations, where the production to dissipation ratio is obtained as a solution to a nonlinear algebraic relation. Three-dimensionality is fully accounted for in the mean flow description of the stress anisotropy. The resulting EARSM has been found to be well suited to integration to the wall and all individual Reynolds stresses can be well predicted by introducing wall damping functions derived from the van Driest damping function. The platform for the model consists of the transport equations for the kinetic energy and an auxiliary quantity. The proposed model can be used with any such platform, and examples are shown for two different choices of the auxiliary quantity.

An explicit algebraic Reynolds stress model for incompressible and compressible turbulent flows

Review and assessment of turbulence models for hypersonic flows

Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

Review and synthesis of roughness-dominated transition correlations for reentry applications

Modern nonintrusive techniques are used to make skin-friction and heat transfer measurements in two shockwave/turbulent boundary-layer interactions (SWTBLIs). The two-dimensional SWTBLI is generated by impingement of an incident oblique shock wave on a flat-plate boundary layer. The three-dimensional SWTBLI results from the interaction of the swept shock generated by a fin with a flat-plate boundary layer. The measurements are made using the global interferometry skin-friction technique for the skin friction and the quantitative infrared thermography technique for the heat transfer rate. The results show that, for the two- and three-dimensional interactions, there is a clear difference in the behavior of skin friction and heat transfer as the strength of the shock is changed. This observation suggests that the analogy between momentum and heat transfer, which is the basis of many simplified physical models, is not valid in SWTBLIs. These new data supplement the previous measurements that include boundary-layer properties, surface pressure distributions, and patterns of the limiting streamlines. Taken together, these data complete a data set that is suited for computational fluid dynamics validation.

Skin Friction and Heat Flux Measurements in Shock Boundary Layer Interaction Flows

Aspects of Hypersonic Boundary Layer Transition Control

Documentation of two-dimensional impinging shock/turbulent boundary layer interaction flow

We present boundary-layer transition data on swept cylinders in hypersonic flow that were obtained in the hypersonic Ludwieg-tube wind tunnel of DLR. Experiments were conducted at M∞ = 5.0 and 6.9 for three swept cylinders with and without end plates and sweep angles of 30, 45, and 60 deg, respectively. To determine the state of the boundary layers, the liquid-crystal technique was applied. In addition, the response of attachment-line transition to surface roughness heights due to trip wires for sweep angles of 45 and 60 deg at M∞ = 5.0 was investigated. Freestream Reynolds numbers based on cylinder diameter for attachment-line transition were about 0.2-0.3 x 10 6 for the end plate disturbances and 0.9-1.2 x 10 6 for no contaminations. Trends of attachment-line transition Reynolds numbers with surface roughness heights are roughly similar to subsonic ones with respect to critical roughness heights. The behavior of the attachment-line transition around the critical roughness heights, however, depended strongly on spanwise Mach numbers. With flow visualizations, off-attachment-line transition could be observed, as well as fine streak pattern in laminar flow region upstream of the transition front, confirming streamwise vortices induced by crossflow instability at hypersonic speeds.

Boundary-layer transition on swept cylinders at hypersonic speeds

We present boundary-layer transition data on swept cylinders in hypersonic flow that were obtained in the hypersonic Ludwieg-tube wind tunnel of DLR Experiments were conducted at M∞ = 50 and 69 for three swept cylinders with and without end plates and sweep angles of 30, 45, and 60 deg, respectively To determine the state of the boundary layers, the liquid-crystal technique was applied In addition, the response of attachment-line transition to surface roughness heights due to trip wires for sweep angles of 45 and 60 deg at M∞ = 50 was investigated Freestream Reynolds numbers based on cylinder diameter for attachment-line transition were about 02-03 x 10 6 for the end plate disturbances and 09-12 x 10 6 for no contaminations Trends of attachment-line transition Reynolds numbers with surface roughness heights are roughly similar to subsonic ones with respect to critical roughness heights The behavior of the attachment-line transition around the critical roughness heights, however, depended strongly on spanwise Mach numbers With flow visualizations, off-attachment-line transition could be observed, as well as fine streak pattern in laminar flow region upstream of the transition front, confirming streamwise vortices induced by crossflow instability at hypersonic speeds

Boundary layer transition on swept cylinders at hypersonic speeds

On montre que des corrections globales ne peuvent rendre compte des effets d'interference des parois laterales en soufflerie, ces effets etant dus dans une large mesure au caractere tridimensionnel de l'ecoulement provenant de l'interaction entre les couches limites sur les parois laterales et le champ de pression produit par le profil

Egon Stanewsky

Papers

Documentation of two-dimensional impinging shock/turbulent boundary layer interaction flow

Boundary-layer transition on swept cylinders at hypersonic speeds

Boundary layer transition on swept cylinders at hypersonic speeds

Sidewall Effects on Airfoil Tests.