Shock-induced Turbulent Boundary Layer Separation Research Articles

Application of Hybrid RANS/LES Methods for the Simulation of Shock-Induced Turbulent Boundary Layer Separation

Application of Hybrid RANS/LES Methods for the Simulation of Shock-Induced Turbulent Boundary Layer Separation

An Experimental Study on the Influence of Vortex Generators on the Shock-Induced Boundary Layer Separation at M=1.4

The results of an investigation into the effects that sub-boundary layer vortex generators (SBVGs) have on reducing normal shock-induced turbulent boundary layer separation are presented. The freestream Mach number and Reynolds number were M=1.45 and R=15.9×106/m, respectively. Total pressure profiles, static pressure distributions, surface total pressure (Preston pressure) distributions, oil flow visualization, and Schlieren photographs were used in the result analysis. The effects of SBVG height and the location upstream of the shock were investigated. A novel tetrahedron shape SBVG with different lengths (30 mm and 60 mm) was used for these experiments. The effect of streamwise location of the longer SBVG on the interaction was also investigated. The location of the shock wave was controlled by an adjustable choke mechanism located downstream of the working section. The results show that an increase in the distance for the longer SBVG from 17.4δR to 25.5δR did not remove the separation entirely, but the shorter SBVG provided higher total pressure distribution within the boundary layer in the recovery region. This also provided a healthier boundary layer profile downstream of the interaction region with lower displacement thickness and shape factor.

Scale analysis and experimental observations of shock-induced turbulent boundary layer separation in nozzles

Time average shock-induced boundary layer separation is investigated using scale analyses, analytical modeling, and experiments. While the study focuses on turbulent boundary layer separation in overexpanded rocket nozzles, many of the analyses presented apply to the broad family of free interaction, shock-separated flows in which the structure of the boundary layer–shock interaction zone is self-similar and independent of the shock generator. The scale analyses lead to two approximate expressions for the wall pressure ratio at separation; over a range of separation Mach numbers, both models provide reasonable predictions of observed separation pressure ratios. The second model, representing a refinement of the first, appears to provide a fairly general description of free interaction separation: the model approximately captures separation pressure ratios observed in supersonic flow over backward facing steps and in the case of overexpanded nozzle flow, provides predictions that are consistent with the free interaction model. Experiments are carried out in a sub-scale nozzle under overexpanded, cold-flow conditions. The principal observations are as follows: (i) For the range of separation Mach numbers investigated ( 5.0 ⩽ M i ⩽ 5.4 ), nominal separation line locations can be predicted with reasonable accuracy using the classical generalized quasi-one-dimensional compressible flow model and an appropriate separation criterion. (ii) Over the same range of overexpanded flow conditions, the time-average pressure rise over the shock interaction zone can be accurately fit by the free interaction model.

Shock-Induced Turbulent Boundary Layer Separation in Supersonic Nozzles.

Shock-Induced Turbulent Boundary Layer Separation in Supersonic Nozzles.