Investigation of three-dimensional shock wave/turbulent-boundary-layer interaction initiated by a single fin

Three-dimensional shock wave/turbulent-boundary-layer interaction of a hypersonic flow passing a single fin mounted on a flat plate at a Mach number of five and unit Reynolds number 3.7×10^7 was conducted by a large-eddy simulation approach. The performed large-eddy simulation has demonstrated good agreement with experimental data in terms of mean flowfield structures, surface pressure distribution, and surface flow pattern. Furthermore, the shock wave system, flow separation structure, and turbulence characteristics were all investigated by analyzing the obtained large-eddy simulation dataset. It was found that, for this kind of three-dimensional shock wave/turbulent-boundary-layer interaction problem, the flow characteristics in different regions have been dominated by respective wall turbulence, free shear layer turbulence, and corner vortex motions in different regions. In the reverse flow region, near-wall quasi-streamwise streaky structures were observed just beneath the main separation vortex, indicating that the transition of the pathway of the separation flow to turbulence may occur within a short distance from the reattachment location. The obtained large-eddy simulation results have provided a clear and direct evidence of the primary reverse flow and the secondary separation flow being essentially turbulent

On the turbulence amplification in shock-wave/turbulent boundary layer interaction

The mechanism of turbulence amplification in shock-wave/boundary layer interactions is reviewed, and a new turbulence amplification mechanism is proposed based on the analysis of data from direct numerical simulation of an oblique shock-wave/flat-plate boundary layer interaction at Mach 2.25. In the upstream part of the interaction zone, the amplification of turbulence is not essentially shear driven, but induced by the interaction of the deceleration of mean flow with streamwise velocity fluctuations, which causes a rapid increase of turbulence intensity in the near-wall region. In the downstream part of the interaction zone, the high turbulence intensity is mainly due to the free shear layer generated in the interaction zone. During the initial stage of turbulence amplification, the characteristics of wall turbulence, including compact velocity streaks, streamwise vortices and an anisotropic Reynolds stress, are well preserved. The mechanism proposed explains the high level of turbulence in the near-wall region observed in some experiments and numerical simulations

Direct numerical simulation of supersonic turbulent flows around a tandem expansion-compression corner

© 2015 AIP Publishing LLC. The M = 2.9 supersonic turbulent flows over a tandem expansion-compression corner configuration with a sharp deflection angle of 25° at three Reynolds numbers Reδ = 20 000, 40 000, and 80 000 were studied by using direct numerical simulation. The flow statistics were validated against available experimental measurements and other numerical predictions. The flow structures and turbulence statistics were detailed visualized and analysed for the Reδ = 40 000 case, especially in the interaction region where flow separation and reattachment occurred. It was found that during the expansion process, the boundary layer exhibited a characteristic two-layer structure also discovered in previous experimental studies, and the turbulence evolved differently within these two layers. In the outer layer, the turbulence was consistently suppressed along the ramp to a large extent, while in the inner layer, it was suppressed only in a small region around the expansion corner, and the near-wall quasi-streamwise vortices were well preserved. Flow patterns near the reattachment line have shown the existence of the Görtler-type vortices, which would largely amplify turbulence fluctuations in the near-wall region, thus promoting the regeneration of wall turbulence that in turn contributed to the redevelopment of a downstream turbulent boundary layer. The Reynolds number effects and the characteristics of coherent structures were also discussed. With the increase of the Reynolds number, the separation bubble size decreased, but the pattern and the characteristic size of wall streamlines near the reattachment line were preserved

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 becomes\ 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.Comment: 34 pages; 9 figure

Large-eddy simulation of shock-wave/turbulent boundary layer interaction with and without SparkJet control

© 2016 Chinese Society of Aeronautics and Astronautics. Production and hosting by Elsevier Ltd. The efficiency and mechanism of an active control device "SparkJet" and its application in shock-induced separation control are studied using large-eddy simulation in this paper. The base flow is the interaction of an oblique shock-wave generated by 8° wedge and a spatially-developing Ma = 2.3 turbulent boundary layer. The Reynolds number based on the incoming flow property and the boundary layer displacement thickness at the impinging point without shock-wave is 20000. The detailed numerical approaches were presented. The inflow turbulence was generated using the digital filter method to avoid artificial temporal or streamwise periodicity. The numerical results including velocity profile, Reynolds stress profile, skin friction, and wall pressure were systematically validated against the available wind tunnel particle image velocimetry (PIV) measurements of the same flow condition. Further study on the control of flow separation due to the strong shock-viscous interaction using an active control actuator "SparkJet" was conducted. The single-pulsed characteristic of the device was obtained and compared with the experiment. Both instantaneous and time-averaged flow fields have shown that the jet flow issuing from the actuator cavity enhances the flow mixing inside the boundary layer, making the boundary layer more resistant to flow separation. Skin friction coefficient distribution shows that the separation bubble length is reduced by about 35% with control exerted

Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed

Documentation of Experimental Data for Hypersonic 3-D Shock Waves

Experimental data for a series of three-dimensional hypersonic swept- and crossing-shock-waves / turbulent boundary layer interaction flows in the vicinity of one and two sharp fins mounted on a flat plate are presented. The data contains detailed documentation of incoming boundary layer, surface pressure and heat flux distributions, skin friction distributions and topological features of the limiting streamlines in the interaction region at M = 5. The deflection angles ß were varied for swept-shock interactions (single fin geometry) from 0° to 27°, and for crossing-shock interaction (double fin geometry) from 8° to 23°