Lipeng Lu

An Optimized Low-Dissipation Monotonicity-Preserving Scheme for Numerical Simulations of High-Speed Turbulent Flows

This paper presents an optimized low-dissipation monotonicity-preserving (MP-LD) scheme for numerical simulations of high-speed turbulent flows with shock waves. By using the bandwidth dissipation optimization method (BDOM), the linear dissipation of the original MP scheme of Suresh and Huynh (J. Comput. Phys. 136, 83–99, 1997) is significantly reduced in the newly developed MP-LD scheme. Meanwhile, to reduce the nonlinear dissipation and errors, the shock sensor of Ducros et al. (J. Comput. Phys. 152, 517–549, 1999) is adopted to avoid the activation of the MP limiter in regions away from shock waves. Simulations of turbulent flows with and without shock waves indicate that, in comparison with the original MP scheme, the MP-LD scheme has the same capability in capturing shock waves but a better performance in resolving small-scale turbulence fluctuations without introducing excessive numerical dissipation, which implies the MP-LD scheme is a valuable tool for the direct numerical simulation and large eddy simulation of high-speed turbulent flows with shock waves.

Large-eddy simulation of 3-D corner separation in a linear compressor cascade

The increase of the thrust/weight ratio of aircraft engines is extremely restricted by different 3-D flow loss mechanisms. One of them is the corner separation that can form at the junction between a blade suction side and a hub or shroud. In this paper, in order to further investigate the turbulent characteristics of corner separation, large-eddy simulation (LES) is conducted on a compressor cascade configuration using NACA65 blade profiles (chord based Reynolds number: 3.82 × 105), in comparison with the previous obtained experimental data. Using the shear-improved Smagorinsky model as subgrid-scale model, the LES gives a good description of the mean aerodynamics of the corner separation, especially for the blade surface static pressure coefficient and the total pressure losses. The turbulent dynamics is then analyzed in detail, in consideration of the turbulent structures, the one-point velocity spectra, and the turbulence anisotropy. Within the recirculation region, the energy appears to concentrate around the largest turbulent eddies, with fairly isotropic characteristics. Concerning the dynamics, an aperiodic shedding of hairpin vortices seems to induce an unsteadiness of the separation envelope.

Numerical Study of the Effect of Secondary Vortex on Three-Dimensional Corner Separation in a Compressor Cascade

Abstract The complex flow structures in a linear compressor cascade have been investigated under different incidences using both the Reynolds-averaged Navier–Stokes (RANS) and delayed detached eddy simulation (DDES) methods. The current study analyzes the development of horseshoe vortex and passage vortex in a compressor cascade based on DDES results and explores the effect of the passage vortex on corner separation using the RANS method. Results show that the effect of horseshoe vortex on three-dimensional corner separation is weak, whereas the effect of passage vortex is dominant. A large vortex breaks into many small vortices in the corner separation region, thereby resulting in strong turbulence fluctuation. The passage vortex transports the low-energetic flow near the endwall to the blade suction surface and enlarges corner separation in the cascade. Hence, total pressure loss increases in the cascade.

Unsteady behavior of corner separation in a compressor cascade: Large eddy simulation and experimental study

The present study carried out a pure large eddy simulation (LES) on a NACA65 linear compressor cascade (chord-based Reynolds number: 382 000), at an incidence angle of 4°. In this configuration, a corner separation at the juncture of the blade suction side and the end-wall is clearly observed experimentally. For the simulation, 852 CPUs are used in parallel and particular attention is paid to the inflow conditions, coupling LES with the calculation of the inlet boundary layer and associated velocity fluctuations. The shear-improved Smagorinsky model, which has been proved to be competent for turbomachines, is used in this study. Numerical results are carefully compared with pressure measurements, two-dimensional and stereo particle image velocimetry. LES is found to be superior to Reynolds-averaged Navier–Stokes approach in predicting the three-dimensional separated flows in the compressor cascade since pressure coefficients and losses are in very good agreement with the experiment. LES has been proven to simulate also the unsteady behavior of the separation. Finally, the experimental and numerical analysis of the inflow angle reveals high fluctuations of incidence upstream the leading edge of the blade near the end-wall, mostly associated with the incoming wall boundary layer.

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.

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

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 Gortler-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 numbereffects 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.

Experimental Investigations of Corner Stall in a Linear Compressor Cascade

In applied research, a lack of understanding of corner stall, i.e. the three-dimensional (3D) separation in the juncture of the endwall and blade corner region, which has limited the efficiency and the stability of compressors. Both Reynolds-averaged Navier-Stokes (RANS) and large eddy simulation (LES) still need to be calibrated for turbomachinery applications. In the fundamental research of the turbulent boundary layer (TBL), there are a lot of findings of the effects of curvature and pressure gradients, which also play an important role in physics of corner stall. The purpose of this thesis is (i) to carry out an experiment in a cascade, (ii) to gain a database that could be used to calibrate both RANS and LES, and (iii) to give some basic explanations of corner stall through investigating the TBL on the suction side at the mid-span which is more complex than those in the basic investigations but simpler than those in a real engine. A detailed and accurate experiment of 3D flow field through a linear compressor cascade has been set up. Experimental data were acquired for a Reynolds number of 3.82×10 ^5 based on blade chord and inlet flow conditions. Measurements have been achieved by hot-wire anemometry, pressure taps on blade and endwall, five-hole pressure probe, oil visualization, 2D particle image velocimetry (PIV),and two-component laser Doppler anemometry (LDA). An original and complete database was thus obtained. The TBL on the suction side at mid-span was investigated. The wall-normal negative pressure gradient restrains the separation, on the contrary to its influence in the corner stall. The streamwise adverse pressure gradient can be responsible for the development of Reynolds stresses. The remarkable phenomenon at measurement stations near the trailing edge of blade is that an inflection point occurs in each profile of the mean streamwise velocity. At this inflection point, the magnitudes of the Reynolds stresses reach their maximum values, and the direction of energy diffusion also changes. The velocity field in the corner stall was presented. Bimodal histograms of velocity exist in the experiment. The bimodal points mainly appear in the region around the mean interface of separated flow and non-separated flow. At a bimodal point the local two velocity components are non-independent from each other, due to the aperiodic interplay of two basic modes in the flow field. Two modes were proposed to interpret the physics of bimodal behaviour.

Computational Model for Stall Inception and Nonlinear Evolution in Axial Flow Compressors

This paper presents a computational model for axial compressor stall inception and its nonlinear evolution using unsteady Reynolds-averaged Navier–Stokes equations as an initial boundary value prob...

Large-eddy simulation of a three-dimensional hypersonic shock wave turbulent boundary layer interaction of a single fin

The large-eddy simulation (LES) of a hypersonic flow passing a single-fin at Mach 5 and Re∞=3.7×107/m was conducted and the three-dimensional (3D) shock wave/turbulent boundary layer interaction (SWTBLI) was studied in the present paper. This is probably a first reported LES of this kind of flows. The newly developed seventh order low-dissipation monotonicity-preserving scheme is used to solver the Euler fluxes and the dynamic Smagorinsky subgrid model is used to take account of the subgrid stress and heat flux. The shock system, flow separation structure, and turbulence characteristic are investigated by analyzing the LES data. The turbulence in the 3D SWTBLI is found to be dominated by small-scale wall turbulence, large-scale free shear turbulence, as well as the corner vortex in different regions. In the reverse flow, the streamwise elongated coherent structures are regenerated beneath the main separation vortex, almost immediately after the flow reattachment.

Numerical Analysis of Three-Dimensional Corner Separation in a Linear Compressor Cascade

(1)Laboratoire de Mecanique des Fluides et d’Acoustique, Ecole Centrale de Lyo´ n, 69130 Ecully, France(2)School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, 100191 Beijing, China∗ Feng.Gao@ec-lyon.frABSTRACTNowadays, the internal flow in aircraft engine compressorscan be quite accurately reproduced at design condition by theCFD tools. However, CFD generally fails to simulate some sin-gular 3D phenomena, near off-design conditions, such as thecorner separation. Studies have pointed out that the separationregions are often over-estimated when the flow state is far fromdesign condition, owing to the turbulence model. Much workis devoted to improving the capability of the turbulence modelin capturing the onset and the extent of the corner separation,which is desired in the designing procedures.In this paper, steady RANS simulations are carried out in thesame configuration as an experiment of Ma et al. These simula-tions are obtained with a high-precision in-house Navier-Stokessolver (Turb’Flow). With the same mesh, an unsteady simulation(URANS) is subsequently presented, in order to investigate theinfluence of a fluctuating inflow.Attention is focused on a specific angle of attack of 4 de-grees, for which the three-dimensional corner separation isclearly observed. For the unsteady simulation, unsteadiness isimposed through perturbations of the angle of attack at the inlet.The results ofthe steady andunsteadycomputationsareanalyzedand compared with those of the experiment. The time-averagedURANS results agree well with the RANS results. The fluctuati nginflow does not show much influence on the mean performanceof the compressor cascade. The onset of the corner separationoccurs earlier in the simulations than in the experiment, consid-ering the blade surface pressure and the passage velocity pro-files. However, the cross-stream extent of the corner separationappears slightly under-estimated by CFD, according to the out-let total pressure losses andthe passagevelocityprofiles. Finally,the URANS simulation allows to recover bi-modal PDFs, as ob-served in the experiment.NOMENCLATUREc Chord lengthc