The Experts below are selected from a list of 252 Experts worldwide ranked by ideXlab platform

Alexander A. Zheltovodov - One of the best experts on this subject based on the ideXlab platform.

  • Direct numerical simulation of supersonic turbulent flows around a tandem expansion-Compression Corner
    Physics of Fluids, 2015
    Co-Authors: Jian Fang, Yufeng Yao, Alexander A. Zheltovodov
    Abstract:

    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.

  • DIRECT NUMERICAL SIMULATION OF SUPERSONIC TURBULENT FLOW IN AN EXPANSION-Compression Corner AT MACH=2.9
    2014
    Co-Authors: Jian Fang, Yufeng Yao, Alexander A. Zheltovodov, Aero-engine Aero-thermodynamics
    Abstract:

    Shock-wave/turbulent boundary layer interactions (SWTBLI) are prevalent phenomena in high-speed flights, which could cause large flow separation and high wall heat flux and strong pressure fluctuation, therefore, affect significantly the aero-thermodynamic loads of the vehicles and the performance of propulsion systems. Among all kinds of SWTBLI, the supersonic expansion–Compression Corner is an important flow configuration consisting in the forebody, intake, combustion chamber and nozzle of high-speed flying vehicle1. In an expansion–Compression Corner, the supersonic flow is accelerated and turned through an expansion fan formed at the expansion Corner (EC), then compressed by a shock-wave formed around the Compression Corner (CC)1. The boundary layer in the Compression Corner region can be either attached or separated, depending on the strength of the shock-wave. This kind of flows is complicated and essentially non-equilibrium due to the strong interaction among the turbulence, expansion-wave, and shock-wave.

  • direct numerical simulation of supersonic turbulent flow in an expansion Compression Corner at mach 2 9
    2014
    Co-Authors: Jian Fang, Yufeng Yao, Alexander A. Zheltovodov, Aeroengine Aerothermodynamics
    Abstract:

    Shock-wave/turbulent boundary layer interactions (SWTBLI) are prevalent phenomena in high-speed flights, which could cause large flow separation and high wall heat flux and strong pressure fluctuation, therefore, affect significantly the aero-thermodynamic loads of the vehicles and the performance of propulsion systems. Among all kinds of SWTBLI, the supersonic expansion–Compression Corner is an important flow configuration consisting in the forebody, intake, combustion chamber and nozzle of high-speed flying vehicle1. In an expansion–Compression Corner, the supersonic flow is accelerated and turned through an expansion fan formed at the expansion Corner (EC), then compressed by a shock-wave formed around the Compression Corner (CC)1. The boundary layer in the Compression Corner region can be either attached or separated, depending on the strength of the shock-wave. This kind of flows is complicated and essentially non-equilibrium due to the strong interaction among the turbulence, expansion-wave, and shock-wave.

  • advances in cfd prediction of shock wave turbulent boundary layer interactions
    Progress in Aerospace Sciences, 2003
    Co-Authors: Doyle Knight, Hong Yan, Argyris G Panaras, Alexander A. Zheltovodov
    Abstract:

    Abstract The paper presents a summary of recent computational fluid dynamics (CFD) simulations of shock wave turbulent boundary layer interactions. This survey was prepared as part of the activity of NATO RTO Working Group 10 which was established in December 1998, and considers results obtained subsequent to the previous survey paper on the same topic by Knight and Degrez (“Shock Wave Boundary Layer Interactions in High Mach Number Flows—A Critical Survey of Current CFD Prediction Capabilities”, AGARD Advisory Report AR-319, Volume II, December 1998). Five configurations are considered: 2-D Compression Corner, 2-D shock impingement, 2-D expansion–Compression Corner, 3-D single fin and 3-D double fin. Recent direct numerical simulations (DNS), large eddy simulations (LES) and Reynolds-averaged Navier–Stokes (RANS) simulations are compared with experiment. The capabilities and limitations are described, and future research needs identified.

  • Large Eddy Simulation of Supersonic Turbulent Flow in Expansion-Compression Corner
    2001
    Co-Authors: Doyle Knight, Hong Yan, Alexander A. Zheltovodov
    Abstract:

    Abstract : A Large Eddy Simulation (LES) methodology has been developed for supersonic turbulent flows with strong shock boundary layer interaction. Results are presented for an expansion-Compression Corner at Mach 3 and compared with experimental data.

Doyle Knight - One of the best experts on this subject based on the ideXlab platform.

  • advances in cfd prediction of shock wave turbulent boundary layer interactions
    Progress in Aerospace Sciences, 2003
    Co-Authors: Doyle Knight, Hong Yan, Argyris G Panaras, Alexander A. Zheltovodov
    Abstract:

    Abstract The paper presents a summary of recent computational fluid dynamics (CFD) simulations of shock wave turbulent boundary layer interactions. This survey was prepared as part of the activity of NATO RTO Working Group 10 which was established in December 1998, and considers results obtained subsequent to the previous survey paper on the same topic by Knight and Degrez (“Shock Wave Boundary Layer Interactions in High Mach Number Flows—A Critical Survey of Current CFD Prediction Capabilities”, AGARD Advisory Report AR-319, Volume II, December 1998). Five configurations are considered: 2-D Compression Corner, 2-D shock impingement, 2-D expansion–Compression Corner, 3-D single fin and 3-D double fin. Recent direct numerical simulations (DNS), large eddy simulations (LES) and Reynolds-averaged Navier–Stokes (RANS) simulations are compared with experiment. The capabilities and limitations are described, and future research needs identified.

  • Large Eddy Simulation of Supersonic Turbulent Flow in Expansion-Compression Corner
    2001
    Co-Authors: Doyle Knight, Hong Yan, Alexander A. Zheltovodov
    Abstract:

    Abstract : A Large Eddy Simulation (LES) methodology has been developed for supersonic turbulent flows with strong shock boundary layer interaction. Results are presented for an expansion-Compression Corner at Mach 3 and compared with experimental data.

  • Large Eddy Simulation of Supersonic Compression Corner Using ENO Scheme
    2001
    Co-Authors: Hong Yan, Doyle Knight, Alexander A. Zheltovodov
    Abstract:

    Abstract : A Large Eddy Simulation of a 25 deg Compression Corner at M = 2.88 and Re(delta) - 2 x 10(exp 4) is performed using an Essentially Non Oscillatory (ENO) scheme. The Favre filtered compressible Navier-Stokes equations are solved using a Monotone Integrated Large Eddy Simulation (MILES) technique on an unstructured grid of tetrahedral cells. The mean flow variables and turbulent shear stress at the incoming flow are in good agreement with experiment and DNS. The separation length scaled by the characteristic scale shows agreement with the experiment. No pronounced pressure plateau is observed compared with experiment at higher Reynolds number.

  • Large Eddy Simulation of a Supersonic Compression Corner Part I
    38th Aerospace Sciences Meeting and Exhibit, 2000
    Co-Authors: Gerald Urbin, Doyle Knight, Alexander A. Zheltovodov
    Abstract:

    A preliminary LES study of the interaction of a Mach 3 adiabatic turbulent boundary layer with the shock system generated by a 25’ Compression Corners is presented. The filtered compressible Navier-Stokes equations are solved on a three-dimensional unstructured grid of tetrahedral cells using a finite volume formulation. The inflow conditions are produced by a resealing-reintroducing method. The subgrid scale motions are modeled using the Monotone Integrated Large Eddy Simulation (MILES) method. The incident shock causes separation of the boundary layer upstream of the point of impingement of the primary shock, with a secondary shock forming near the separation. Qualitative data on the flowfield structure are presented.

  • Structure of supersonic turbulent flow past a swept Compression Corner
    AIAA Journal, 1992
    Co-Authors: Doyle Knight, C. C. Horstman, Seymour M Bogdonoff
    Abstract:

    The structure of the shock wave/turbulent boundary-layer interaction generated by a 3D swept Compression Corner has been investigated through a combined experimental and theoretical research program. The flowfield geometry is defined by the streamwise Compression angle alpha and the sweep angle lambda of the Corner. The present study examines two different configurations, namely (alpha, lambda) = (24 deg, 40 deg) and (24 deg, 60 deg) at Mach 3 and Re sigma infinity about 9 x 10 exp 5. The theoretical model is the 3D Reynolds-averaged compressible Navier-Stokes equations with turbulence incorporated using a turbulent eddy viscosity. The calculated flowfields display general agreement with experimental data for surface pressure and good agreement with experimental flowfield profiles of pitot pressure and yaw angle. The principal feature of the flowfield is a large vortical structure approximately aligned with the Corner. The entrainment of incoming fluid into the vortical structure is strongly affected by the sweep angle lambda. Viscous (turbulent and molecular) effects appear to be important only in the immediate vicinity of the surface and in an isolated region within the interaction and near the Corner.

Eric C. Marineau - One of the best experts on this subject based on the ideXlab platform.

  • amplification and structure of streamwise velocity fluctuations in Compression Corner shock wave turbulent boundary layer interactions
    Journal of Fluid Mechanics, 2019
    Co-Authors: M. A. Mustafa, Nicholaus J. Parziale, M. S. Smith, Eric C. Marineau
    Abstract:

    In this work, we study the effect of the Compression-Corner angle on the streamwise turbulent kinetic energy (sTKE) and structure in Mach 2.8 flow. Krypton tagging velocimetry (KTV) is used to investigate the incoming turbulent boundary layer and flow over , , and Compression Corners. The experiments were performed in a 99 % and 1 % Kr gas mixture in the Arnold Engineering Development Complex (AEDC) Mach 3 Calibration Tunnel (M3CT) at . A figure of merit is defined as the wall-normal integrated sTKE ( ), which is designed to identify turbulence amplification by accounting for the root-mean-squared (r.m.s.) velocity fluctuations and shear-layer width for the different geometries. We observe that the increases as an exponential with the Compression-Corner angle near the root when normalized by the boundary-layer value. Additionally, snapshot proper orthogonal decomposition (POD) is applied to the KTV results to investigate the structure of the flow. From the POD results, we extract the dominant flow structures and compare each case by presenting mean-velocity maps that correspond to the largest positive and negative POD mode coefficients. Finally, the POD spectrum reveals an inertial range common to the boundary-layer and each Compression-Corner flow that is present after the first dominant POD modes.

  • Amplification and structure of streamwise-velocity fluctuations in Compression-Corner shock-wave/turbulent boundary-layer interactions
    Journal of Fluid Mechanics, 2019
    Co-Authors: M. A. Mustafa, Nicholaus J. Parziale, M. S. Smith, Eric C. Marineau
    Abstract:

    In this work, we study the effect of the Compression-Corner angle on the streamwise turbulent kinetic energy (sTKE) and structure in Mach 2.8 flow. Krypton tagging velocimetry (KTV) is used to investigate the incoming turbulent boundary layer and flow over , , and Compression Corners. The experiments were performed in a 99 % and 1 % Kr gas mixture in the Arnold Engineering Development Complex (AEDC) Mach 3 Calibration Tunnel (M3CT) at . A figure of merit is defined as the wall-normal integrated sTKE ( ), which is designed to identify turbulence amplification by accounting for the root-mean-squared (r.m.s.) velocity fluctuations and shear-layer width for the different geometries. We observe that the increases as an exponential with the Compression-Corner angle near the root when normalized by the boundary-layer value. Additionally, snapshot proper orthogonal decomposition (POD) is applied to the KTV results to investigate the structure of the flow. From the POD results, we extract the dominant flow structures and compare each case by presenting mean-velocity maps that correspond to the largest positive and negative POD mode coefficients. Finally, the POD spectrum reveals an inertial range common to the boundary-layer and each Compression-Corner flow that is present after the first dominant POD modes.

Ulrich Rist - One of the best experts on this subject based on the ideXlab platform.

  • Two-dimensional numerical investigations of small-amplitude disturbances in a boundary layer at Ma=4.8: Compression Corner versus impinging shock wave
    Physics of Fluids, 2004
    Co-Authors: Alessandro Pagella, Andreas Babucke, Ulrich Rist
    Abstract:

    Two-dimensional direct numerical simulations and linear stability theory investigations have been carried out for a Compression ramp at Ma54.8 and compared to earlier results of a laminar boundary layer with impinging shock wave. The inflow parameters in both flows were identical; the ramp angle of the Compression Corner was chosen to cause a separation bubble, which has exactly the same length compared to the case with impinging shock. It turned out, that the two cases are almost identical for the base flow properties. This is in accordance with similarity assumptions, e.g., free interaction theory, which for smaller Reynolds numbers states, that the boundary layer should be independent of the sort of shock-boundary layer interaction. However, linear stability theory results differ near the Corner and the impinging shock, respectively. Direct numerical simulations of small-amplitude disturbances, which were introduced into the laminar boundary layer, also behave in a very similar way. Amplitude distributions exhibit the same characteristics. The according distributions of the ramp flow have slightly larger amplitudes than the case with impinging shock. © 2004 American Institute of Physics. @DOI: 10.1063/1.1738414#

Sudhir L. Gai - One of the best experts on this subject based on the ideXlab platform.

  • Hypersonic Compression Corner flow with large separated regions
    Journal of Fluid Mechanics, 2019
    Co-Authors: Sudhir L. Gai, Amna Khraibut
    Abstract:

    The structure of large-scale hypersonic boundary layer separation and reattachment is studied numerically using a flat plate/Compression Corner geometry. Apart from verifying the large scale separation characteristics in hypersonic flow, a detailed discussion of secondary separation and fragmentation into multiple vortices embedded within the main recirculation region is presented. The unique relation between the second minimum in shear stress and the scaled angle is highlighted in the context of the reverse flow singularity of Smith ( Proc. R. Soc. Lond.  A, vol. A420, 1988, pp. 21–52) and it appears that for a small wall temperature ratio, such a singularity is unlikely. It is shown that the size of the separation can be estimated in terms of Burggraf’s expression based on asymptotic theory.

  • Establishment of steady separated flow over a Compression-Corner in a free-piston shock tunnel
    Shock Waves, 1997
    Co-Authors: Samuel George Mallinson, Sudhir L. Gai, Neil Mudford
    Abstract:

    The time required to establish steady separated Compression–Corner flow is examined under hypervelocity conditions in a free-piston shock tunnel. This time is reasonably well described using previous perfect gas analyses. The results suggest that, provided the nozzle reservoir enthalpy is 20 MJ kg\(^{-1}\) or less, there is sufficient time to establish steady separated flow before driver gas contamination becomes a significant problem in the present facility.

  • The interaction of a shock wave with a laminar boundary layer at a Compression Corner in high-enthalpy flows including real gas effects
    Journal of Fluid Mechanics, 1997
    Co-Authors: Samuel George Mallinson, Sudhir L. Gai, Neil Mudford
    Abstract:

    The high-enthalpy, hypersonic flow over a Compression Corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg -1 to 19 MJ kg -1 , with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7 x 10 4 to 2.7 x 10 5 . The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the Compression Corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction.

  • High-enthalpy, hypersonic Compression Corner flow
    AIAA Journal, 1996
    Co-Authors: Samuel George Mallinson, Sudhir L. Gai, Neil Mudford
    Abstract:

    The results of an experimental investigation of high-enthalpy, hypersonic flow over sharp leading-edge Compression Corners are presented and discussed. In particular, the possible effects of real gas behavior are examined. Measurements have been made of the heat transfer and pressure distributions for flat plate and Compression Corner flow. Some flow visualization data have also been obtained. Test flows were generated using a free-piston shock tunnel operating in the reflected mode. The reservoir enthalpy ranged from 3 to 19 MJ kg -1 , giving freestream speeds of 2.3-5.5 km s -1 . For these conditions, the flow remains laminar throughout. The flat plate data for both high- and low-enthalpy flows are in agreement with the reference enthalpy method for heat transfer and the weak interaction theory for pressure. Also, the measured flat plate boundary-layer thickness compares well with an expression strictly valid for perfect gas flows only. The high- and low-enthalpy Compression Corner flows have upstream influence and plateau pressure behavior similar to perfect gas flow. That is, real gas effects for the present flows appear to be negligible. This is consistent with the essentially chemically frozen viscous and inviscid flow upstream of the interaction.

  • Leading-Edge Bluntness Effects in High Enthalpy, Hypersonic Compression Corner Flow
    AIAA Journal, 1996
    Co-Authors: Samuel George Mallinson, Sudhir L. Gai, Neil Mudford
    Abstract:

    An experimental study of the combined effects of leading-edge bluntness and real gas behavior on shock wave/boundary-layer interaction has been performed. Pressure and heat transfer distributions have been measured over a Compression Corner for a range of Corner angles, including the datum case of flat plate flow. On the flat plate and upstream of the Corner, the pressure and heat transfer for the blunt leading-edge configuration were found to lie above the corresponding sharp leading values. On the ramp, there was a considerable reduction in the pressure and heat transfer when the leading edge was blunt. Also, the extent of the interaction was seen to be smaller with the blunt leading edge. These results are similar to those from perfect gas studies. The differences between the sharp and blunt leading-edge data appear to be less pronounced at the higher enthalpy. This is thought likely to be because of the reduced shock standoff, which occurs as a result of dissociation. A comparison between the heat transfer data on the ramp face and predictions from the generalized reference enthalpy theory was seen to be reasonable. The upstream influence and plateau pressure were found to be in fair agreement with data from low enthalpy experiments. A C CD c d h hr L lu M n Pr p qw Rex