Jetting in Strong Shock Reflections Through Low Isentropic Exponent Gases: Experiments and Navier-Stokes Simulations

2019 ◽  
pp. 769-776
Author(s):  
S. SM. Lau-Chapdelaine ◽  
Q. Xiao ◽  
M. I. Radulescu
Keyword(s):  
Strong Shock ◽  
Navier Stokes ◽  
Isentropic Exponent ◽  
Shock Reflections

scholarly journals Choke flutter instability sources tracking with linearized calculations

2019 ◽  
Vol 30 (9) ◽  
pp. 4155-4166
Author(s):  
Pierre Duquesne ◽  
Quentin Rendu ◽  
Stephane Aubert ◽  
Pascal Ferrand
Keyword(s):  
Shock Wave ◽  
Superposition Principle ◽  
Strong Shock ◽  
Source Tracking ◽  
Navier Stokes ◽  
Test Case ◽  
Navier Stokes Equation ◽  
Local Vibration ◽  
Limited Region ◽  
Content Type

Purpose The choke flutter is a fluid-structure interaction that can lead to the failure of fan or compressor blade in turbojet engines. In ultra high bypass ratio (UHBR) fans, the choke flutter appears at part-speed regimes and at low or negative incidence when a strong shock-wave chokes the blade to blade channel. The purpose of this study is to locate the main excitation sources and improving the understanding of the different work exchange mechanisms. This work contributes to avoiding deficient and dangerous fan design. Design/methodology/approach In this paper, an UHBR fan is analyzed using a time-linearized Reynolds-averaged Navier–Stokes equation solver to investigate the choke flutter. The steady-state and the imposed vibration (inter blade phase angle, reduced frequency and mode shape) are selected to be in choke flutter situation. Superposition principle induced by the linearization allow to decompose the blade in numerous small subsections to track the contribution of each local vibration to the global damping. All simulations have been performed on a two-dimensional blade to blade extraction. Findings Result analysis points to a restricted number of excitation sources at the trailing edge which induce a large part of the work exchange in a limited region of the airfoil. Main phenomena suspected are the shock-wave motion and the shock-wave/boundary layer interaction. Originality/value An original excitation source tracking methodology allowed by the linearized calculation is addressed and applied to a UHBR fan test case.

The persistence of regular reflection during strong shock diffraction over rigid ramps

2001 ◽  
Vol 431 ◽  
pp. 273-296 ◽  
Author(s):  
L. F. HENDERSON ◽  
K. TAKAYAMA ◽  
W. Y. CRUTCHFIELD ◽  
S. ITABASHI
Keyword(s):  
Stokes Equations ◽  
Mach Reflection ◽  
Strong Shock ◽  
Navier Stokes ◽  
Regular Reflection ◽  
Navier Stokes Equations ◽  
Von Neumann ◽  
Initial Appearance ◽  
Self Similar ◽  
Initial Shock

We report on calculations and experiments with strong shocks diffracting over rigid ramps in argon. The numerical results were obtained by integrating the conservation equations that included the Navier–Stokes equations. The results predict that if the ramp angle θ is less than the angle θe that corresponds to the detachment of a shock, θ < θe, then the onset of Mach reflection (MR) will be delayed by the initial appearance of a precursor regular reflection (PRR). The PRR is subsequently swept away by an overtaking corner signal (cs) that forces the eruption of the MR which then rapidly evolves into a self-similar state. An objective was to make an experimental test of the predictions. These were confirmed by twice photographing the diffracting shock as it travelled along the ramp. We could get a PRR with the first exposure and an MR with the second. According to the von Neumann perfect gas theory, a PRR does not exist when θ < θe. A viscous length scale xint is a measure of the position on the ramp where the dynamic transition PRR → MR takes place. It is significantly larger in the experiments than in the calculations. This is attributed to the fact that fluctuations from turbulence and surface roughness were not modelled in the calculations. It was found that xint → ∞ as θ → θe. Experiments were done to find out how xint depended on the initial shock tube pressure p0. The dependence was strong but could be greatly reduced by forming a Reynolds number based on xint. Finally by definition, regular reflection (RR) never interacts with a boundary layer, while PRR always interacts; so they are different phenomena.

Experimental and Numerical Study of Shock Wave Interaction with Perforated Plates

2004 ◽  
Vol 126 (3) ◽  
pp. 399-409 ◽  
Author(s):  
A. Britan ◽  
A. V. Karpov ◽  
E. I. Vasilev ◽  
O. Igra ◽  
G. Ben-Dor ◽  
...  
Keyword(s):  
Shock Wave ◽  
Numerical Study ◽  
Wave Interaction ◽  
Wave Pattern ◽  
Incident Shock ◽  
Incident Shock Wave ◽  
Navier Stokes ◽  
Navier Stokes Equation ◽  
Perforated Plates ◽  
Shock Reflections

The flow developed behind shock wave transmitted through a screen or a perforated plat is initially highly unsteady and nonuniform. It contains multiple shock reflections and interactions with vortices shed from the open spaces of the barrier. The present paper studies experimentally and theoretically/numerically the flow and wave pattern resulted from the interaction of an incident shock wave with a few different types of barriers, all having the same porosity but different geometries. It is shown that in all investigated cases the flow downstream of the barrier can be divided into two different zones. Due immediately behind the barrier, where the flow is highly unsteady and nonuniform in the other, placed further downstream from the barrier, the flow approaches a steady and uniform state. It is also shown that most of the attenuation experienced by the transmitted shock wave occurs in the zone where the flow is highly unsteady. When solving the flow developed behind the shock wave transmitted through the barrier while ignoring energy losses (i.e., assuming the fluid to be a perfect fluid and therefore employing the Euler equation instead of the Navier-Stokes equation) leads to non-physical results in the unsteady flow zone.

Investigation of Shock-Wave/Boundary-Layer Interaction on Aeroelastic Stability: Towards Active Control

2016 ◽  
Author(s):  
Quentin Rendu ◽  
Yannick Rozenberg ◽  
Stéphane Aubert ◽  
Pascal Ferrand
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Pressure Fluctuations ◽  
Strong Shock ◽  
Separated Flow ◽  
Navier Stokes ◽  
Pressure Waves ◽  
Aeroelastic Stability ◽  
Design And Optimization ◽  
Boundary Layer Interaction

In order to predict oscillating loads on a structure, time-linearized methods are fast enough to be routinely used in design and optimization steps of a turbomachine stage. In this work, frequency-domain time-linearized Navier-Stokes computations are proposed to predict the unsteady separated flow generated by an oscillating bump in a transonic nozzle. We also investigate the interaction of backward traveling pressure waves and moving surface on the unsteady behavior of a strong shock-wave with separated boundary-layer. This case is representative of transonic stall flutter of a compressor blade submitted to downstream stator potential effects. The influence of frequency is first investigated on a generic oscillating bump to identify the most unstable configuration. Introducing back pressure fluctuations, we then show that the aeroelastic stability of the system depends on the phase-shift between the fluctuations and the bump motion. Finally, we propose to actively control the instability by generating backward traveling pressure waves at prescribed amplitude, frequency and phase.

A Hybrid Lattice Boltzmann Flux Solver for Simulation of Viscous Compressible Flows

2016 ◽  
Vol 8 (6) ◽  
pp. 887-910 ◽  
Author(s):  
L. M. Yang ◽  
C. Shu ◽  
J. Wu
Keyword(s):  
Lattice Boltzmann ◽  
Compressible Flows ◽  
Stokes Equations ◽  
Strong Shock ◽  
Viscous Flows ◽  
Lattice Boltzmann Model ◽  
Navier Stokes ◽  
Inviscid Flows ◽  
Navier Stokes Equations ◽  
Viscous Compressible Flows

AbstractIn this paper, a hybrid lattice Boltzmann flux solver (LBFS) is proposed for simulation of viscous compressible flows. In the solver, the finite volume method is applied to solve the Navier-Stokes equations. Different from conventional Navier-Stokes solvers, in this work, the inviscid flux across the cell interface is evaluated by local reconstruction of solution using one-dimensional lattice Boltzmann model, while the viscous flux is still approximated by conventional smooth function approximation. The present work overcomes the two major drawbacks of existing LBFS [28–31], which is used for simulation of inviscid flows. The first one is its ability to simulate viscous flows by including evaluation of viscous flux. The second one is its ability to effectively capture both strong shock waves and thin boundary layers through introduction of a switch function for evaluation of inviscid flux, which takes a value close to zero in the boundary layer and one around the strong shock wave. Numerical experiments demonstrate that the present solver can accurately and effectively simulate hypersonic viscous flows.

Shocks generated in a confined gas due to rapid heat addition at the boundary. II. Strong shock waves

1984 ◽  
Vol 393 (1805) ◽  
pp. 331-351 ◽  
Keyword(s):  
Shock Wave ◽  
Ad Hoc ◽  
Numerical Solutions ◽  
Stokes Equations ◽  
Strong Shock ◽  
Mean Free Path ◽  
Present Theory ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Heat Addition

A study is made of the birth and evolution of a strong shock wave in an inert gas due to rapid energy deposition at a boundary. The gas is confined between infinite parallel planes separated by a distance large compared with the molecular mean free path. Heat flux at the wall rises from zero to a finite constant value during an interval that is a modest multiple of the intermolecular collision time. The thermomechanical response of the gas near the boundary is described by the complete Navier-Stokes equations in a layer with a thickness that is a few molecular mean free paths. Numerical solutions show how a spatial pressure variation is generated adjacent to the boundary, which then propagates away as an almost steady shock wave. If heat addition is continued a thicker high-temperature expanding layer develops in which the pressure remains uniform. This expanding layer acts like a piston, or a contact surface, the speed of which is calculated to leading order. In this way the present theory provides a rational basis for the ad hoc piston analogies used by earlier authors. In particular it shows the importance of power as a crucial factor in the determination of shock strength.

A Flux-Split Solution Procedure for Unsteady Inlet Flow Calculations

1992 ◽  
Vol 114 (2) ◽  
pp. 198-204 ◽  
Author(s):  
H. S. Pordal ◽  
P. K. Khosla ◽  
S. G. Rubin
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Unsteady Flow ◽  
Sparse Matrix ◽  
Strong Shock ◽  
Solution Procedure ◽  
Navier Stokes ◽  
Time Varying ◽  
Two Dimensional ◽  
Boundary Layer Interaction

The solution of the reduced Navier Stokes (RNS) equations is considered using a flux-split procedure. Unsteady flow in a two dimensional engine inlet is computed. The problems of unstart and restart are investigated. A sparse matrix direct solver combined with a domain decomposition strategy is used to compute the unsteady flow field. Strong shock-boundary layer interaction, time varying shocks, and time varying recirculation regions are efficiently captured.

Unsteady Navier-Stokes Simulation of a Transonic Flutter Cascade Near Stall Conditions Applying Algebraic Transition Models

2005 ◽  
Author(s):  
Hans Thermann ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Stokes Equations ◽  
Strong Shock ◽  
Turbulence Models ◽  
Navier Stokes ◽  
Implicit Time ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Transition Models ◽  
Operating Points

Due to the trend in the design of modern aeroengines to reduce weight and to realize high pressure ratios, fan and first stage compressor blades are highly susceptible to flutter. At operating points with transonic flow velocities and high incidences stall flutter might occur involving strong shock-boundary layer interactions, flow separation and oscillating shocks. In this paper, results of unsteady Navier-Stokes flow calculations around an oscillating blade in a linear transonic compressor cascade at different operating points including near stall conditions are presented. The nonlinear unsteady Reynolds-averaged Navier-Stokes equations are solved time-accurately using implicit time-integration. Different Low-Reynolds-Number turbulence models are used for closure. Furthermore, empirical algebraic transition models are applied to enhance the accuracy of prediction. Computations are performed two-dimensionally as well as three-dimensionally. It is shown that, for the steady calculations, the prediction of the boundary layer development and the blade loading can be substantially improved compared with fully turbulent computations when algebraic transition models are applied. Furthermore, it is shown that the prediction of the aerodynamic damping in the case of oscillating blades at near stall conditions can be dependent on the applied transition models.

Unsteady Navier-Stokes Simulation of a Transonic Flutter Cascade Near-Stall Conditions Applying Algebraic Transition Models

2005 ◽  
Vol 128 (3) ◽  
pp. 474-483 ◽  
Author(s):  
Hans Thermann ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Stokes Equations ◽  
Strong Shock ◽  
Turbulence Models ◽  
Navier Stokes ◽  
Implicit Time ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Transition Models ◽  
Operating Points

Due to the trend in the design of modern aeroengines to reduce weight and to realize high pressure ratios, fan and first-stage compressor blades are highly susceptible to flutter. At operating points with transonic flow velocities and high incidences, stall flutter might occur involving strong shock-boundary layer interactions, flow separation, and oscillating shocks. In this paper, results of unsteady Navier-Stokes flow calculations around an oscillating blade in a linear transonic compressor cascade at different operating points including near-stall conditions are presented. The nonlinear unsteady Reynolds-averaged Navier-Stokes equations are solved time accurately using implicit time integration. Different low-Reynolds-number turbulence models are used for closure. Furthermore, empirical algebraic transition models are applied to enhance the accuracy of prediction. Computations are performed two dimensionally as well as three dimensionally. It is shown that, for the steady calculations, the prediction of the boundary layer development and the blade loading can be substantially improved compared with fully turbulent computations when algebraic transition models are applied. Furthermore, it is shown that the prediction of the aerodynamic damping in the case of oscillating blades at near-stall conditions can be dependent on the applied transition models.

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