Thanks for all the advice so far on this problem. I am still having the same problem with the M6 wing validation. The simulations work fine for any subsonic flow, however once the flow goes supersonic over the upper surface, the resulting pressure distribution is incorrect. A plot of the Mach number over the upper surface is attached and it shows 2 distinct regions of high speed flow. I have also done this simulation if fluent and it gives accurate results. in CFX, the region of high speed flow at the leading edge is smaller than fluent giving a smaller region of low pressure at the leading edge than is required. The aft region does not exist in fluent and in CFX it therefore gives a region of low pressure over the aft section where it should not be. I have a Y+ value of 1 or less on the wing.
Any ideas on where to proceed.
Do you know what is causing the 2 regions? A shock? Separation? Something else? That would give you a clue as to where to start looking.
I am not sure what is causing this problem. I have tried diffrent meshes, even without inflation layers, diffrent turbulence models, diffrent turb models. I have left turbulent wall functions on automatic, is it worth trying these.
I can't make heads or tails of your image, by the way.
This is the basic setup I have been using.
Fluids List; Air Ideal Gas
Reference Pressure; 1 atm
Buoyancy Option; Non Buoyant
Domain Motion; Stationary
Heat Transfer Option; Total Energy
Inc Viscous Work Term Select
Turbulence Model; SST
Transitional Turbulence Select
Turbulent Wall Functions; Automatic
Far Field ( Inlet )
Flow Regime Option Subsonic
Mass and Momentum
Cart. Vel. & Pressure
U = 270.715
V = 27.74
W = 0
Turbulence Option; Low Intensity 1%
Heat Transfer, Static Temperature = 263
Downstream ( Outlet )
Flow Regime Option; Subsonic
Mass and Momentum Option;
Average Static Pressure, Relative Pressure; 0 Pa
Pressure Averaging; Average Over Whole Outlet
Inboard ( Symmetry ) Symmetry
Wing ( Wall )
Wall Influence on flow - Wall, no-slip / No-Slip
Heat Transfer - Adiabatic
Global Initial Conditions
Velocity Type; Cartesian
Cartesian Velocity Components; Automatic With Value, U = 270.715 m/s V = 28.74 m/s
Static Pressure Option; Automatic
Turbulence Kinetic Energy Option; Automatic
Turbulence Eddy Dissipation (Select) Option; Automatic
Solver Control Criteria
Advection Scheme Option; High Resolution
Convergence Control; Automatic Timescale
Max. Iterations; 1000
Length Scale Option Aggressive
Convergence Criteria; Residual Type, Max
Residual Target; 1e-6
The domain is a parabolic one with a symmetry plane. the wing is level and the domain is inclined for the angle of attack.
The pressure plot on the surface is fine for subsonic and low angle of attack. At the M0.84 and 5 degrees, the plot on the upper surface disagrees with the wind tunnel data.
The mach number distribution in this region shows two areas of high speed flow. There should be only one region of high speed flow.
Your mach number plot gives the game away - if your contour lines are jagged like that then your mesh is too coarse. You need a finer grid. Your results on this coarse grid are rubbish.
Thanks for the help.
I currently have 280 points on the upper surface, expanding away at a rate of 1.2. There is a boundary layer with ten layers giving a Y+ of 1.
Is it the number of points or the expansion rate that I should be concentrating on.
Do a sensitivity study on all of them.
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