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Supersonic flow over an airfoil. FLUENT Lift don't match up with Shock Expansion Tech

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Old   March 25, 2012, 15:02
Default Supersonic flow over an airfoil. FLUENT Lift don't match up with Shock Expansion Tech
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Hi everybody,
I’m a student of Beng(Hons) Aero Mech engineering. I’m involved in an undergraduate project focused in Compressible flow/ Supersonic ( M=1.5/2.5). Air flow over a Double wedge airfoil. After that 3D.

What I’m doing firstly is comparing Lif&Drag ratios against the values calculated with the Shock Expansion Technique. Once the 2D tests done in double wedged airfoils will be correct and compared with those ones calculated by hand (excel sheet) I will validate them to move on 3D models.
I know that my FLUENT configuration is rough and it could be difficult to get correct Lif&Drag ratios.

The first test I’m doing is over a symmetrical double wedge airfoil. t/c=4% with c=100. AoA 4 degrees.
I’ve already designed the farfield and the BC’s under GAMBIT. With a NURBS “cone” due to refine the mesh in that area using Bi-exponent method. Pressure farfields are chosen for the BC’s. The final mesh has more than 400.000 elements, aspect ratio 1-3 and the most skewed element has 0.7 ratio. (Figure1)

It’s not the best but I think that my laptop can “die” if I mesh it more accurately when I iterate under FLUENT. And I’m an undergraduate student and in order to the amount af tests I would like to do I thought that it could be a good start point.
Let’s go to the main point. My FLUENT configuration is:
Scaled correctly / Reorder domain till get 1 / yplus around 150 ( in order to the mesh, not the best)
  • Density Based solver
  • K-epsilon viscosity method
  • Ideal gas / Sutherland viscosity
  • Operating pressure: 0 (supersonic)
  • Gauge pressure ( Pressure Farfield BC’s ) : aprox. 26436.26 Pa (10.000 meters altitude hypothesis)
  • Initialize / Iterate.

Not messages given back during the iterations. Everything correct. BUT NEVER CONVERGES.(Figure2)
I can bear the no convergence event BUT:
Report/forces in lift I get 0.37 while in the Shock Expansion Technique (Hand calculation) I get 0.17 which is correct and reliable. So it’s like the double and not valid.(Figure 3)

I know that the Shock Expansion Technique only gives the Wave Drag but in Lifts I should get encouraging results. Not the double.

I tried to describe accurately the configuration set up in order that somebody could understand what I’m doing. Hope it helps.

Could anybody give some clue in how to set a Supersonic flow over an airfoil? (I’ve already done the Cornwell University tutorials.)
Thank you in advance,
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Old   August 21, 2013, 07:43
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I think gauge pressure must be derived using p0/p at your respective no......
I'm also doing same thing [2d then 3d and mach no. range] but on different shape of diamond airfoil.....altough I'm also a newbie
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