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I am simulating wind tunnel tests for a NACA64210 airfoil with the following conditions:

$P_\infty=101300$ $Pa$ ; $\rho = 1.23$ $kg/m^3$ ; $U_\infty = 29.99$ $m/s$

I performed the simulation with Solidworks Flow Simulation, and I obtained the following plots.

Pressure Plot

Velocity Plot

My question is about if these plots make sense and can be correct.

For instance, I am wondering why in the velocity plot, the velocity is not zero at the leading edge (provided it should be a stagnation point), maybe it's because the wind tunnel speed is low.

On the other hand, as far as I know, the pressure distribution in a wing should be smaller in the upper side of the wing than in the lower to successfully create lift, in this case it seems to be equal at both surfaces. Also, there's usually a suction closer to the leading edge that I cannot detect here.

I just need a review of my plots to know if they are correct or not, and if they are not correct what might be the cause of it.

P.S: The only setting that I was not sure about what to input, was that I enabled that the air was laminar and turbulent. I'm not sure if I should have stated only laminar, instead.

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    $\begingroup$ The stagnation point is a mathematical concept, in reality, the velocity will not be exactly zero. However, it should be very low. I would advise you to determine the Cp, Cl and Cd, and see if they make sense.This will also answer your question about the pressure above and below the foil. $\endgroup$ – ROIMaison Dec 29 '15 at 13:32
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Qualitatively they seem to be correct:

  • Highest pressure at stagnation point and maximal pressure in diagram close to total pressure: $$101843.84\ Pa \approx P_{tot}= P_{\infty}+\frac{\rho}{2}\cdot U_\infty^2 =101853\ Pa$$
  • Speed drop at stagnation point.
  • The slight flow deflection downwards behind the airfoil shows that your airfoil is generating lift (I am assuming your free stream flow is in x-direction).
  • Low pressure (dark blue) region above airfoil is bigger than below.

As @ROIMaison wrote the correct way to validate the results is looking at the aerodynamic coefficients and pressure ditribution $c_p$. In a quick search I found some polars calculated with xfoil you should be able to replicate. With xfoil you can calculate the pressure distribution too.

If you are only interested in lift at small angles of attack laminar flow simulation should be enough. If you also want to calculate drag I would set turbulent air. To validate your setup (model, boundary conditions,...) for turbulent flow there is a great NASA webpage.

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