how to get more accurate accurate coefficient of aerodynamic of subsonic airfoil
I donīt know what can i do to have an accurate coefficient of aerodynamic of airfoil, at a reynolds number of 3*10^6. I try to do several simulations,
with different types of turbulence models, such as spallart-allmaras, k-epsilon,
k-omega and even reynolds stress, but the outcome was not accurate enough, taking into account real wind tunnel data.
I realize that spallart-allmaras model is not the best one, due to it was designed for transonic speeds,
but i think that i using K-W model is best result with modify the turbulence intensity parameter with magnitude less than 0.1%..but i fell that the result still
not acurate enough for cd (drag coefficient). i have done the grid mesh with more number of cell..but still not accurate..
i need some suggestion what my mistake occure and what should i do to get the better result for cd that approach the experimental data..i try simulate the naca 4415 type
please help me..:)
what solver are you using for your simulations ?
Also do you know what was your wind tunnel turbulence intensity ?
i am using pressure based solver..i follow the turbulence intensity based of the velocity...that for subsonic category turbulence intensity must less than 0.1 %...i am validated the simulation result with experimantal book in theory of wind section, by Ira and Abbot..so i don't know the value of turbulent intensity in the case...what parameters must i change...i am once change the parameter area..that the default value is 1 but i was change with more than 1. so the result is better..but still not enough accurate with experimental data in that book. what should i do ? please help me..give me more information...thaks alot of
If you are using Fluent as I suspect you won't be able to obtain the correct Cd simply because Fluent doesn't take into account the transition from laminar to turbulent ( maybe if you will use Fluent 12 you will be able to use the new model of Menter ).
All the Abbott cases were done at Mach around 0.2 so you will have large zone of laminar flows on your airfoil.
You won't be able to obtain the experimental Cd unless you take into account the transition region.
However you should be able to obtain the experimental Cl with 10-15% error if you set up correctly your case.
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