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October 25, 2015, 16:59 
Airfoil 2D cfd simulation

#1 
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Filippo
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Hi everyone,
I'm working on an exam project consisting on find aerodinamics coefficients and pressure trend around a 2D airfoil using gambit and openfoam. almost every parameter is given by the professor and I should use the openfoam 2d incompressible airfoil tutorial as a base for the simulation, so the only thing I can really tweak is the mesh in gambit.... my airfoil is a NACA 2415 (chord=4,875m, 2°angle) and other parameters are: Re= 6 x 10^6 freestreamvelocity=18 m/s cinematic viscosity= 1.5 x 10^5 nut= 1.2 and nuTilda= 2.2 (both calculated with Spalart equations) the turbulent model is SpalartAlmaras and wall function is nutUSpaldingWallFunction talking about the mesh, I had to put the airfoil in a domain sizing (given values) 5xchord from airfoil on front,top and bottom sides, while 10xchord on the backside I used a boundary layer around the airfoil and the wake and triangle pavè outside it...... talking about first cell size, we were asked to keep y+ in a range from 30 to 300, that correspond to a first cell height from 1mm to 10mm.... so i cretated the boundary layer with a 2mm first size and a 1.1 grow factor then i extracted the mesh and put it on openfoam, it extruded the 2d mesh with a 2m cell in z direction.... after this I changed all the tutorial parameters with my initial conditions, just replacing values I started the simulation but here are the problems: 1)aerodinamic coefficients differ from tabulate ( Theory of Wing Sections) values, Cl seems preety accurate (0.018) but Cd is extremely inaccurate, should be +0.01 and I get 0.045.....why this? 2)talking about the referement Area to calculate the coefficients, I put Aref = 2 x chord cause when I try to chekMesh, it gives that domain in z goes from 1 to +1.......am I doing this right? 3)my simulation doesn't reach convergence, I have a tollerance of E^5 but the minimum I could reach with initial residuals is E^3, why this? I think that my problems lie on the mesh (given the fact that al conditions have been provided by the professor)......so I would like to know some tips such as: what is a good interval size for the airfoil meshing? what is a good far field cell size? and what a good growing factor? thanks Last edited by Cocorito90; October 25, 2015 at 18:00. 

October 25, 2015, 17:15 

#2  
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Filippo
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here is my mesh in gambit:
http://imgur.com/a/AkyMY here the checkMeshlog: Quote:
Quote:
Quote:
Last edited by Cocorito90; October 25, 2015 at 19:36. 

October 25, 2015, 19:41 

#3  
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Filippo
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here initial conditions:
Quote:
Quote:
Quote:
Quote:


October 26, 2015, 05:26 

#4 
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Filippo
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I solved the drag problem by myself , I put wrong formulas for nut e nuTilda, now I get much more correct drag coefficient...
anyway, I still have 2 problems: 1) with new nut e nuTilda I now get a lift coefficient that is still close to zero, but positive....comparing with the data graph I can't understand if it should be positive or negative (2°), so could this be still corrrect? 2) as I've previously stated, after 500 iterations my residuals don't reach convergence, even if coefficients don't change by much in the last part of the simulation, could this be normal? my residuals plot in 500 iterations: http://imgur.com/0SYY9KY Last edited by Cocorito90; October 26, 2015 at 06:44. 

October 27, 2015, 05:33 

#5 
Senior Member

Hi,
500 iterations is not a lot, just continue running for more iterations, the plot looks ok up to this point. Just try and see what happens. It may be that they will flatten, or they will continue to get to a very low level. A couple of 1000 iterations is nothing unusual. Regards, Tom 

October 27, 2015, 08:11 

#6 
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Filippo
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ok now I have another problem, I have found the graphs for NACA 2415 on the website http://airfoiltools.com/ and it says that at 2 degrees I should get Cl=0.0256 (this is different from the graph I've shown you before, don't know why)........anyway, now I've made a rougher mesh and it converges in about 700 iterations, bu there' s a weird thing happening:
if I widen the mesh in the boundary layer on the trail (in x direction) Cl grows susbstantially, and with an interval size of about 200, I can get Cl=0.0255.... but this seems like a trick for me, why is this happening? what phisical explanation for this? 

October 27, 2015, 09:04 

#7  
Senior Member

Hi,
I do not know what you mean by this: I think you mean coarser resolution in chord direction? Or this: Quote:
If you are changing the resolution in chord direction I think you may be altering the separation point (if any) near the trailing edge, or modifying the discrete geometry with respect to the actual geometry (maybe even changing the camber line). I guess pictures would help a lot... 

October 27, 2015, 09:35 

#8 
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Filippo
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this is my mesh and what I mean for 200 interval size: http://imgur.com/a/pZtN9
my chord is 4,875m (4875 in gambit size) increasing that 200 to (let's say) 250 make the Cl pass from 0.0255 to 0.0285....and reducing it to (let's say) 100 make the Cl go to 0.014 or something like that.... why this? 

October 27, 2015, 10:19 

#9 
Senior Member

I would suggest to look at the pressure distribution on the airfoil as a function of your interval size. It seems like the pressure distribution is very sensitive to this and it should not be like that. You need to find the resolution you need where a further refinement does not significantly change the result. This is called a mesh sensitivity study, which you need to perform to find a grid independent solution. If the terms in bold are not clear, please look for some more literature on them. The area where you are changing the resolution is the pressure recovery area in the wake of your airfoil. It is very important in the pressure distribution on the airfoil and by that the lift and drag. This is one of the difficulties of CFD I would say.
Regards, Tom 

October 27, 2015, 12:03 

#10 
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Filippo
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I found that refining the mesh (interval size) on the boundary layer trail stabilize the Cl around 0.0034, but according to this graph I should get small and negative number....


October 27, 2015, 12:30 

#11 
Senior Member

Well small and negative or small and positive is just a small absolute error. It may well be within the experimental uncertainty (maybe they actually measured at 2.1 degree) so I would say that your result is rather good. Best option would be to take a couple of different angles of attack and try to match the slope of the lift curve. I would never expect to be spoton to a single point of an experiment and around zero small absolute errors are big relative errors, but I would not be too worried about that, since there are always small discrepancies that are hard to account for. Also try to compare the pressure coefficient on the airfoil surfaces with the experiment. If they are similar I think you would have a good simulation result. In my opinion it makes more sense to compare this than a single point on a graph. Since the pressure distribution should tell you if you have (approximately) the correct value because you solved correctly the flow around the airfoil or if you happen to find two errors that cancel each other.


October 27, 2015, 15:04 

#12 
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Filippo
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I'm really a retarded, I could've tried before with another angle......tried +2°, it should be 0.4, mine is 0.42


December 8, 2016, 03:21 
Naca2415

#13 
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H salem
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Hello Dear Foamers,
I am trying to simulate the NACA2415 at 0 angle of attack. I checked my boundary conditions and Mesh several time and the solution will converge, but I can not take the right Cl, Cd from post processing. I am using the openfoam recommendation for calculating the Cl from its user guide but the Cl will be two times the real one. It should be 0.2 but my results shows the 0.4. what should I do????? I am really exhausted. the whole my setup and BC's are attached here. please let me know the right answer. 

December 12, 2016, 17:23 

#14  
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Joshua
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Quote:
Hope this helps! Joshua 

December 19, 2016, 03:43 
polt cl ,cd

#15 
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hi.how to draw (cl,cd alpha )plot for airfoil by fluent?


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