XFOIL: number of digits in output

jack1980 · January 18, 2013, 02:50

Hi, is it possible to increase the number of digits in the polar output values of CL, CD, and CM? I think I did this once, but really can't remember how ...

Thanks for your help!

truffaldino · January 19, 2013, 12:50

Increasing number of digits in x-foil does not make sense, since, even with standard number of digits, errors related with methods applied exceed truncation errors by orders of magnitude.

jack1980 · January 19, 2013, 14:19

Thank you very much for your reply. I guess I should have mentioned that I'm asking this because I would to see if I can increase the smoothness of sensitivities, calculated from finite differences.

luigi22 · October 11, 2016, 05:10

Hi,
I have to calculate Cl Cd for different type of airfoils to vary the angle of attack (alpha), I do in this way:
naca0018
pane
oper
visc 100000
iter 1000
alfa 0
pacc naca0018.pol
aseq -30.0 30.0 2.0
pacc
quit
The results are different from the results of wind tunnel, maybe cause the large variation of alpha. what do you think?

January 18, 2013, 02:50	XFOIL: number of digits in output	#1
jack1980 Senior Member Jouke de Baar Join Date: Oct 2009 Posts: 127 Rep Power: 16	Hi, is it possible to increase the number of digits in the polar output values of CL, CD, and CM? I think I did this once, but really can't remember how ... Thanks for your help!

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January 19, 2013, 12:50		#2
truffaldino Senior Member Join Date: Jan 2011 Posts: 249 Blog Entries: 5 Rep Power: 17	Increasing number of digits in x-foil does not make sense, since, even with standard number of digits, errors related with methods applied exceed truncation errors by orders of magnitude.

January 19, 2013, 14:19		#3
jack1980 Senior Member Jouke de Baar Join Date: Oct 2009 Posts: 127 Rep Power: 16	Thank you very much for your reply. I guess I should have mentioned that I'm asking this because I would to see if I can increase the smoothness of sensitivities, calculated from finite differences.

October 11, 2016, 05:10		#4
luigi22 New Member luigi Join Date: Oct 2016 Posts: 6 Rep Power: 9	Hi, I have to calculate Cl Cd for different type of airfoils to vary the angle of attack (alpha), I do in this way: naca0018 pane oper visc 100000 iter 1000 alfa 0 pacc naca0018.pol aseq -30.0 30.0 2.0 pacc quit The results are different from the results of wind tunnel, maybe cause the large variation of alpha. what do you think?