X-Foil help and Reynolds number help

[PNSD]

Hello I have just got myself xfoil and having fun.
Though now i thought it would be a good idea to use it for a uni project.
Thing is I can plot pressure distribution for different alphas for one airfoil. I can also plot on the same graph the multiple alphas so i can compare.
however I can not find a way to plot a simple lift curve graph. ie Cl against Alpha???

Any help would be good thanks.

[Ciro Pabón]

I'm not sure if this program is what you're looking for:

https://engineering.purdue.org/AAE/Acad ... m_5/Coeffs

[PNSD]

I cant access the webpage. Something about security?

[tahadar]

have you tried the user manual? In my current uni project two of my groupmates were doing x-foil analysis for wind turbine blades and every time they got stuck they seemed to find the solution in the manual.

[Ciro Pabón]

Well, my browser (Opera) says something like "The server certificate does not match the server's name. Accept?". When I say "Yes", I can see it.

Just in case, I copied the program here: http://cirospictures.googlepages.com/Coeffs.htm

I browsed the manual and tried to read the program (in my mind ), but I haven't downloaded the software, so I don't know. It's for a particular profile and a certain set of alpha angles, it's also full of parameters (Mach number, Reynolds, aspect ratio, speed of sound, span, fuselage width, you name them) so I guess you have to change them according to your needs. Please, take in account I have no idea of aerodynamics, I'm into roads...

Tahadar is totally right. There must be a forum somewhere, I'd add.

[PNSD]

Thanks but xfoil is not recognizing those commands :s I have 6.96 and thats for 6.94 so could be why.

Cant see anything in the manual either... though when i plot the pressure dist it does show Cl results, so i will simply record the results of cl against alpha and then do my own graphs.

whilst i have a topic open there is another question.

this is for a model glider wing.

to get the accurate lift curve i needed to work out the reynolds number to input into xfoil.

re = rho*v*l / u

edit - the glider wil be tested in a sports hall so i assume the u is standard air at standard room temp

rho = 1.225
v = 4m/s ( the speed the glider will be launched at )
l = chord length, atm it is 0.12metres

and for u i have a value of 1.78*10^-5 givng a reynolds number of 330033.71

I input the data into the NASA applet on their site and I got 33933 Re.

Is this the correct way of calculating it and therefore my lift curve? because even right up to alpha=20 i had no drop in Cl.

[Gecko]

By default xfoil does the inviscid calculation, which will not show separation. You need to turn on the viscous calculation, I believe Visc is the command. Then to get the plots of lift vs alpha, you need to switch the mode invoking the Seqp command. Then, you need to do Pacc command so that the data will be written to a file. Finally, do a sweep over alpha with the ASeq command. All of this is described in the manual,

http://web.mit.edu/drela/Public/web/xfoil/xfoil_doc.txt

Cheers and good luck!

[PNSD]

I'll have to go through more carefully then, thankyou !