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Solved: Pressure (Lift) Distribution along airfoil chord (and software)

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Brad Bock

Member
Joined
Nov 2, 2015
Messages
13
Location
Defiance, Ohio
A while back I needed to determine the pressure distribution along an airfoil chord for various airfoils at various angles of attack. I looked everywhere, including here without any real success, so I decided to look deeper and try to figure it out myself. It took me a while to figure out how this is calculated and I decided my best option was to make an Excel program that would run it. The whole process took about a month and since I ran across quite a few people looking for the same answer I decided to share the software I created here. The software seems to be pretty accurate but at high angles of attack the CL results start to slightly vary with the results from Xfoil. I left all the pages and calculations open in case anyone with knowledge cares to take a look at how the values are calculated and find any mistakes. Just as well, it works pretty well for people who want to know the pressure distribution but don't know how to find it.

You can download the software here. I call it CPcalc...... https://www.dropbox.com/s/33or71wd88w4niq/CPcalc.xlsm?dl=0

You must have Xfoil v6.99 to use this software. I will give the tutorial below for how to run an airfoil through both programs

-Open Xfoil
-Select a 4 or 5 digit naca airfoil or you can import your own if you have it evenly divided between top and bottom into 331 coordinates (the middle being zero) as xfoil cannot do this on an imported airfoil. Only on a Naca 4 or 5 digit airfoil. To select a naca 4 or 5 digit airfoil, simply type it into Xfoil and press enter. For example, for a naca 0014, simply type naca 0014 and press enter
-Type pcop and press enter. this will evenly split the airfoil between top and bottom. CPcalc needs this to work
-Type oper and press enter
-Type visc and press enter
-Type in a reynolds number and press enter. An easy way to determine reynolds number is airfoil chord in meters times airspeed in meters per second times 70,000. ( c(m)*V(m/s)*70000=Re)
-Determine whether you want to calculate by angle of attack or CL. for angle of attack type a and then the angle, then press enter. For example "a 6". For a CL of say 1.12, type "c 1.12" and press enter.
-To export the raw CP data type cpwr and press enter, then choose a file name followed by .txt . For example I usually name the file CP.txt ..... when it aks you for the file name type it in and press enter, being sure to follow it with .txt

-Open CPcalc and be sure to enable macros. It will not work unless they are enabled. In the "Input Airfoil Data" tab you will see a button at the top right titled "Load New Airfoil". Click this button and select the txt file you just saved in Xfoil. (It will be in the Xfoil folder unless for some reason you moved it).
-Click yes you want to replace data when the popups ask you.
-The rest is pretty self explanatory.

I hope it proves useful to you and if anyone has any corrections to the calculations I welcome them as I taught myself how to do this
 

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