Hi

These parameters have been determined for each airfoil.

You can see the files that you put into them.

Hi Shahrukh.

Sajjad is right, but I doubt if you would be using a predefined airfoil if your objective is to optimize the airfoil itself.

In order to calculate the camber percent chord and thickness percent chord, I suggest you follow these steps:

You will have to do these calculations for all configurations of your airfoil.

Hope this helps.

Hi Sajjad,

Thank you for providing the data file of design parameters of airfoil. It is very helpful, but as pointed out by Saumitra, my objective is to optimize the design of airfoil by varying the parameters.  

Hi Saumitra,

Thank You for your guidance. It is very useful information that I was searching for. Now I can easily optimize the design of airfoil. 

The method explain by mr joshi is an appropriate way but its incomplete without the airfoil equation, to calculate camber and thickness percentage i suggest you to read basic text books of aerodynamics in which the experiment results of several airfoils available in which you found the equation of airfoil with camber and without camber and through that equation you can generate points for a particular type of profile.

please follow this page where you can find the equation of both kind of airfoil

http://en.wikipedia.org/wiki/NACA_airfoil

as well to generate data directly can do it over here

NACA 4 digits

http://airfoiltools.com/airfoil/naca4digit

NACA 5 digits

http://airfoiltools.com/airfoil/naca5digit

Regards,

Shrish Shukla

Hi Shrish Sir,

Thank You for your help. I am currently using a java application named "JavaFoil" developed by Dr. Martin Hepperle. This applet directly gives me the data points for different airfoil series namely NACA 4-digit, 5-digit, 6-digit and NACA 07-series and 16-series. It has a user interface through which I can modify the camber and thickness of the airfoil, and  the required data points of the modified airfoil design is obtained which can be easily imported in ANSYS. 

Hai Shahrukh,

CUrrently I'm doing modification same as you. I'm trying to change the t/c location (x/c) without changing the t/c itself. Do you have any idea on how to do it? I'm using JAVAFOIL too to import the coordinate into ANSYS. I've got all the details of my airfoil there and measured the t/c at x/c in solidwork. It agrees well with javafoil. However, I couldnt change the t/c without changing the location. can you give any advise on doing so?

Hi Nur,

If you are using airfoil software JAVAFOIL, then you can easily change the thickness location(xt/c) without changing the thickness (t/c). Just check the attached image. I've highlighted the option (xt/c) to change it without changing the thickness (t/c).

Regards,

Shahrukh  

thanks!! i found the solution weeks ago. i did notice t/c change a little bit, buat only at the point value which not a significant change. anyhow, thanks for answering. by the way, since you're studying about an airfoil, did you calculate the value of Cd? since the equation containing Cd is a drag force equation and the equation consists Area parameter, how do you calculate the area of an airfoil? especially in my case, it's definitely a 3D airfoil with span. unfortunately, my leading edge span is a little shorter than my trailing edge span. any idea?