Calculate Half-dimond Airfoil CL, CD, CM, Drag & Lift

[KEYOFDARK]

Hello Guys,

This Is My First Day ,first Topic,first Question In Here And I Would Like Some Body To Help Me As Soon As Possible,i Dont Know If I Am In The Right Place Or Not But Let Me Give It A Try...

My Question Is Related To Half-dimond Airfoil,i Would Like To Know How To Calculate The Cl,cd,cm,drag And Lift, In Detail,pleasezzzzz,

I Will Be Very Thankfull...let Me Here From U Soon,

Keyofdark

 

[mab623]

Hello,

Im not sure if you still need this result but I will post it anyway. This is done in detail in the book by anderson "compressible flow with historical perspective".

What is your incoming mach number? Is the incoming flow at Standard Temperature and Pressure?

Remember that Lift can be calculated by the closed surface integral of the pressure time with your unit normal vector dotted with the vertical unit vector times the infinitesimal area. Remember that even if your fluid is inviscid you have form drag or pressure drag to deal with. This calculation is the same as lift however you must dot the outward normal vector with the horizontal unit vector to resolve the forces in their proper direction.

Rememeber that once you have lift Cl= L/qC were q is the dynamic pressure and c is the cord length of your airfoil. Cd is the same however you use lift instead of drag. Remember that you use free stream values to calculate q.

I suggest you really review your volume and area integral skills as they are so very important for Aersopace Engineers. Hope I was of some help.

 

[mab623]

Also, you must remember that you will not generate any lift without an angle of attack of your airfoil is symetric!