Joshrk22
Sierra Hotel
- Joined
- Feb 26, 2006
- Posts
- 230
We really need a technical forum on this board. Anyways, I'm doing a physics (high school) project relating AoA and lift. Is there a formula to convert the coefficient of lift to angle of attack? I have found the Max CL for the airfoil but I want to convert that to angle of attack so I know the critical AoA.
I know for thin airfoils you can use AoA=2*pi*radians but I am using NACA 2412 (Cessna 172/182 airfoil) so the chord % is 12% with camber % being 2%. How can I adjust the equation to account for these factors?
Next, for the calculation of max CL I have:
CLmax=(2*L)/(S*rho*V^2)
L = 2,100 lbs/2.2=954.5*9.8 = 9354.5 N
S = (174 ft^2)/(3.28 ft)^2 = 16.17 m^2
rho = 1.225 kg/m^3
V = 48 KIAS/1.15 = 41.7 mi/h *(1609/3600) = 18.7 m/s
CLmax = (2*9354.5)/(16.17*1.225*18.7^2)
CLmax = 2.71
This seems kinda high for the CLmax on a 172. Is this right? Please help all you aeronautical engineers!
I know for thin airfoils you can use AoA=2*pi*radians but I am using NACA 2412 (Cessna 172/182 airfoil) so the chord % is 12% with camber % being 2%. How can I adjust the equation to account for these factors?
Next, for the calculation of max CL I have:
CLmax=(2*L)/(S*rho*V^2)
L = 2,100 lbs/2.2=954.5*9.8 = 9354.5 N
S = (174 ft^2)/(3.28 ft)^2 = 16.17 m^2
rho = 1.225 kg/m^3
V = 48 KIAS/1.15 = 41.7 mi/h *(1609/3600) = 18.7 m/s
CLmax = (2*9354.5)/(16.17*1.225*18.7^2)
CLmax = 2.71
This seems kinda high for the CLmax on a 172. Is this right? Please help all you aeronautical engineers!
