First of all I'd like to ask if anyone is aware of a reliable method to calculate the wing Max CL with regards to aspect ratio. I know of several methods to calculate the lift curve slope but this will not tell me when the wing will actually stall and what Cl i get at that point.
I have also gathered data for a few aircraft of similar size to what I am designing in order to compare them and to be able to calculate what CL max they are able to get from their wing. The results I get seems to be way of.
Corby Kestrel
Wing area 6,55m2
Gross weight 340kg
Stall speed 39mph
Solving for CL from the lift equation gives CL=L/qS. I get a CL max of 2.73!!
Midget Mustang
Wing area 6,32m2
Gross weight 431kg
Stall Speed 60mph Clean
I make the CL max 1.52
Sonerai I
Wing area 6.97m2
Gross weight 318kg
Stall speed 45mps
I make the CL max 1.8
Sonex Onex
Wing are 7.84m2
Gross weight 431kg
Stall speed 50mps clean
I make the CL max 1.76
What am I doing wrong? These CL values look very incorrect to me?
The onex for example is using the NACA 64-415 airfoil. The smooth section data for that gives a Cl max of just over 1.6. With the rule of thumb CL max=cl max*0.9 it would be 1,44 but I even find this a high number given the fact that the aspect ratio of the onex is only about 4.7. Either way 1.44 is a long way of from 1.76.
Thanks for your help!
I have also gathered data for a few aircraft of similar size to what I am designing in order to compare them and to be able to calculate what CL max they are able to get from their wing. The results I get seems to be way of.
Corby Kestrel
Wing area 6,55m2
Gross weight 340kg
Stall speed 39mph
Solving for CL from the lift equation gives CL=L/qS. I get a CL max of 2.73!!
Midget Mustang
Wing area 6,32m2
Gross weight 431kg
Stall Speed 60mph Clean
I make the CL max 1.52
Sonerai I
Wing area 6.97m2
Gross weight 318kg
Stall speed 45mps
I make the CL max 1.8
Sonex Onex
Wing are 7.84m2
Gross weight 431kg
Stall speed 50mps clean
I make the CL max 1.76
What am I doing wrong? These CL values look very incorrect to me?
The onex for example is using the NACA 64-415 airfoil. The smooth section data for that gives a Cl max of just over 1.6. With the rule of thumb CL max=cl max*0.9 it would be 1,44 but I even find this a high number given the fact that the aspect ratio of the onex is only about 4.7. Either way 1.44 is a long way of from 1.76.
Thanks for your help!
