L
Liras
Hi,
I'm trying to choose airfoils and establish wing dimensions (and high lift devices if needed) for my design. I have a low stall speed requirement (Vs0 = 42km/h [with flaps]), so I'm looking into high-lift airfoils with good stall and thick enough for a light spar. I'm trying to balance CLmax vs wing area. It's a single seater UL design with no more than 180kg gross weight.
I have a problem though. It's hard for me to estimate the actual, practical CLmax. My Reynolds Number is 1 milion. AirfoilTools.com says USA35A would give me some CLmax = 1.8, airfoildb.com, on the other hand, goes over 2. My wings will probably be of rectangular planform and AR~7, so I conservatively multiply the CLmax by 0.9. So, I've got: 1.62 and 1.8. That's quite a difference. My choice of high lift devices and wing area depends quite heavily on the airfoil's CLmax and 0.2 is a lot.
So, a question to the more experienced than me, which prediction is closer to the truth? Am I right to multiply by .9?
P.S.: At cruise Cl of 0.25 USA35A generates quite a bit of drag (Cd = 0.012-13).. I won't be going fast, some 120km/h but still.. maybe you have some better suggestions for airfoils for this kind of plane? Maybe a combo of 2? Just remember the stall speed req...
Thanks,
Liras
I'm trying to choose airfoils and establish wing dimensions (and high lift devices if needed) for my design. I have a low stall speed requirement (Vs0 = 42km/h [with flaps]), so I'm looking into high-lift airfoils with good stall and thick enough for a light spar. I'm trying to balance CLmax vs wing area. It's a single seater UL design with no more than 180kg gross weight.
I have a problem though. It's hard for me to estimate the actual, practical CLmax. My Reynolds Number is 1 milion. AirfoilTools.com says USA35A would give me some CLmax = 1.8, airfoildb.com, on the other hand, goes over 2. My wings will probably be of rectangular planform and AR~7, so I conservatively multiply the CLmax by 0.9. So, I've got: 1.62 and 1.8. That's quite a difference. My choice of high lift devices and wing area depends quite heavily on the airfoil's CLmax and 0.2 is a lot.
So, a question to the more experienced than me, which prediction is closer to the truth? Am I right to multiply by .9?
P.S.: At cruise Cl of 0.25 USA35A generates quite a bit of drag (Cd = 0.012-13).. I won't be going fast, some 120km/h but still.. maybe you have some better suggestions for airfoils for this kind of plane? Maybe a combo of 2? Just remember the stall speed req...
Thanks,
Liras
