scanflyer

Hi,

I do not know if this is the right forum to ask this question, but I will give it a try.

I am trying to find a NACA-aerofoil for a given wing geometry (AR = 8, Taper: 0.35, Cruise CL = 0.05 at 0.6 degrees angle). Do any of you know if it is possible to find an aerofoil that will give this wing the best L/D-ratio at this cruise AoA?
I have done some analysis on several aerofoils on my wing using the XFLR5-software, but the best L/D-drag is usually achieved at an angle of attack of 3-4 degrees.

Or maybe the only solution would be to change the wing geometry? The cruise lift coefficient of 0.05 is very small, but due to some geometrical problems this can maximum be set to 0.1, giving an AR of 5. But even at this CL, I struggle to find the aerofoil giving the best L/D-ratio, or at least very close to it.

Hope someone can help me out with this.

Himat

You specify: Cruise CL = 0.05 at 0.6 degrees angle of attack.
A question here, why do you specify both Cl and angle of attack?
With a given weight, wing area and wing plan form you fix the wing Cl by choosing the cruise speed. You then need to find an airfoil that have low drag at this Cl. The angle of attack is then a free variable and you get one depending on the airfoil selected. The angle of attack will probably be small, but is there a reason for it to be exactly 0.6 degrees?

scanflyer

Thanks for your reply. I first found a weight and wing area, then I found the CL required for the area and weight. To find the AoA I used the lifting line theory; (2*pi*area*alpha)/(area+2) = CL. This formula is just an approximation about what angle I can expect to cruise at, but I assume this gives a reasonable result? And due to this low angle, the L/D is quite low, at least from the results I get from the software.

Himat

I am not that well versed in lifting line theory, it gives an approximation of what angle of attack to expect, but I do not know if it cater for how airfoil camber and thickness distribution modify the lift vs. angle of attack. The expression (2*pi*area*alpha)/(area+2) = CL at least have no term that account for airfoil camber.

Shoe

You've probably moved on, but the lift curve slope of an elliptical lift distribution predicted by lifting line theory is:

2*pi*Aspect_Ratio/(2+Aspect_Ratio)

You will not get sensible results if you use Area in place of Aspect_Ratio. One way you can tell there's something fishy with the "area formula" in the posts above is that you will get different answers if you measure the area in square inches instead of square feet. That's a sure sign that something is amiss.