# Equations - Nacra curve

Hi,

I am trying to use the custom option in DSM to create an Airfoil section based on the NACRA profile, currently with no luck ( lack of experience and knowlege hopefully).  The actual equation from Wikipedia is:

yt = 5t[0.2969Sqrtx - 0.1260x - 0.3516x2 + 0.2843x3 - 0.1015x4]

Nacra Airfoil @ Wiki

My attemp in the equation tool is:

5[t][(0.2969*Sqrt[t])-(0.1260*[t])-(0.3516*Pow([t],2))+(0.2843*Pow([t],3))-(0.1015*Pow([t],4))]

All I get is a straight line :(

Idealy I would want the airfoil section to be 300mm wide from which i can make a surface.

Many thanks

Hi,

I added a couple of round brackets where the power function (Pow) is applied to the chord position and defined a separate parameter for the thickness to chord length ratio [a]. This example shows a NACA 0015 (a=0.15, t= 0->1, scale=50 - adjust to requirement) symmetric profile:-

For upper airfoil surface:
y= 5*([a])*((0.2969*Sqrt([t]))-(0.1260*[t])-(0.3516*Pow(([t]),2.0))+(0.2843*Pow(([t]),3.0))-(0.1015*Pow(([t]),4.0)))

x= ([t])

For lower airfoil surface (sign of y is negative):
y= -5*([a])*((0.2969*Sqrt([t]))-(0.1260*[t])-(0.3516*Pow(([t]),2.0))+(0.2843*Pow(([t]),3.0))-(0.1015*Pow(([t]),4.0)))

x= ([t])

Click the green tick to finish each curve separately.

Check the screenshot if the formatting of my comment gets messed up. At the trailing edge (t=1), the airfoil isn't zero thickness using the default 4th coefficient. Change it from 0.1015 to 0.1036 for zero thickness so both upper and lower surfaces converge.

I haven't tried equations for cambered profiles yet; this is what most real world aircraft use. Example: NACA 2412 used in Airbus A380s. The first two digits of NACA designation indicate the airfoil has an asymmetric profile (with camber).

Regards,
JC