# How can I calculate airfoil thickness?

Given this information... ... how can I find the thickness for a "NASA/Langley LS(1)-0417MOD general aviation airfoil" at 90% and 99% chord?

You have already done half the work. With the precise name it was easy to find the coordinates here and plot them. Yes, the coordinates look fine (no steps, no distortions) Now either measure the thickness directly from the plot or subtract the lower contour point at 90% and 99% from its corresponding value on the upper side. You will need to interpolate between the two adjacent points, though.

Since you seem to need the thickness to calculate the lift curve gradient, you might just rely on the NASA technical paper which contains measurements and airfoil data on just this one airfoil. Be sure to correct for the proper Reynolds number and check whether your application uses the same thick trailing edge that is used in the coordinate file.

• Thankyou verymuch sir, actually. what should i do to calculate thickness at Y90% chord and Y99% chord. how to calculate it? – Ade Firmansyah Oct 18 '16 at 19:13
• how do i calculate with theorytical – Ade Firmansyah Oct 18 '16 at 19:23
• because i was trying search used CAD to find value thickness at Y90% chord and Y99% chord. and i have the value, then i try equation at the picture. but the value of cl alpha be divided cl alpha theory more small from NACA 0015. – Ade Firmansyah Oct 18 '16 at 19:24
• i mean how do i calculate use theoretical method to find value, thickness at Y90% and 99%. – Ade Firmansyah Oct 18 '16 at 19:27
• @AdeFirmansyah: This depends on the airfoil. Most airfoils are shaped to yield a desired pressure distribution and the above mentioned method is the only way - there is no "theoretical method". This is also true for the LS(1)-0417. Old 4-digit NACA airfoils use a fifth-order equation for the thickness distribution along the chord, and there is an analytical method, but using the technique described above is easier. – Peter Kämpf Oct 19 '16 at 4:51