# meanings of the numbers in the airfoil NACA 65(12)-10, where to find and plot

I am trying to plot an airfoil named NACA 65(12)-10 in javafoil. I found it in a few research articles. However, I am not sure from what series the airfoil is from. I tried looking it up on Airfoiltools website but to no avail. I am unsure if it is from the 6-digit series because it doesn't follow the naming nomenclature, such as NACA 65-210. Can anyone help me understand the name and how I can plot it? thanks.

• There must be an error in the airfoil number. This one would have a width of the laminar bucket of 2.4 c$_L$, way too wide for any laminar airfoil. Also, for a proper 6-series name there need to be 3 digits past the hyphen, not two. Please explain where you found this odd number! Commented Nov 26, 2022 at 19:22
• I found a research article link using the airfoil and a few more research articles relating to the topic link Commented Nov 27, 2022 at 5:11

## 2 Answers

According to this NACA report:

"In the designation NACA 65-(12)10, the first two digits, 65, refer to a particular thickness distribution; the second two digits, 12, indicate the isolated airfoil lift coefficient in tenths; and the third pair of digits, 10, denotes the maximum thickness in percent chord."

It seems that that airfoil is especially used for gas turbine compressor blades.

It looks like they're stretching the NACA definition -- this is fairly common, but it would be nice if they said exactly what they're doing in the paper or cited the paper that did.

Theory of Wing Sections by Abbott and von Doenhoff is your go-to reference for this kind of information.

Normally, a 65 series is numbered 65 XYY. The X is the design lift coefficient in tenths and YY is the t/c in percent. so, a 65310 would have a design cl=0.3 and t/c of 10%.

In this case, it looks like your airfoil has an extreme amount of camber such that the design cl=1.2 and it has a t/c of 10%.

I quickly generated something this in OpenVSP and it looks about right. Right now, OpenVSP will only go up to cl=1.0. I modified my local copy and generated this airfoil with cl=1.2 and t/c=10%.

I searched back through a few of the references that use this airfoil and this data set. It appears that they've thickened the trailing edge and rounded the absolute end. The theoretical airfoil in the above image has an extremely sharp cusped TE that would be impossible to manufacture. Digging back further, the airfoil used becomes even less clear.