[giddo244]

Hi all, for a school project I'm looking at the lift against angle of attack for a 3D NACA 2414 wing at about a Reynold's number of 200,000. However, I am getting results that are about 30% lower than expected compared to an experiment and 2 other CFD studies I found which more or less agree with each other. My apologies for a wall of text but I would greatly appreciate it anyone's help to identify places that may be causing this error. I have attached an imgur link for my screenshots and images: https://imgur.com/a/TL8pADi

1. Mesh
Firstly, my wing has a chord of 140mm and a span of 240mm. I am using a C-domain 40 times the chord. For inflation layers, I've set first layer height to be 0.02786mm, 15 maximum layers to get an eventual y+ = 1 (I have been using k-w SST). My edge and face sizing around the wing is 1.4mm and I have a body of influence with 5mm element size.
2. Fluent
Pressure-based, steady. I have been using the k-w SST model. However, trying different models (k-e realiable and SA) did not really improve my results. I have opted to use a constant value for air density and viscosity. For my boundary conditions, I will use my 10 degree angle of attack case. My reference values are taken from the inlet. For methods, I used second order upwind. For report definitions I used the vector compotents. And lastly, I intialise from the inlet.
3. Results
I get the following results. They are lower than expected vs experimental data at the same Reynold's number and lower than other CFD studies I found.

I really would appreciate any input!

[LoGaL]

Interestingly, you are catching the correct Cl and Cd at 0 AoA. You sure you are not messing up your trigonometry calculations for Vx and Vy?

[giddo244]

Oh my I hope its as simple as that, but I don't think this was the mistake I made.
Here are the trig numbers I used: https://imgur.com/0ZpRLom