[spaceman776]

Hi, I am doing a CFD analysis of transonic (M=0.8) flow over a NACA-0012 airfoil at an angle of attack 2º as practice. I followed a Cornell tutorial on creating the geometry and mesh exactly but that was for subsonic flow of V=1. So I used the same mesh and geometry but I made changes in Fluent to account for compressibility and such. I used a C-Mesh that extends 12.5m on both sides of the origin (located on the tip of the trailing edge) with the airfoil length being 1m.

My setup:
General > Density-based solver and steady
Models > Energy On and Inviscid
Materials > Air > Density > Ideal Gas
Boundary Conditions > Inlet and Outlet are set to Pressure Far-Field with gauge pressure set to 0, M=0.8, and the x and y components set to the cosine and sine of 2º respectively.

Using the setup above I get the following results. It looks 99% correct but the bottom surface won't show any shocks for some reason. Here is what it should look like at 1.25º. I set it up for 1.25º to compare it with the prior link and the bottom surface still doesn't match correctly. I am not sure if I made an error when changing my setup from subsonic to supersonic. Anyone know what could be the cause of this? Thanks in advance!