[ChaldExals]

Hi Everyone,

I ran a CFD analysis on NACA 0006 wing at 0 angle of attack. The root length,tip length and span of the wing is 10cm,8cm,10cm respectively.The freestream velocity is 300 m/s and pressure is 1 atm (sea level conditions). When the analysis is completed the drag acting on the wing is 5N which corresponds to a drag coefficient of 0,005. I am new to CFD but this number seemed so small. Can this be correct? For further information: in the mesh part, I have a minimum orthogonal quality of 0.0138,maximum skewness of 0.95 and maximum aspect ratio of 792 (I checked orthogonal quality and aspect ratio from Fluent). The first layer thickness is 2.5e-6m. I have 15 number of layers with a growth rate of 1.2. In the setup part I converted the mesh to polyhedral mesh. I used density-based solver,ideal gas model and SST k-w turbulence model for air. I also share the mesh and mach number contour I get. Thank you in advance.

[LuckyTran]

I'm trying a bunch of different combinations like 5*2/300^2/.01^2 ~ 1, how do you get 0.0005?

[ChaldExals]

I obtained the surface area from 3D CAD software. So my calculation is : (5*2)/(1.2*0.01838257*300^2) ~ 0.005

[LuckyTran]

Your projected frontal area should be (10 cm * 6%*10 cm) and your planform area (10 cm x 10 cm). Your reference area is going to be one of those or close to one of those and you need to use the drag coefficient corresponding to either definition. If you use the planform area then you get really small values for Cd. Drag coefficient normally uses the frontal area but sometimes the planform area is used in order to keep the same area as the lift coefficient.

[ChaldExals]

Oh I get it. But still isnt this force very small for such an high velocity?