CFD Analysis of a Swept Back Wing
This project performs a computational analysis to verify the experimental results carried out on NACA 64015A Sweptback wing. The case used here is at sweep angle of 20 degree and zero degree Angle of Attack. (Reference: NASA Technical note D338 -
NACA 64015A airfoil geometry was imported from the UIUC Airfoil Coordinates Database and the wing geometry was created on Solidworks.
Figure 1: Swept Wing CAD Model.
For generating mesh, snappyHexMesh was used. A box of dimension 15mm x 10mm x 5mm was created which acted as the farfield. The swept wing of wing span 2.5mm was placed 5 mm aft from the inlet face of the box. Prior to meshing the surfaces were grouped into topological entity sets namely Inlet, Outlet, Walls and the SweptWing. This allowed to set different refinement levels for small and large surfaces easily.
Figure 2: snappyHexMesh
Three refinement layers of mesh were generated over the wing surface growing further to the farfield surface.
For the computational analysis, Steady State Incompressible fluid flow type was selected with K-Omega-SST Turbulence Model with SIMPLE solver. Time step of 1000 sec were solved with time step length of 2 Sec. The analysis was carried out on 4 cores, taking 110 mins. The picture below shows the convergence history.
Figure 3: Convergence History.
The post- processing was carried out on the SimScale Platform. The picture below shows the Velocity and Pressure contours map over the mid plan of the wing span.
Figure 4: Pressure Contour Plot.
Figure 5: Velocity Contour Plot.
To verify the experimental results, pressure were plotted over a line on the across the airfoil cross section. The pressure co – efficient Cp was plotted against x/c of the airfoil and comapred with experimental results. The graphs below shows Cp vs x/c plot (left) and the experimental results (Right).