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NACA 0012 Airfoil Simulation
5 Attachment(s)
Hello, Dear OpenFOAM Users,
I am trying to validate NACA 0012 airfoil simulation using OpenFOAM 5.0x using simpleFoam solver. I created a C-H type mesh using Pointwise meshing tool and successfully exported to OpenFOAM. My simulations inputs are as follow: (1): Free Stream Velocity = 45 m/s (2): Mach Number = 0.13 (3): Free Stream Pressure ( (101325 pa) and Temperature (300 k ) (4): Density of air = 1.225 (5): Kinematic Viscosity = 1.4607*10^-5 (6): Reynolds Number = 3.0807 *10^6 I run this case for Angle of Attack (A.O.A) = 10 degrees and I got the results for Lift and Drag Coefficients like Cd = -0.00339582 Cl = 0.0797399 which seems incorrect. Can someone help me how to get the right values of Cl and Cd ? Attachment 67969 Attachment 67970 Attachment 67971 Attachment 67972 Attachment 67969 Attachment 67970 Attachment 67971 Attachment 67972 Attachment 67973 |
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