Abstract:
In this thesis, the result of experimental study of aerodynamic coefficients, CL, CD, Cm,
coefficients of pressure, CP, and flowfield around NACA 0012 airfoil has been described. An
open loop subsonic wind tunnel, AF100, was used to conduct the experiment for two
reynolds number 10.15 104 and 15.23 104
and with different angle of attack- 0 degree, 5
degree, 10 degree, 15 degree, 20 degree. Pitot tube was used to measure the flow velocity.
Again yawmeter was used to collect and calculated data of flow angularity around airfoil.
The data of lift, drag and pitching moment were collected from VDAS for calculating CL, CD
and Cm. Pressure coefficients Cp were calculated from the pressure readings of 20 pressure
tappings around the midplane of the airfoil. The result shows that no lift is generated at 0
degree AOA, but with the increase of AOA, significant amount of lift and drag is generated.
Vortex was found to occur at the trailing edge due to flow separartion. The result shows that
with the increasing AOA, flow separation region is getting increased. For greater reynolds
number, the phenomena is more prominent.
