Week 7 - Minor Project
#) Objective - To set up the simulation over the NACA 0012 airfoil . That is a symmetrical airfoil.
To find out the drag coefficient vs the angle of attack .
The lift coefficient vs the angle of attack .
Compare the effect of turbulence odels on the above results .
The inlet Reynolds number is of 2 * 10^5
The nomenclature of the airfoil is as follows -
The above image shows the basic nomenclature of the airfoil system . Aero0oil is the cross section of the airfoil normasl to span .
The chamber of the symmetrical airfoil is 0 because the chamber line and the chord line coincides with it .
Chord is the distance between the leading and trailing edge .
Aspect ratio is the ratio between wing span and the mean chord .
Angle of attack is the angle between the relative chord and the direction of the wind .
Aerofoil theory - The aerofoil is a streamlined body designed in such a way that the separation occures at the extreme rear end of the body . Such that it will produce small wake and the subsequently small pressure drag . Even for the high Reynolds number the pressure drag is very small the skin fr4iction drag is the largest drag component .
The first two dijit in the NACA aerofoil representsw the chamber of the aerofoil and the last two digit represents the thickness of the aerofoil . So thickness is the 12 percent of the chord .
#) The geometry setup -
The aerofoil coordinates are imported into the spaceclaim and the geometry of the airfoil is created in spaceclaim .
The two dimensional domain of the aerofoil is created in the spacecloaim using the following dimensions as follows -
For subsonic flow the results obtained are as follows-
We take a freestream velocity (v) = 100 m/sec
To take into account of the angle of attack of alpha = 5 degrees the x component of velocity = 100 * cos 5 = 99.619 m/sec
The y component of velocity vy = 100 *sin5= 6.67 m/sec
The contours of velocity magnitude obtained is-
The contours of static pressure obtained is-
The lift coefficients obtained is- 0.25 and the drag coefficient obtained is - 0.33
For 10 degree angle of attack we got the velocity contours as follows-
For 15 degree aoa the velocity contour obtained is -
The pressure contour is -
The lift coefficient value has dropped as we go from the 5 degree aoa to 10 degrees and to 15 degrees aoa since aoa is getting closer to the stall aoa .
#) For the supersoinic flow we consider the freestream velocity = 350 m/sec.
The x component of nvelocity vx = 350 * cos5 =
The y component of velocity vy = 350 *sin5 =
The contour of velocity obtained is -
The static pressure contour obtained is -
The lift and drqag coefficient data obtained are-
For 10 degree angle of attack-
The velocity contour obtained is -
For 15 degree aoa velocity contour is -
the lift and drag coefficent data are-
Conclusion - The NACA 0012 airfoil is well suited for the subsonic flows but in the supersonic regime the shock waves are created which creates lparge gradient variation in the lift and drag coefficients.