Naaca 0012 Airfoil Essay

Determining the Drag Polar of the NACA 0012 Airfoil

Wiphu Duangoupa

Ira A. Fulton Schools of Engineering, Tempe, AZ, 85281

The purpose of this lab is to compute the plots for the drag polar of the NACA 0012 airfoil by measuring the lift and drag coefficient of the airfoil in the wind tunnel. The drag polar plots was computed using force balance and wind tunnel environment. The drag polar were simulated using the Reynold’s numbers of 50000, 100000, and 150000 at angle of attacks from -15 to 15 with one degree increment. The plots were created by mainly using the calculated values for the coefficient of lift and the coefficient of drag and their respective uncertainties. It is shown in the drag polar that higher Reynolds number will cause

The set up for the experiment can be seen in figure 2. Where the finite airfoil in figure1 is attached to a force balance that is located right below the wind tunnel. The force balance is attached directly on to the 0012 airfoil. The force balance is the instrument that was used to directly measure the lift and drag forces that were acting on the airfoil.

To start the experiment the ambient pressure and the air temperature was measured using the barometer and thermocouple respectively. The uncertainties were given by the TA, which will also be taken into consideration during the data processing. The air density can be calculated by assuming the air inside the wind tunnel is idea and using a manipulated version of the ideal gas law as shown in equation 1. By using the calculated value for the air density, the V∞ was calculated using equation

Before operating the wind tunnel the NACA 0012 airfoil should be tightly secured, positioned to -15 angle of attack and connected to the force balance located below it. The wind tunnel started to operate at a frequency that will give an approximate Reynolds number of 50000. Once the Reynolds number is stead the airfoil should be rotated after the LABVIEW program collects the data. The airfoil will continually increase its angle of attack until it reaches 15, afterward the steps are repeated for the Reynolds number of 100000 and 150000. Once all the data is collected the coefficient of lift and drag needed to be calculated in order to find the drag pole. To calculate the coefficients equation 3 and 4 was used since the dimensions of the airfoil is known. The Final steps of the data processing was to calculate the velocity using the Sutherland law, which is located in equation 5. After determining the air viscosity the Reynolds number can be calculated with equation 6, which consist of density , velocity , chord length and air