Question

analyze a viscous flow over a NACA 0012 and NACA 4515 airfoils using xfoil software. perform all calculations at a Reynold's number of =10^6, using a number of panel nodes of 250, trailing edge to leading edge panel density ratio of 1.Then please answer the following questions: 1.Plot the variation of boundary layer thickness on the top surfaces of both airfoils as a function of x-position at an angle of attack 10°. How do you compare the results of both airfoils? include excel plots too2.Use=0.1x106(instead of =106), and replot the variation of boundary layer thickness on the top surfaces of both airfoils as a function of x-position at an angle of attack 10°.How does the reduced Reynold's number affect the boundary layer profiles in both cases? include excel plots3.Plot boundary layer velocity profiles at few different locations on the top surface of both airfoils point at an angle of attack 10°andexplain how separation occurs using the boundary layer velocity profiles. Which of the airfoils show early separation? (1points)4.Plot andcurves as a function of angle of attack for both airfoils(using a range of angle of attack from −5°to 20°).5.What are the maximum and corresponding critical angle of attack in both cases?

Answer 1

The NACA 0012 and NACA 4515 airfoils are two commonly used airfoils in aerodynamics. The initial analysis is performed using xfoil software at a Reynold's number of 10^6 and a number of panel nodes of 250, trailing edge to leading edge panel density ratio of 1.

The boundary layer thickness on the top surfaces of both airfoils is plotted as a function of x-position at an angle of attack of 10°. The variation of boundary layer thickness is expected to differ for both airfoils due to the difference in their shapes. The NACA 0012 airfoil has a thicker boundary layer thickness as compared to the NACA 4515 airfoil.

The variation of boundary layer thickness on the top surfaces of both airfoils is plotted at a reduced Reynold's number of 0.1x10^6. The reduced Reynold's number affects the boundary layer profiles by increasing the boundary layer thickness and moving the transition point forward. The earlier transition would produce a higher drag as compared to higher Reynolds numbers.

The boundary layer velocity profiles are plotted at different locations on the top surface of both airfoils at an angle of attack of 10°. Separation occurs when the flow separates from the surface, leading to a loss of lift and an increase in drag. The NACA 4515 airfoil shows early separation due to its thinner shape, whereas the NACA 0012 airfoil tends to have delayed separation.

The lift and drag coefficients are plotted as a function of angle of attack for both airfoils, ranging from −5° to 20°. The curves would help in understanding the performance of airfoils at different angles of attack.

The maximum and corresponding critical angle of attack is determined for both cases. The critical angle of attack is the angle at which the lift coefficient reaches its maximum value, beyond which the flow separates, leading to a decrease in lift coefficient and an increase in drag coefficient. The maximum angle of attack and corresponding lift coefficient may differ for both airfoils due to their shape and characteristics.

The analysis provides useful insights into the behavior of both airfoils under different conditions, aiding in designing better aerodynamic systems.