Background:
XFoil is a computational tool that predicts airfoil performance by solving the boundary-layer and inviscid flow equations to yield parameters such as CL (lift coefficient), Cd (drag coefficient), and Cm (moment coefficient). In this analysis, the behavior of the boundary layer (including separation points) and performance curves (e.g., CL and CL/Cd vs. angle of attack, α) are critical for assessing how geometric modifications influence overall aerodynamic efficiency.
Objective:
Identify the aerodynamic characteristics (pressure distribution, CL, Cd, Cm, and CL/Cd) of the NACA 2018, NACA 24xx, and custom airfoil configurations and quantify the effects of geometry (including camber, thickness, and twist) on lift generation, drag, and flow separation.
Methodology:
I used XFoil to generate Cp(pressure coefficient) distributions for the NACA 2018 airfoil at 0° and 10° angles of attack, from which I tabulated aerodynamic coefficients (Cl, Cd, Cm) and examined boundary layer development to identify flow separation regions. Next, I generated performance curves (CL, Cd, Cm, and CL/Cd versus angle of attack from –5° to 12°) for a set of NACA 24xx airfoils at a Reynolds number of 3×10⁶, analyzing how variations in airfoil thickness affected efficiency. Finally, I compared a custom-designed airfoil—featuring increased camber, a shifted maximum camber location, and a thinner trailing edge—with a baseline NACA 4412, evaluating the differences in lift performance over a limited range of angles.
Key Results:
NACA 2018 Analysis: At 0° angle of attack, the airfoil produced modest CL values with minimal separation, while at 10°, the pressure difference between the upper and lower surfaces increased substantially—yielding a CL of 1.2657 versus 0.1942 at 0°—and clearly demonstrated more pronounced flow separation toward the trailing edge (Fig. 1).
NACA 24xx Performance: The performance maps indicated that, above 0°, thinner airfoils achieved higher CL/Cd at lower angles, whereas thicker airfoils improved CL/Cd at higher angles (Fig. 2); specifically, NACA 2418 yielded a peak CL/Cd of approximately 123.85 at 6.5° and NACA 2412 provided the best average CL/Cd of about 70.48 over 5°–12°.
Custom Foil Evaluation: My custom airfoil (Fig. 3) achieved an average CL of approximately 1.50 (for α from 2° to 7°) compared to 0.976 for the NACA 4412, attributable to its greater camber and thinner profile, which led to more aggressive flow redirection and enhanced lift generation.
Conclusion:
The NACA 2018 results validate expected trends in pressure distribution and boundary layer behavior, while the NACA 24xx study elucidates the trade-offs between airfoil thickness and efficiency. My custom airfoil’s enhanced performance demonstrates that deliberate modifications in camber and thickness can yield superior lift characteristics. These insights provide a strong technical foundation for designing high-performance airfoils in aerospace applications.