Airfoils
Geometry that matters
An aerofoil is described by its chord line, camber line, thickness distribution and leading-edge radius. Candidates should be able to define chord, camber, angle of attack and angle of incidence without confusion. Angle of attack is measured between the chord line and the relative airflow; angle of incidence is a fixed rigging angle between chord and a fuselage reference line. Questions frequently swap these two.
Camber generates lift at zero angle of attack and shifts the lift curve upward and the zero-lift angle negative. Increasing thickness increases maximum lift but also drag; a very sharp leading edge promotes early separation, while a rounded one tolerates higher angles before stalling. Symmetric sections have no camber and no pitching moment about the aerodynamic centre; cambered sections produce a nose-down pitching moment that is roughly constant with angle of attack until the stall.
The lift curve
Plot lift coefficient against angle of attack. The slope is approximately 0.1 per degree for thin sections in incompressible flow, linear until near the critical angle, then the curve peaks and falls away. The stall angle for light aircraft sections is typically 15–18 degrees. Understanding that the curve is linear, then rounded, then abruptly descending helps you answer questions about approach to the stall and about the effect of flaps, ice and Reynolds number.
Traps and a worked example
Traps: assuming a higher-lift aerofoil always improves performance (it usually costs drag); assuming flaps increase the stall angle (they usually reduce it while raising maximum lift); and forgetting that a rough or contaminated leading edge reduces maximum lift and may reduce the stall angle.
Rule of thumb: lift-curve slope about 0.1 per degree means 10 degrees of angle of attack yields a lift coefficient near 1.0 for a cambered section.
Worked example: if a section has zero-lift angle −2° and slope 0.1 per degree, at 8° angle of attack the lift coefficient is approximately 0.1 × (8 + 2) = 1.0.