Questions: Airfoil Aerodynamics

Question 1

A pilot increases angle of attack past the stall angle. Which of the following best describes what happens to the airfoil's aerodynamic performance?

• A. Lift drops to zero immediately and the aircraft falls at the rate determined by gravity alone
• B. Lift continues to increase as the angle of attack increases, but drag increases faster
• C. The lift coefficient drops below its maximum value and drag increases sharply, but the wing still produces some lift
• D. The pressure on the upper and lower surfaces equalizes, eliminating both lift and drag

Answer: (C)

Stall does not mean zero lift — it means the lift coefficient has passed C_L,max and is now decreasing as angle of attack increases further. The boundary layer separates from the upper surface (beginning at the trailing edge), disrupting the low-pressure suction distribution that generates most of the lift. Separation spreads forward with increasing alpha, progressively destroying lift and dramatically increasing form drag. The wing still produces some lift throughout — recovery from stall by reducing alpha is possible precisely because the flow reattaches when the adverse pressure gradient is relaxed. Option A is the most common misconception.

Question 2

A NACA 2412 airfoil and a NACA 0012 airfoil are both at an angle of attack of 4°. Which generates more lift, and why?

• A. NACA 0012, because its symmetric shape distributes pressure more evenly and efficiently
• B. NACA 2412, because its 2% camber shifts the zero-lift angle to a negative value, increasing C_L at every angle of attack
• C. They generate the same lift because both have the same thickness (12%) and are at the same angle of attack
• D. NACA 2412, because its higher thickness provides a larger surface area for pressure to act on

Answer: (B)

Camber shifts the zero-lift angle α_L=0 to a negative value. Thin airfoil theory gives C_L = 2π(α − α_L=0). For a symmetric NACA 0012, α_L=0 = 0°, so C_L = 2π(4°) ≈ 0.44. For the cambered NACA 2412, α_L=0 is approximately −2°, so C_L = 2π(4° − (−2°)) = 2π(6°) ≈ 0.66. The cambered airfoil generates lift even at zero geometric angle of attack; at 4° it generates significantly more. Thickness affects stall characteristics and structural depth, not the lift curve slope or zero-lift angle.

Question 3

For a subsonic airfoil, the aerodynamic center (where the pitching moment coefficient is constant regardless of angle of attack) is located at the quarter-chord point.

• T. True
• F. False

Answer: True

True, and this is a fundamental result of thin airfoil theory that holds approximately for most practical subsonic airfoil shapes regardless of camber. The aerodynamic center's fixed location (independent of alpha) makes it the natural reference point for stability analysis — if you know C_m at the aerodynamic center at one angle of attack, you know it at all angles of attack. This is why aircraft wing spars are often placed near the quarter-chord and why tail sizing calculations are referenced to the aerodynamic center. The center of pressure (where the resultant force acts) is a different point that moves as alpha changes.

Question 4

An airfoil with greater thickness will generally stall at a lower angle of attack than a thinner airfoil of the same camber.

• T. True
• F. False

Answer: False

False — the relationship between thickness and stall angle is not monotonically negative. Very thin airfoils actually tend to stall more abruptly via leading-edge separation bubbles, often at relatively low angles of attack, with little warning. Moderate thickness delays this leading-edge stall mechanism and often yields a higher C_L,max and a more gradual trailing-edge stall that pilots can detect (through buffeting) before full separation. Very thick airfoils eventually have worse characteristics from increased form drag, but 'thicker = lower stall angle' is a misconception — the Common Misconceptions section explicitly notes this.

Question 5

Why does increasing angle of attack increase lift up to a maximum, but further increases cause lift to decrease? Explain the role of the boundary layer.

Model answer: Increasing angle of attack sharpens the suction peak near the leading edge on the upper surface, strengthening the low-pressure region and thus increasing lift. However, this also steepens the adverse pressure gradient — the flow must decelerate from the suction peak back to the trailing edge against an increasingly unfavorable pressure rise. The boundary layer on the upper surface, which has been losing momentum due to viscosity, eventually cannot overcome this adverse gradient and separates from the surface, beginning at the trailing edge. Separated flow cannot maintain the organized low-pressure distribution that generates lift, so C_L stops increasing and then decreases. When separation spreads to the leading edge region, the wing stalls: C_L drops sharply, drag rises, and the pitching moment changes abruptly.

The boundary layer is the link between inviscid lift theory (which predicts C_L rising indefinitely with alpha) and the real-world stall. Without viscosity and its associated boundary layer, an airfoil would not stall. The boundary layer accumulates adverse effects along the upper surface until it can no longer follow the surface contour — a fundamentally viscous phenomenon that inviscid potential flow theory cannot capture.