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For example, a NACA 2412 airfoil uses a 2% camber (first digit) 40% (second digit) along the chord of a 0012 symmetrical airfoil having a thickness 12% (digits 3 and 4) of the chord.
In February 1976, work commenced to automate the methods contained in the USAF Stability and Control DATCOM, specifically those contained in sections 4, 5, 6 and 7.The work was performed by the McDonnell Douglas Corporation under contract with the United States Air Force in conjunction with engineers at the Air Force Flight Dynamics Laboratory in Wright-Patterson Air Force Base.
For example, an airfoil of the NACA 4-digit series such as the NACA 2415 (to be read as 2 – 4 – 15) describes an airfoil with a camber of 0.02 chord located at 0.40 chord, with 0.15 chord of maximum thickness. Finally, important concepts used to describe the airfoil's behaviour when moving through a fluid are:
English: Selected airfoils in nature and various vehicles, with their approximate chord length indicated. Sources for the shapes of the airfoils: Low-speed ULM wing: drawn over own photo of low-cost, low-speed ultralight; Propeller blade: drawn over own photo of a sliced WW2-era bomber propeller
Further development resulted in two patents and a family of airfoils known as the KF airfoil and KFm airfoils (for Kline–Fogleman modified). The two patents, US Patent # 3,706,430 and US Patent # 4,046,338, refer to the introduction of a step on either the bottom (KFm1) or the top (KFm2) of an airfoil , or on both the top and bottom (KFm4).
a=chord, b=thickness, thickness-to-chord ratio = b/a The F-104 wing has a very low thickness-to-chord ratio of 3.36%. In aeronautics, the thickness-to-chord ratio, sometimes simply chord ratio or thickness ratio, compares the maximum vertical thickness of a wing to its chord.
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Mean aerodynamic chord (MAC) is defined as: [6] = (), where y is the coordinate along the wing span and c is the chord at the coordinate y.Other terms are as for SMC. The MAC is a two-dimensional representation of the whole wing. The pressure distribution over the entire wing can be reduced to a single lift force