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is the half thickness at a given value of x (centerline to surface), t is the maximum thickness as a fraction of the chord (so t gives the last two digits in the NACA 4-digit denomination divided by 100). In this equation, at x = 1 (the trailing edge of the airfoil), the thickness is not quite zero. If a zero-thickness trailing edge is required ...
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.
Camber is a complex property that can be more fully characterized by an airfoil's camber line, the curve Z(x) that is halfway between the upper and lower surfaces, and thickness function T(x), which describes the thickness of the airfoils at any given point. The upper and lower surfaces can be defined as follows:
, is the coefficient of lift of a specific section of the airfoil, t is the airfoil thickness at a given section, c is the chord length at a given section, is a factor established through CFD analysis: K = 0.87 for conventional airfoils (6 series), [4] K = 0.95 for supercritical airfoils.
The Mitsubishi A6M "Zero" is a long-range carrier-capable fighter aircraft formerly manufactured by Mitsubishi ... Airfoil: root: MAC118 or NACA 2315; tip: MAC118 or ...
Therefore, the Drag coefficient on a supersonic airfoil is described by the following expression: C D = C D,friction + C D,thickness + C D,lift. Experimental data allow us to reduce this expression to: C D = C D,O + KC L 2 Where C DO is the sum of C (D,friction) and C D,thickness, and k for supersonic flow is a function of the Mach number. [3]
The thermal boundary layer thickness is similarly the distance from the body at which the temperature is 99% of the freestream temperature. The ratio of the two thicknesses is governed by the Prandtl number. If the Prandtl number is 1, the two boundary layers are the same thickness.
The fact of the matter is, the NACA 4-digit series was empirically derived from real airfoils, none of which actually have a zero-thickness trailing edge (obviously). However, most people these days are probably using these airfoil equations for computational studies, where it is best to close the airfoil.