p is the location of maximum camber (10 p is the second digit in the NACA xxxx description).įor this cambered airfoil, because the thickness needs to be applied perpendicular to the camber line, the coordinates and, of respectively the upper and lower airfoil surface, become:.m is the maximum camber (100 m is the first of the four digits),.The formula used to calculate the mean camber line is: The simplest asymmetric foils are the NACA 4-digit series foils, which use the same formula as that used to generate the 00xx symmetric foils, but with the line of mean camber bent. The camber line is shown in red, and the thickness – or the symmetrical airfoil 0012 – is shown in purple. The equation for a cambered 4-digit NACA airfoil The leading edge approximates a cylinder with a radius of: to −0.1036) will result in the smallest change to the overall shape of the airfoil. If a zero-thickness trailing edge is required, for example for computational work, one of the coefficients should be modified such that they sum to zero. Note that in this equation, at ( x/ c) = 1 (the trailing edge of the airfoil), the thickness is not quite zero. t is the maximum thickness as a fraction of the chord (so 100 t gives the last two digits in the NACA 4-digit denomination).is the half thickness at a given value of x (centerline to surface), and.x is the position along the chord from 0 to c,.The formula for the shape of a NACA 00xx foil, with "xx" being replaced by the percentage of thickness to chord, is: Plot of a NACA 0015 foil, generated from formula The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.Įquation for a symmetrical 4-digit NACA airfoil The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. Four-digit series airfoils by default have maximum thickness at 30% of the chord (0.3 chords) from the leading edge. The NACA four-digit wing sections define the profile by: įor example, the NACA 2412 airfoil has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord. 1.2 The equation for a cambered 4-digit NACA airfoil.1.1 Equation for a symmetrical 4-digit NACA airfoil.Once the critical angle of attack is reached (generally around 14 degrees) the aerofoil will stall. The subsequent loss of static pressure creates a pressure difference between the upper and lower surfaces that is called lift and opposes the weight of an aircraft (or thrust that opposes drag).Īs the angle of attack (the angle between the chord line and relative air flow) is increased, more lift is created. Basically this states that total pressure is equal to static pressure (due to the weight of air above) plus dynamic pressure (due to the motion of air).Īir that travels over the top surface of the aerofoil has to travel faster and thus gains dynamic pressure. The basic principle behind an aerofoil is described by bernoullis theorem. ![]() Alternatively, a jet would have a thin wing with minimal camber to allow it to cruise at high speeds. For example, a crop duster may have a thick, high camber wing that produces a large amount of lift at low speed. Point of Maximum Thickness = Thickest part of the wing expressed as a percentage of the chordīy altering each of the above features of an aerofoil, the designer is able to adjust the performance of the wing so that it is suitable for it's particular task.Denotes the amount of curvature of the wing Mean Camber Line = Line drawn half way between the upper and lower surface of the aerofoil.Chord = Line connecting the leading and trailing edge. ![]()
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