## Wiki

The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described using a series of digits following the word "NACA". The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

This repository contains basic data on the NACA airfoils: profile coordinates and performance data, including lift coefficients, drag coefficients, and pitching moment coefficients for various Reynolds numbers. For each coefficient distribution by the attack angle, corresponding relations are constructed and can be used in design calculations via SplineCloud API.

## 4-digit series

The NACA four-digit wing sections define the profile by:

• First digit describing maximum camber as percentage of the chord.
• Second digit describing the distance of maximum camber from the airfoil leading edge in tenths of the chord.
• Last two digits describing maximum thickness of the airfoil as percent of the chord.

For 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.

The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.

## 5-digit series

The NACA five-digit series describes more complex airfoil shapes. Its format is LPSTT, where:

• L: a single digit representing the theoretical optimal lift coefficient at ideal angle of attack CLI = 0.15 L (this is not the same as the lift coefficient CL),
• P: a single digit for the x coordinate of the point of maximum camber (max. camber at x = 0.05 P),
• S: a single digit indicating whether the camber is simple (S = 0) or reflex (S = 1),
• TT: the maximum thickness in percent of chord, as in a four-digit NACA airfoil code.

For example, the NACA 23112 profile describes an airfoil with design lift coefficient of 0.3 (0.15 × 2), the point of maximum camber located at 15% chord (5 × 3), reflex camber (1), and maximum thickness of 12% of chord length (12).

## 6-series

An improvement over 1-series airfoils with emphasis on maximizing laminar flow. The airfoil is described using six digits in the following sequence:

• The number "6" indicating the series.
• One digit describing the distance of the minimum pressure area in tenths of the chord.
• The subscript digit gives the range of lift coefficient in tenths above and below the design lift coefficient in which favorable pressure gradients exist on both surfaces.
• A hyphen.
• One digit describing the design lift coefficient in tenths.
• Two digits describing the maximum thickness as percent of chord.
• "a=" followed by a decimal number describing the fraction of chord over which laminar flow is maintained. a=1 is the default if no value is given.

For example, the NACA 612-315 a=0.5 has the area of minimum pressure 10% of the chord back, maintains low drag 0.2 above and below the lift coefficient of 0.3, has a maximum thickness of 15% of the chord, and maintains laminar flow over 50% of the chord.

The Reynolds number is a dimensionless value that measures the ratio of inertial forces to viscous forces and descibes the degree of laminar or turbulent flow. Systems that operate at the same Reynolds number will have the same flow characteristics even if the fluid, speed and characteristic lengths vary.

Reynolds number range from 50,000 to 1,000,000 in approximatly logarithmic steps.

Ncrit value is used to model of the turbulence of the fluid or roughness of the airfoil.

Current repository contains data for average wind tunnel with Ncrit=9

Mach number has been left at the default value of zero.

## References

http://airfoiltools.com/