18.4. Airfoil File Format

VBM results will only be as accurate as the lift and drag data supplied for each blade section. The blade is represented by a stack of airfoils and therefore multiple airfoil files can be used. Rotating components can have high tip velocities, therefore, if possible, the airfoil data file should contain lift and drag data for a range of Reynolds and Mach numbers sufficient to cover the entire spectrum of aerodynamic conditions that will be encountered. Since the receding blades of helicopters in forward flight can experience reversed flow, the data must be provided for the full –180˚ ≤ AoA ≤ +180 ˚ spectrum of angles of attack. The names of the airfoil data files present in the working directory should always have the .dat suffix, as in naca0012.dat. The airfoil file follows the format listed in Table 18.1: Format of the Airfoil Data File.

Table 18.1: Format of the Airfoil Data File

`Airfoilxyz`	Name of the airfoil (same as the airfoil file name and 30 characters maximum)
`Itot`	Total number of and tables in the file (25 airfoil tables maximum (cl plus cd))
`Cl`	Label at the beginning of the table. (10 characters maximum)
`Re`	Reynolds number of the table
`Ma`	Mach number of the table
`Jtot`	Number of lines in the subsequent table
`-180.0 0.00`	Angle of attack and values (two blanks-separated columns and 250 data points maximum)
`…`
`-45.0 -1.50`
`…`
`0.0 0.0`
`…`
`45.0 1.50`
`…`
`180.0 0.00`
`Cd`	Label at the beginning of the table. (10 characters maximum)
`Re`	Reynolds number of the table
`Ma`	Mach number of the table
`Jtot`	Number of lines in the subsequent table
`-180.0 0.006`	Angle of attack and values (two blanks-separated columns and 250 data points maximum)
`…`
`180.0 0.006`
`…`	Repeat, if more and tables are available.

The following shows an example of an airfoil data file.

naca0015
2
cl
100000.0
0.1
 41
     -180.00  0.0000 
     -172.00  0.7800 
     : 
        0.00  0.0000 
     :
      172.50 -0.7800
      180.00  0.0000
cd
100000.0
0.1 
71
     -180.00 0.0220 
     -175.00 0.0620 
     : 
        0.00 0.0088 
      :
      175.00 0.0620 
      180.00 0.0220