XFOIL         Version 6.94
  
 Calculated polar for: 20-32C AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.7085   0.01057   0.00301  -0.1301  0.4844  0.0211
   0.500   0.7633   0.01103   0.00328  -0.1298  0.4528  0.0290
   1.000   0.8198   0.01099   0.00287  -0.1298  0.4239  0.0284
   1.500   0.8753   0.01114   0.00281  -0.1298  0.3976  0.0384
   2.000   0.9312   0.01104   0.00317  -0.1303  0.3751  0.2708
   2.500   0.9803   0.01020   0.00350  -0.1292  0.3568  1.0000
   3.000   1.0348   0.01063   0.00376  -0.1291  0.3408  1.0000
   4.500   1.1916   0.01250   0.00469  -0.1281  0.2207  1.0000
   5.000   1.2258   0.01580   0.00685  -0.1256  0.0064  1.0000
   5.500   1.2769   0.01645   0.00757  -0.1248  0.0069  1.0000
   6.000   1.3276   0.01713   0.00845  -0.1239  0.0085  1.0000
   6.500   1.3770   0.01795   0.00948  -0.1226  0.0101  1.0000
   7.000   1.4239   0.01905   0.01087  -0.1209  0.0130  1.0000
   7.500   1.4644   0.02085   0.01306  -0.1180  0.0170  1.0000
   8.000   1.5035   0.02257   0.01506  -0.1148  0.0234  1.0000
   8.500   1.5345   0.02472   0.01753  -0.1102  0.0313  1.0000
   9.000   1.5492   0.02735   0.02040  -0.1037  0.0349  1.0000
   9.500   1.5495   0.03100   0.02432  -0.0968  0.0371  1.0000
  10.000   1.5447   0.03601   0.02953  -0.0915  0.0379  1.0000
  10.500   1.5440   0.04133   0.03504  -0.0873  0.0359  1.0000