XFOIL         Version 6.94
  
 Calculated polar for: AH 88-K-130/20                                  
  
 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   0.4978   0.01135   0.00680  -0.1158  0.7795  0.8654
   0.500   0.5439   0.01152   0.00699  -0.1139  0.7740  0.8807
   1.000   0.5944   0.01147   0.00696  -0.1132  0.7674  0.8891
   1.500   0.6523   0.01140   0.00687  -0.1143  0.7614  0.8955
   2.000   0.7097   0.01107   0.00653  -0.1148  0.7560  0.9006
   2.500   0.7657   0.01058   0.00601  -0.1148  0.7463  0.9060
   3.000   0.8204   0.01030   0.00584  -0.1150  0.7363  0.9114
   3.500   0.8801   0.00987   0.00542  -0.1160  0.7266  0.9165
   4.000   0.9303   0.00952   0.00515  -0.1150  0.7133  0.9223
   4.500   0.9831   0.00901   0.00469  -0.1144  0.6826  0.9270
   5.000   1.0361   0.00904   0.00471  -0.1143  0.6565  0.9312
   5.500   1.0760   0.00933   0.00485  -0.1114  0.6032  0.9375
   6.000   1.1031   0.01026   0.00543  -0.1065  0.5111  0.9433
   7.000   1.0677   0.01619   0.00968  -0.0845  0.1883  0.9783
   7.500   1.0740   0.01972   0.01252  -0.0799  0.0580  1.0000
   8.000   1.1009   0.02210   0.01479  -0.0777  0.0249  1.0000
   8.500   1.1247   0.02473   0.01755  -0.0749  0.0141  1.0000
   9.500   1.1630   0.03069   0.02384  -0.0688  0.0092  1.0000
  10.000   1.1883   0.03322   0.02653  -0.0666  0.0084  1.0000
  10.500   1.2093   0.03620   0.02972  -0.0640  0.0063  1.0000
  11.000   1.2243   0.03991   0.03364  -0.0607  0.0055  1.0000
  11.500   1.2413   0.04421   0.03827  -0.0567  0.0050  1.0000
  12.000   1.2644   0.04861   0.04308  -0.0537  0.0051  1.0000
  12.500   1.2627   0.06179   0.05754  -0.0469  0.0074  1.0000
  13.000   1.2214   0.07270   0.06903  -0.0430  0.0085  1.0000