XFOIL         Version 6.94
  
 Calculated polar for: RAF 19 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   0.8474   0.01294   0.00553  -0.1515  0.4696  0.0166
   0.500   0.8923   0.01260   0.00503  -0.1487  0.4491  0.0440
   1.000   0.9337   0.01262   0.00493  -0.1453  0.4261  0.0555
   1.500   0.9730   0.01279   0.00506  -0.1417  0.4020  0.0699
   2.500   1.0496   0.01400   0.00592  -0.1338  0.3621  0.1037
   3.000   1.0898   0.01445   0.00635  -0.1306  0.3496  0.1138
   3.500   1.1263   0.01502   0.00678  -0.1269  0.3374  0.1227
   4.000   1.1680   0.01542   0.00724  -0.1242  0.3296  0.1299
   4.500   1.2062   0.01587   0.00760  -0.1210  0.3209  0.1364
   5.000   1.2442   0.01644   0.00815  -0.1178  0.3133  0.1439
   5.500   1.2829   0.01673   0.00843  -0.1149  0.3037  0.1495
   6.000   1.3118   0.01726   0.00871  -0.1105  0.2692  0.1536
   6.500   1.3394   0.01813   0.00932  -0.1061  0.2380  0.1577
   7.000   1.3592   0.01950   0.01049  -0.1009  0.1941  0.1620
   8.000   1.3238   0.02724   0.01744  -0.0828  0.0061  0.1674
   8.500   1.3494   0.02896   0.01926  -0.0800  0.0065  0.1736
   9.000   1.3743   0.03096   0.02143  -0.0776  0.0077  0.1820
   9.500   1.3975   0.03331   0.02395  -0.0755  0.0089  0.1928
  10.000   1.4167   0.03622   0.02708  -0.0736  0.0094  0.2052
  10.500   1.4317   0.03978   0.03087  -0.0719  0.0097  0.2188
  11.000   1.4400   0.04424   0.03560  -0.0705  0.0098  0.2329
  11.500   1.4415   0.04975   0.04141  -0.0695  0.0099  0.2477
  12.000   1.4368   0.05648   0.04844  -0.0692  0.0102  0.2638
  12.500   1.4241   0.06511   0.05741  -0.0703  0.0104  0.2885