XFOIL         Version 6.94
  
 Calculated polar for: NACA M14 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3206   0.01212   0.00411  -0.0586  0.7004  0.0453
   0.500   0.3752   0.01149   0.00337  -0.0581  0.6819  0.0444
   1.000   0.4303   0.01109   0.00292  -0.0579  0.6631  0.0463
   1.500   0.4858   0.01083   0.00264  -0.0576  0.6447  0.0663
   3.000   0.6576   0.00899   0.00263  -0.0587  0.5507  1.0000
   3.500   0.7095   0.00927   0.00271  -0.0579  0.5079  1.0000
   4.000   0.7600   0.00975   0.00291  -0.0570  0.4410  1.0000
   4.500   0.8114   0.01026   0.00324  -0.0564  0.3897  1.0000
   5.000   0.8389   0.01412   0.00504  -0.0534  0.0252  1.0000
   5.500   0.8885   0.01493   0.00578  -0.0525  0.0081  1.0000
   6.000   0.9383   0.01567   0.00672  -0.0515  0.0075  1.0000
   6.500   0.9869   0.01654   0.00793  -0.0503  0.0076  1.0000
   7.000   1.0326   0.01773   0.00938  -0.0486  0.0079  1.0000
   7.500   1.0722   0.01950   0.01145  -0.0461  0.0083  1.0000
   8.000   1.1014   0.02207   0.01431  -0.0422  0.0089  1.0000
   8.500   1.1222   0.02533   0.01782  -0.0372  0.0095  1.0000
   9.000   1.1453   0.02998   0.02269  -0.0326  0.0102  1.0000
   9.500   1.1835   0.03246   0.02549  -0.0299  0.0116  1.0000
  10.000   1.1763   0.04113   0.03519  -0.0227  0.0175  1.0000