XFOIL         Version 6.94
  
 Calculated polar for: NACA M23 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   0.3233   0.01830   0.01085  -0.0268  0.5569  0.0310
   0.500   0.3809   0.01703   0.00913  -0.0268  0.5481  0.0371
   1.000   0.4370   0.01634   0.00822  -0.0271  0.5397  0.0430
   1.500   0.4942   0.01581   0.00754  -0.0276  0.5316  0.0554
   2.000   0.5513   0.01449   0.00592  -0.0274  0.5233  0.0510
   2.500   0.6058   0.01391   0.00535  -0.0273  0.5155  0.0534
   3.000   0.6606   0.01352   0.00507  -0.0274  0.5072  0.0586
   3.500   0.7152   0.01334   0.00488  -0.0274  0.4991  0.0704
   5.500   0.9925   0.01215   0.00485  -0.0416  0.3435  1.0000
   6.000   1.0226   0.01744   0.00838  -0.0424  0.0097  1.0000
   6.500   1.0668   0.01841   0.00946  -0.0412  0.0083  1.0000
   7.000   1.1084   0.01954   0.01079  -0.0397  0.0091  1.0000
   7.500   1.1429   0.02119   0.01270  -0.0375  0.0103  1.0000
   8.000   1.1682   0.02333   0.01510  -0.0346  0.0119  1.0000
   8.500   1.1591   0.02764   0.01964  -0.0302  0.0128  1.0000
   9.000   1.1592   0.03233   0.02457  -0.0278  0.0145  1.0000
   9.500   1.1504   0.03835   0.03075  -0.0261  0.0159  1.0000
  10.000   1.1436   0.04400   0.03650  -0.0236  0.0182  1.0000
  10.500   1.1643   0.04636   0.03897  -0.0187  0.0238  1.0000