XFOIL         Version 6.94
  
 Calculated polar for: JN-153 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   0.7723   0.03093   0.02357  -0.1437  0.4977  0.0345
   0.500   0.8365   0.03041   0.02296  -0.1448  0.4958  0.0494
   1.000   0.8847   0.03097   0.02499  -0.1459  0.4921  0.5219
   1.500   0.8875   0.03434   0.02914  -0.1412  0.4823  0.6720
   2.000   0.9249   0.03377   0.02899  -0.1368  0.4797  1.0000
   2.500   0.9895   0.03354   0.02847  -0.1376  0.4781  1.0000
   3.000   1.0530   0.03352   0.02822  -0.1384  0.4768  1.0000
   3.500   1.1121   0.03412   0.02861  -0.1389  0.4752  1.0000
   4.000   0.8859   0.05172   0.04691  -0.1200  0.4427  1.0000
   4.500   0.9434   0.05158   0.04661  -0.1196  0.4416  1.0000
   5.000   1.0016   0.05132   0.04621  -0.1191  0.4408  1.0000
   5.500   1.0587   0.05105   0.04584  -0.1185  0.4400  1.0000