XFOIL         Version 6.94
  
 Calculated polar for: Apex 16 (normalized using XFOIL date021206)     
  
 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   0.2425   0.01389   0.00612  -0.0652  0.8087  0.0317
   1.000   0.3730   0.01081   0.00549  -0.0688  0.7929  0.7045
   1.500   0.4649   0.01075   0.00562  -0.0761  0.7812  0.7686
   2.000   0.4657   0.01079   0.00579  -0.0644  0.7680  0.8012
   2.500   0.5057   0.01049   0.00560  -0.0607  0.7530  0.8409
   3.000   0.5621   0.01003   0.00517  -0.0601  0.7372  0.8762
   3.500   0.6795   0.00964   0.00471  -0.0724  0.7095  0.9092
   4.000   0.7079   0.00974   0.00500  -0.0669  0.6923  0.9766
   4.500   0.7824   0.00976   0.00497  -0.0710  0.6644  1.0000
   5.000   0.8251   0.00991   0.00519  -0.0687  0.6387  1.0000
   5.500   0.8515   0.01005   0.00521  -0.0628  0.6004  1.0000
   6.000   0.8874   0.01040   0.00549  -0.0590  0.5612  1.0000
   6.500   0.9171   0.01095   0.00582  -0.0541  0.5049  1.0000
   7.000   0.9424   0.01170   0.00635  -0.0487  0.4424  1.0000
   7.500   0.9564   0.01284   0.00716  -0.0415  0.3615  1.0000
   8.000   0.9480   0.01472   0.00841  -0.0311  0.2533  1.0000
   8.500   0.9378   0.01671   0.00995  -0.0212  0.1692  1.0000
   9.000   0.9359   0.01890   0.01161  -0.0135  0.0817  1.0000