XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-1010 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.8289   0.01391   0.00638  -0.1338  0.1865  0.8026
   1.000   0.8705   0.01575   0.00757  -0.1311  0.0800  0.8164
   1.500   0.9182   0.01687   0.00869  -0.1291  0.0629  0.8302
   2.500   1.0083   0.01948   0.01137  -0.1242  0.0441  0.8574
   3.000   1.0527   0.02110   0.01302  -0.1217  0.0401  0.8708
   4.000   1.1461   0.02500   0.01740  -0.1173  0.0340  0.8988
   4.500   1.1915   0.02719   0.01980  -0.1150  0.0313  0.9130
   6.000   1.2965   0.04134   0.03564  -0.1042  0.0279  1.0001
   6.500   1.3144   0.04816   0.04320  -0.0993  0.0276  1.0001
   7.000   1.3107   0.05626   0.05204  -0.0922  0.0274  1.0001
   7.500   1.2800   0.06474   0.06114  -0.0831  0.0273  1.0001
   8.000   1.2294   0.07366   0.07053  -0.0749  0.0273  1.0001
   8.500   1.1715   0.08546   0.08272  -0.0734  0.0277  1.0001