XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0712 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.5348   0.00886   0.00472  -0.1229  0.8692  0.7053
   0.500   0.5913   0.00888   0.00451  -0.1221  0.7953  0.7109
   1.000   0.6288   0.01044   0.00484  -0.1176  0.5205  0.7141
   1.500   0.6706   0.01243   0.00553  -0.1154  0.2509  0.7187
   2.000   0.7199   0.01358   0.00610  -0.1143  0.1376  0.7225
   2.500   0.7723   0.01436   0.00666  -0.1135  0.0980  0.7268
   3.000   0.8264   0.01505   0.00724  -0.1132  0.0758  0.7313
   3.500   0.8781   0.01580   0.00795  -0.1121  0.0631  0.7347
   4.500   0.9805   0.01757   0.00972  -0.1100  0.0507  0.7431
   6.000   1.1279   0.02079   0.01321  -0.1057  0.0370  0.7556
   6.500   1.1759   0.02220   0.01473  -0.1041  0.0349  0.7609
   7.000   1.2212   0.02370   0.01638  -0.1020  0.0334  0.7649
   7.500   1.2672   0.02540   0.01819  -0.1003  0.0320  0.7704
   8.000   1.3068   0.02844   0.02156  -0.0975  0.0303  0.7739
   8.500   1.3443   0.03189   0.02541  -0.0947  0.0292  0.7794
   9.000   1.3812   0.03345   0.02726  -0.0914  0.0281  0.7835
   9.500   1.4120   0.03602   0.03017  -0.0876  0.0267  0.7888
  10.000   1.4310   0.03953   0.03411  -0.0822  0.0254  0.7929
  10.500   1.4298   0.04397   0.03901  -0.0742  0.0247  0.7979
  11.000   1.4085   0.04967   0.04524  -0.0650  0.0243  0.8025
  11.500   1.3672   0.05781   0.05396  -0.0571  0.0242  0.8070
  12.000   1.3030   0.06998   0.06673  -0.0539  0.0243  0.8107