XFOIL         Version 6.94
  
 Calculated polar for: SOKOLOV 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   0.5255   0.01141   0.00376  -0.1273  0.6790  0.0945
   0.500   0.5837   0.01125   0.00362  -0.1274  0.6581  0.1388
   1.000   0.6415   0.01116   0.00354  -0.1275  0.6382  0.1671
   2.500   0.8140   0.01085   0.00415  -0.1278  0.5855  0.5181
   3.000   0.8635   0.01009   0.00420  -0.1258  0.5705  1.0000
   3.500   0.9205   0.01038   0.00444  -0.1257  0.5540  1.0000
   4.000   0.9769   0.01069   0.00465  -0.1253  0.5349  1.0000
   4.500   1.0327   0.01104   0.00494  -0.1248  0.5138  1.0000
   5.000   1.0880   0.01139   0.00526  -0.1244  0.4918  1.0000
   5.500   1.1423   0.01175   0.00559  -0.1236  0.4548  1.0000
   6.000   1.1948   0.01224   0.00598  -0.1228  0.4057  1.0000
   6.500   1.2455   0.01299   0.00656  -0.1218  0.3591  1.0000
   7.000   1.2918   0.01423   0.00746  -0.1203  0.3017  1.0000
   7.500   1.3397   0.01528   0.00841  -0.1189  0.2635  1.0000
   8.000   1.3834   0.01666   0.00949  -0.1171  0.1824  1.0000
   8.500   1.4193   0.01874   0.01124  -0.1142  0.1390  1.0000
   9.000   1.4575   0.02038   0.01279  -0.1116  0.0988  1.0000
   9.500   1.4804   0.02317   0.01520  -0.1068  0.0516  1.0000
  10.000   1.5000   0.02564   0.01771  -0.1016  0.0424  1.0000
  10.500   1.5092   0.02829   0.02046  -0.0950  0.0382  1.0000
  11.000   1.5162   0.03156   0.02386  -0.0895  0.0351  1.0000
  11.500   1.5235   0.03532   0.02781  -0.0852  0.0323  1.0000
  12.000   1.5258   0.03999   0.03256  -0.0818  0.0304  1.0000
  12.500   1.5334   0.04452   0.03731  -0.0793  0.0284  1.0000
  13.000   1.5367   0.04959   0.04241  -0.0766  0.0270  1.0000
  13.500   1.5422   0.05469   0.04780  -0.0749  0.0252  1.0000
  14.000   1.5473   0.05978   0.05297  -0.0732  0.0237  1.0000
  14.500   1.5458   0.06603   0.05955  -0.0720  0.0219  1.0000
  15.000   1.5511   0.07136   0.06508  -0.0705  0.0217  1.0000
  15.500   1.5523   0.07732   0.07105  -0.0708  0.0201  1.0000
  16.000   1.5545   0.08337   0.07737  -0.0702  0.0200  1.0000
  17.000   1.5230   0.10230   0.09719  -0.0737  0.0188  1.0000
  18.000   1.4845   0.12419   0.11961  -0.0827  0.0166  1.0000
  18.500   1.4675   0.13616   0.13180  -0.0904  0.0145  1.0000
  19.000   1.4565   0.14733   0.14327  -0.0972  0.0150  1.0000