XFOIL         Version 6.94
  
 Calculated polar for: E176  (8.83%)                                   
  
 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   0.3298   0.00659   0.00182  -0.0640  0.6582  1.0000
   0.500   0.3840   0.00670   0.00181  -0.0636  0.6465  1.0000
   1.000   0.4386   0.00685   0.00185  -0.0633  0.6352  1.0000
   1.500   0.4934   0.00704   0.00191  -0.0630  0.6241  1.0000
   2.000   0.5483   0.00718   0.00200  -0.0627  0.6115  1.0000
   2.500   0.6033   0.00732   0.00212  -0.0625  0.5974  1.0000
   3.000   0.6582   0.00747   0.00224  -0.0622  0.5820  1.0000
   3.500   0.7130   0.00763   0.00243  -0.0619  0.5661  1.0000
   4.000   0.7676   0.00781   0.00261  -0.0616  0.5482  1.0000
   4.500   0.8218   0.00797   0.00282  -0.0612  0.5221  1.0000
   5.000   0.8744   0.00825   0.00302  -0.0606  0.4774  1.0000
   5.500   0.9235   0.00896   0.00340  -0.0596  0.3867  1.0000
   6.000   0.9646   0.01079   0.00448  -0.0580  0.2260  1.0000
   6.500   0.9988   0.01343   0.00614  -0.0557  0.0567  1.0000
   7.000   1.0390   0.01524   0.00773  -0.0536  0.0139  1.0000
   7.500   1.0787   0.01696   0.00969  -0.0512  0.0114  1.0000
   8.000   1.1140   0.01887   0.01186  -0.0484  0.0110  1.0000
   8.500   1.1417   0.02125   0.01447  -0.0446  0.0109  1.0000
   9.000   1.1678   0.02360   0.01701  -0.0407  0.0102  1.0000
   9.500   1.1868   0.02629   0.01991  -0.0360  0.0098  1.0000
  10.000   1.2083   0.02978   0.02369  -0.0320  0.0097  1.0000
  10.500   1.2310   0.03413   0.02845  -0.0286  0.0098  1.0000
  11.000   1.2418   0.04001   0.03484  -0.0248  0.0100  1.0000
  11.500   1.2314   0.04750   0.04288  -0.0211  0.0107  1.0000