XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -4.500  -0.1086   0.02588   0.01910  -0.0812   0.6847   0.5768
  -4.000  -0.0625   0.02892   0.02201  -0.0753   0.6791   0.6258
  -3.500  -0.0228   0.03140   0.02448  -0.0692   0.6724   0.6572
  -3.000   0.0184   0.03215   0.02504  -0.0673   0.6654   0.6829
  -2.500   0.0654   0.03239   0.02501  -0.0663   0.6608   0.7011
  -2.000   0.1025   0.03342   0.02595  -0.0656   0.6551   0.7184
  -1.500   0.1347   0.03467   0.02715  -0.0642   0.6477   0.7355
  -1.000   0.1770   0.03526   0.02761  -0.0631   0.6433   0.7490
  -0.500   0.2063   0.03717   0.02947  -0.0627   0.6383   0.7616
   0.000   0.1845   0.04299   0.03544  -0.0618   0.6312   0.7741
   0.500   0.1807   0.04718   0.03964  -0.0604   0.6289   0.7866
   1.000   0.1854   0.05197   0.04445  -0.0606   0.6368   0.7978
   1.500   0.2305   0.05476   0.04715  -0.0627   0.6399   0.8108
   3.500   0.2645   0.06969   0.06219  -0.0613   0.6541   0.8636
   4.000   0.3122   0.07203   0.06452  -0.0623   0.6433   0.8814
   4.500   0.3295   0.07422   0.06677  -0.0606   0.6287   0.9012
   5.000   0.3931   0.07769   0.07033  -0.0631   0.6235   0.9345
   6.000   0.4345   0.08372   0.07640  -0.0657   0.5940   1.0002
   6.500   0.5075   0.08829   0.08093  -0.0710   0.5875   1.0002
   7.500   0.5358   0.09582   0.08849  -0.0720   0.5616   1.0002
   8.000   0.5879   0.09988   0.09257  -0.0744   0.5524   1.0002
   8.500   0.5919   0.10394   0.09667  -0.0739   0.5394   1.0002
   9.000   0.6550   0.10892   0.10172  -0.0766   0.5334   1.0002
  10.000   0.6623   0.11728   0.11023  -0.0757   0.5084   1.0002