XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 1098 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   0.5083   0.01320   0.00697  -0.1173  0.6568  0.5964
   0.500   0.5642   0.01337   0.00716  -0.1180  0.6546  0.5994
   1.000   0.6195   0.01341   0.00727  -0.1186  0.6521  0.6028
   1.500   0.6747   0.01356   0.00749  -0.1191  0.6491  0.6061
   2.000   0.7302   0.01369   0.00767  -0.1196  0.6457  0.6098
   2.500   0.7871   0.01379   0.00781  -0.1204  0.6422  0.6142
   3.000   0.8466   0.01386   0.00785  -0.1215  0.6387  0.6180
   3.500   0.9034   0.01410   0.00811  -0.1223  0.6339  0.6217
   4.000   0.9552   0.01393   0.00809  -0.1220  0.6277  0.6264
   4.500   1.0118   0.01375   0.00800  -0.1225  0.6217  0.6312
   5.000   1.0729   0.01368   0.00794  -0.1239  0.6167  0.6366
   5.500   1.1240   0.01384   0.00823  -0.1235  0.6109  0.6417
   6.000   1.1748   0.01370   0.00829  -0.1230  0.6038  0.6488
   6.500   1.2342   0.01340   0.00804  -0.1238  0.5965  0.6562
   7.000   1.2785   0.01322   0.00803  -0.1219  0.5851  0.6636
   7.500   1.3306   0.01287   0.00778  -0.1213  0.5737  0.6726
   8.500   1.3848   0.01267   0.00794  -0.1107  0.5328  0.6940
   9.000   1.3832   0.01324   0.00849  -0.1004  0.4896  0.7079
   9.500   1.3505   0.01531   0.01033  -0.0865  0.4334  0.7246
  10.000   1.3209   0.01844   0.01336  -0.0754  0.3858  0.7473
  10.500   1.2927   0.02237   0.01725  -0.0664  0.3406  0.7801
  11.000   1.2695   0.02626   0.02124  -0.0585  0.2999  0.8584