XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1098 AIRFOIL
                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -5.000  -0.3360   0.03785   0.02928  -0.0780   0.8196   0.0928
  -4.500  -0.3239   0.03511   0.02910  -0.0757   0.8201   0.5045
  -4.000  -0.3673   0.03778   0.03177  -0.0643   0.8544   0.5355
  -3.000  -0.4043   0.04167   0.03567  -0.0430   0.9245   0.6026
  -2.000  -0.3573   0.04920   0.04329  -0.0218   0.8985   0.7017
  -1.000  -0.2771   0.04992   0.04346  -0.0206   0.8718   0.7291
  -0.500  -0.2323   0.05059   0.04388  -0.0216   0.8598   0.7367
   1.000  -0.0677   0.05429   0.04692  -0.0311   0.8226   0.7554
   2.000   0.0279   0.05685   0.04922  -0.0356   0.7934   0.7654
   2.500   0.0736   0.05843   0.05072  -0.0372   0.7784   0.7711
   3.000   0.1237   0.06029   0.05249  -0.0400   0.7632   0.7764
   3.500   0.1700   0.06213   0.05431  -0.0414   0.7472   0.7820
   4.000   0.2194   0.06420   0.05638  -0.0434   0.7310   0.7887
   4.500   0.2773   0.06652   0.05869  -0.0463   0.7141   0.7952
   5.500   0.4179   0.06512   0.05731  -0.0485   0.6393   0.8132
   6.000   0.4646   0.06605   0.05832  -0.0491   0.6170   0.8235
   6.500   0.5131   0.06681   0.05919  -0.0494   0.5979   0.8353
   7.000   0.5657   0.06740   0.05992  -0.0499   0.5812   0.8506
   7.500   0.6191   0.06763   0.06033  -0.0502   0.5653   0.8700