XFOIL         Version 6.8
  
 Calculated polar for: e387                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =    .000     Re =      .500 e 6     Ncrit =   7.000
  
   alpha     CL        CD       CDp       CM    Top Xtr Bot Xtr
  ------- -------- --------- --------- -------- ------- -------
  -2.000    .1714    .00939    .00210   -.0791   .6658   .0192
  -1.500    .2269    .00866    .00176   -.0793   .6534   .1303
  -1.000    .2820    .00811    .00163   -.0795   .6417   .2888
   -.500    .3345    .00721    .00163   -.0794   .6310   .6094
    .000    .3892    .00635    .00164   -.0783   .6197   .9839
    .500    .4476    .00647    .00160   -.0789   .6091  1.0000
   1.000    .5032    .00664    .00162   -.0788   .5987  1.0000
   1.500    .5593    .00677    .00168   -.0789   .5875  1.0000
   2.000    .6149    .00693    .00179   -.0789   .5765  1.0000
   2.500    .6703    .00712    .00190   -.0788   .5636  1.0000
   3.000    .7254    .00731    .00203   -.0787   .5486  1.0000
   3.500    .7806    .00749    .00222   -.0786   .5325  1.0000
   4.000    .8354    .00770    .00243   -.0785   .5147  1.0000
   4.500    .8898    .00793    .00268   -.0783   .4886  1.0000
   5.000    .9425    .00830    .00295   -.0779   .4488  1.0000
   5.500    .9918    .00906    .00341   -.0771   .3675  1.0000
   6.000   1.0316    .01101    .00452   -.0754   .2091  1.0000
   6.500   1.0733    .01267    .00569   -.0738   .1102  1.0000
   7.000   1.1104    .01475    .00722   -.0717   .0223  1.0000
   7.500   1.1534    .01605    .00853   -.0700   .0128  1.0000
   8.000   1.1958    .01725    .00990   -.0683   .0110  1.0000
   8.500   1.2336    .01876    .01157   -.0660   .0101  1.0000
   9.000   1.2623    .02087    .01389   -.0626   .0093  1.0000
   9.500   1.2768    .02340    .01667   -.0575   .0090  1.0000
  10.000   1.2891    .02587    .01934   -.0528   .0088  1.0000