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 =      .250 e 6     Ncrit =   7.000
  
   alpha     CL        CD       CDp       CM    Top Xtr Bot Xtr
  ------- -------- --------- --------- -------- ------- -------
  -2.000    .1739    .01093    .00319   -.0808   .7209   .0976
  -1.500    .2275    .01006    .00293   -.0810   .7067   .3019
  -1.000    .2746    .00883    .00293   -.0793   .6936   .6902
   -.500    .3371    .00832    .00271   -.0795   .6813  1.0000
    .000    .3925    .00848    .00264   -.0795   .6682  1.0000
    .500    .4479    .00868    .00263   -.0794   .6558  1.0000
   1.000    .5032    .00891    .00268   -.0793   .6446  1.0000
   1.500    .5584    .00913    .00277   -.0792   .6330  1.0000
   2.000    .6136    .00936    .00295   -.0791   .6209  1.0000
   2.500    .6686    .00961    .00316   -.0790   .6092  1.0000
   3.000    .7233    .00987    .00336   -.0788   .5966  1.0000
   3.500    .7777    .01011    .00360   -.0785   .5818  1.0000
   4.500    .8852    .01055    .00420   -.0778   .5463  1.0000
   5.000    .9384    .01077    .00454   -.0773   .5241  1.0000
   5.500    .9901    .01100    .00485   -.0765   .4897  1.0000
   6.000   1.0396    .01142    .00520   -.0754   .4362  1.0000
   7.000   1.1131    .01509    .00755   -.0708   .1677  1.0000
   7.500   1.1433    .01766    .00948   -.0680   .0613  1.0000
   8.000   1.1727    .02011    .01174   -.0648   .0218  1.0000
   8.500   1.2022    .02223    .01406   -.0614   .0183  1.0000
   9.000   1.2268    .02426    .01633   -.0576   .0164  1.0000
   9.500   1.2417    .02669    .01896   -.0532   .0150  1.0000
  10.000   1.2453    .03002    .02251   -.0487   .0143  1.0000