XFOIL         Version 6.94
  
 Calculated polar for: GOE 570 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1600   0.02951   0.02067  -0.0212  0.4930  0.1951
   0.500   0.2099   0.02963   0.02074  -0.0216  0.4882  0.2092
   1.500   0.2143   0.03401   0.02527  -0.0113  0.4740  0.2287
   2.500   0.3087   0.03418   0.02641  -0.0133  0.4649  0.4532
   3.000   0.3790   0.03339   0.02577  -0.0161  0.4611  0.5283
   3.500   0.2382   0.04393   0.03668   0.0030  0.4462  0.5201
   4.000   0.3006   0.04306   0.03600   0.0012  0.4438  0.5889