XFOIL         Version 6.94
  
 Calculated polar for: WORTMANN FX 77-W-120 AIRFOIL                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0063   0.04596   0.03740  -0.0612  0.7251  0.1628
  -2.750   0.0180   0.04398   0.03525  -0.0609  0.7206  0.1592
  -2.500   0.0429   0.04280   0.03376  -0.0615  0.7157  0.1551
  -2.250   0.0689   0.04185   0.03237  -0.0619  0.7108  0.1502
  -2.000   0.0970   0.04132   0.03124  -0.0617  0.7065  0.1462
  -1.750   0.1255   0.04037   0.02993  -0.0614  0.7026  0.1463
  -1.500   0.1495   0.03993   0.02953  -0.0625  0.6984  0.1533
  -1.250   0.1760   0.04005   0.02934  -0.0633  0.6941  0.1628
  -1.000   0.2076   0.03978   0.02894  -0.0645  0.6902  0.1795
  -0.750   0.2392   0.03927   0.02845  -0.0649  0.6866  0.1998
  -0.500   0.2621   0.03936   0.02856  -0.0648  0.6832  0.2303
  -0.250   0.2756   0.03965   0.02939  -0.0649  0.6805  0.2933
   0.000   0.3620   0.03859   0.02939  -0.0750  0.6753  1.0000
   0.250   0.3713   0.04042   0.03099  -0.0746  0.6746  1.0000
   0.500   0.3788   0.04229   0.03264  -0.0739  0.6737  1.0000
   0.750   0.3860   0.04415   0.03430  -0.0731  0.6724  1.0000
   1.000   0.3896   0.04616   0.03616  -0.0721  0.6720  1.0000
   1.250   0.3814   0.04863   0.03853  -0.0706  0.6756  1.0000
   1.500   0.3791   0.05081   0.04057  -0.0691  0.6792  1.0000
   1.750   0.3843   0.05284   0.04245  -0.0682  0.6818  1.0000
   2.000   0.3927   0.05493   0.04441  -0.0678  0.6856  1.0000