XFOIL         Version 6.94
  
 Calculated polar for: GOE 464 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.8827   0.01834   0.01271  -0.1564  0.5475  0.0176
   0.500   0.9376   0.01743   0.01160  -0.1560  0.5270  0.0276
   1.000   0.9900   0.01712   0.01125  -0.1552  0.5047  0.0558
   1.500   1.0440   0.01638   0.01018  -0.1545  0.4816  0.0675
   2.500   1.1206   0.00781   0.00162  -0.1476  0.4328  0.0937
   3.000   1.1734   0.00792   0.00147  -0.1467  0.4141  0.0965
   3.500   1.2214   0.00765   0.00107  -0.1459  0.3986  0.0972
   4.000   1.2700   0.00752   0.00083  -0.1449  0.3845  0.0986
   4.500   1.3204   0.00768   0.00081  -0.1436  0.3727  0.1021
   5.000   1.3661   0.00745   0.00052  -0.1424  0.3603  0.1033
   5.500   1.4099   0.00762   0.00044  -0.1402  0.3314  0.1083
   6.000   1.4457   0.00773   0.00033  -0.1374  0.2781  0.1101
   6.500   1.4730   0.00863   0.00073  -0.1330  0.2128  0.1158
   7.000   1.4935   0.00948   0.00134  -0.1276  0.1667  0.1195
   7.500   1.4960   0.01100   0.00251  -0.1195  0.1144  0.1257
   8.000   1.4971   0.01284   0.00409  -0.1122  0.0701  0.1341
   8.500   1.4701   0.01691   0.00797  -0.1029  0.0028  0.1386
   9.000   1.4870   0.01873   0.00994  -0.0993  0.0028  0.1551