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GOE 494 AIRFOIL (goe494-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 494 AIRFOIL (goe494-il)
Reynolds number: 500,000
Max Cl/Cd: 121.92 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe494-il-500000-n5.txt
Download as CSV file: xf-goe494-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 494 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3316   0.09708   0.09484  -0.0347   0.9974   0.0054
  -8.250  -0.3189   0.09294   0.09071  -0.0384   0.9942   0.0054
  -8.000  -0.3059   0.08912   0.08690  -0.0420   0.9905   0.0054
  -7.750  -0.2919   0.08466   0.08244  -0.0466   0.9867   0.0055
  -7.500  -0.2806   0.08035   0.07816  -0.0507   0.9800   0.0057
  -7.250  -0.2645   0.07605   0.07384  -0.0558   0.9741   0.0057
  -7.000  -0.2439   0.07317   0.07095  -0.0605   0.9696   0.0062
  -6.750  -0.2186   0.06962   0.06739  -0.0667   0.9666   0.0070
  -6.500  -0.1950   0.06479   0.06254  -0.0740   0.9595   0.0073
  -6.250  -0.1616   0.05852   0.05623  -0.0849   0.9550   0.0075
  -6.000  -0.1276   0.05156   0.04921  -0.0966   0.9471   0.0078
  -5.750  -0.0768   0.04108   0.03858  -0.1151   0.9407   0.0086
  -5.500  -0.0166   0.02557   0.02252  -0.1369   0.9322   0.0098
  -5.250   0.0263   0.01420   0.00972  -0.1459   0.9274   0.0141
  -5.000   0.0535   0.01500   0.01061  -0.1457   0.9217   0.0150
  -4.750   0.0817   0.01522   0.01080  -0.1458   0.9164   0.0168
  -4.500   0.1109   0.01444   0.00976  -0.1463   0.9112   0.0204
  -4.250   0.1390   0.01407   0.00916  -0.1464   0.9050   0.0226
  -3.750   0.1950   0.01248   0.00721  -0.1472   0.8923   0.0261
  -3.500   0.2228   0.01216   0.00679  -0.1473   0.8860   0.0281
  -3.250   0.2503   0.01168   0.00616  -0.1473   0.8791   0.0296
  -3.000   0.2782   0.01112   0.00544  -0.1474   0.8725   0.0303
  -2.750   0.3056   0.01060   0.00479  -0.1473   0.8642   0.0306
  -2.500   0.3330   0.01024   0.00431  -0.1472   0.8553   0.0314
  -2.250   0.3603   0.00981   0.00377  -0.1470   0.8453   0.0314
  -2.000   0.3876   0.00938   0.00325  -0.1469   0.8349   0.0310
  -1.750   0.4148   0.00903   0.00281  -0.1467   0.8242   0.0306
  -1.500   0.4420   0.00874   0.00243  -0.1465   0.8132   0.0303
  -1.250   0.4692   0.00851   0.00212  -0.1463   0.8030   0.0300
  -1.000   0.4966   0.00830   0.00187  -0.1462   0.7939   0.0298
  -0.750   0.5239   0.00814   0.00165  -0.1460   0.7861   0.0296
  -0.500   0.5513   0.00802   0.00148  -0.1459   0.7785   0.0295
  -0.250   0.5787   0.00792   0.00134  -0.1458   0.7717   0.0294
   0.000   0.6061   0.00785   0.00124  -0.1456   0.7651   0.0295
   0.250   0.6335   0.00782   0.00116  -0.1455   0.7594   0.0298
   0.500   0.6607   0.00780   0.00112  -0.1453   0.7519   0.0303
   0.750   0.6876   0.00781   0.00109  -0.1450   0.7433   0.0314
   1.000   0.7145   0.00784   0.00109  -0.1447   0.7343   0.0334
   1.250   0.7415   0.00779   0.00114  -0.1445   0.7256   0.0713
   1.500   0.7679   0.00755   0.00128  -0.1444   0.7122   0.2461
   1.750   0.7934   0.00756   0.00141  -0.1439   0.6919   0.3127
   2.000   0.8191   0.00755   0.00155  -0.1435   0.6725   0.3889
   2.250   0.8446   0.00723   0.00177  -0.1432   0.6545   0.6122
   2.750   0.8876   0.00728   0.00201  -0.1404   0.5366   1.0000
   3.000   0.9077   0.00801   0.00235  -0.1390   0.4541   1.0000
   3.250   0.9216   0.00955   0.00299  -0.1368   0.2698   1.0000
   3.500   0.9410   0.01055   0.00350  -0.1356   0.1691   1.0000
   3.750   0.9592   0.01172   0.00410  -0.1342   0.0537   1.0000
   4.000   0.9826   0.01224   0.00456  -0.1334   0.0315   1.0000
   4.250   1.0068   0.01264   0.00496  -0.1327   0.0210   1.0000
   4.500   1.0309   0.01307   0.00537  -0.1320   0.0146   1.0000
   4.750   1.0541   0.01364   0.00602  -0.1310   0.0115   1.0000
   5.000   1.0775   0.01415   0.00662  -0.1302   0.0101   1.0000
   5.250   1.1009   0.01462   0.00713  -0.1294   0.0085   1.0000
   5.500   1.1210   0.01558   0.00816  -0.1280   0.0071   1.0000
   5.750   1.1428   0.01627   0.00893  -0.1269   0.0067   1.0000
   6.000   1.1636   0.01712   0.00988  -0.1256   0.0062   1.0000
   6.250   1.1836   0.01809   0.01094  -0.1242   0.0057   1.0000
   6.500   1.2036   0.01907   0.01205  -0.1228   0.0054   1.0000
   6.750   1.2249   0.01978   0.01279  -0.1218   0.0049   1.0000
   7.000   1.2432   0.02115   0.01424  -0.1203   0.0044   1.0000
   7.250   1.2633   0.02236   0.01561  -0.1190   0.0041   1.0000
   7.500   1.2831   0.02401   0.01745  -0.1175   0.0039   1.0000
   7.750   1.3031   0.02603   0.01971  -0.1161   0.0036   1.0000
   8.000   1.3226   0.02847   0.02244  -0.1146   0.0034   1.0000
   8.250   1.3404   0.03136   0.02566  -0.1128   0.0032   1.0000
   8.500   1.3546   0.03506   0.02976  -0.1106   0.0031   1.0000
   8.750   1.3622   0.04057   0.03580  -0.1072   0.0031   1.0000
   9.000   1.3616   0.04707   0.04284  -0.1029   0.0031   1.0000
   9.250   1.3561   0.05281   0.04905  -0.0987   0.0032   1.0000
   9.500   1.3468   0.05789   0.05448  -0.0946   0.0033   1.0000
   9.750   1.3330   0.06231   0.05917  -0.0903   0.0033   1.0000
  10.000   1.3144   0.06591   0.06297  -0.0858   0.0034   1.0000
  10.250   1.2938   0.06988   0.06709  -0.0823   0.0034   1.0000
  10.500   1.2727   0.07434   0.07173  -0.0802   0.0034   1.0000
  10.750   1.2509   0.07945   0.07701  -0.0795   0.0034   1.0000
  11.000   1.2297   0.08508   0.08280  -0.0803   0.0034   1.0000
  11.250   1.2062   0.09199   0.08985  -0.0827   0.0035   1.0000
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