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GOE 494 AIRFOIL (goe494-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 494 AIRFOIL (goe494-il)
Reynolds number: 200,000
Max Cl/Cd: 97.04 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe494-il-200000.txt
Download as CSV file: xf-goe494-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3581   0.09359   0.09028  -0.0234   1.0000   0.0289
  -7.000  -0.3688   0.09247   0.08922  -0.0205   1.0000   0.0293
  -6.750  -0.3778   0.09108   0.08790  -0.0182   1.0000   0.0297
  -6.500  -0.3830   0.08932   0.08620  -0.0171   1.0000   0.0302
  -6.250  -0.3872   0.08747   0.08440  -0.0164   1.0000   0.0308
  -6.000  -0.3765   0.08446   0.08142  -0.0202   0.9982   0.0320
  -5.750  -0.3202   0.07920   0.07607  -0.0400   0.9917   0.0339
  -5.500  -0.2747   0.07367   0.07046  -0.0529   0.9865   0.0342
  -5.250  -0.2572   0.06668   0.06348  -0.0565   0.9835   0.0356
  -5.000  -0.2398   0.06393   0.06073  -0.0567   0.9782   0.0378
  -4.750  -0.2007   0.05960   0.05632  -0.0652   0.9737   0.0430
  -4.500  -0.1289   0.05074   0.04721  -0.0862   0.9710   0.0480
  -4.250  -0.1096   0.04875   0.04521  -0.0861   0.9656   0.0509
  -4.000  -0.0441   0.04129   0.03737  -0.1019   0.9622   0.0613
  -3.750   0.0043   0.03721   0.03295  -0.1101   0.9595   0.0744
  -3.500   0.0330   0.03543   0.03128  -0.1115   0.9573   0.0797
  -3.250   0.0837   0.03164   0.02685  -0.1196   0.9531   0.1017
  -3.000   0.1100   0.02961   0.02499  -0.1204   0.9485   0.1080
  -2.750   0.1492   0.02769   0.02290  -0.1243   0.9459   0.1331
  -2.500   0.1895   0.02602   0.02106  -0.1279   0.9440   0.1594
  -2.250   0.2517   0.01876   0.01208  -0.1329   0.9439   0.0674
  -2.000   0.2850   0.01767   0.01070  -0.1339   0.9391   0.0677
  -1.750   0.3218   0.01649   0.00922  -0.1354   0.9347   0.0647
  -1.500   0.3624   0.01548   0.00804  -0.1376   0.9321   0.0628
  -1.250   0.4039   0.01462   0.00714  -0.1402   0.9302   0.0629
  -1.000   0.4337   0.01413   0.00668  -0.1405   0.9242   0.0644
  -0.750   0.4679   0.01367   0.00624  -0.1418   0.9200   0.0672
  -0.500   0.5050   0.01329   0.00586  -0.1435   0.9172   0.0718
  -0.250   0.5371   0.01299   0.00558  -0.1443   0.9126   0.0798
   0.000   0.5670   0.01276   0.00539  -0.1447   0.9068   0.0936
   0.250   0.6011   0.01187   0.00546  -0.1462   0.9035   0.4001
   0.500   0.6313   0.01063   0.00542  -0.1464   0.9000   1.0000
   0.750   0.6579   0.01076   0.00546  -0.1461   0.8928   1.0000
   1.000   0.6908   0.01073   0.00536  -0.1468   0.8889   1.0000
   1.250   0.7168   0.01088   0.00551  -0.1464   0.8819   1.0000
   1.500   0.7476   0.01083   0.00544  -0.1466   0.8760   1.0000
   1.750   0.7745   0.01082   0.00543  -0.1461   0.8670   1.0000
   2.000   0.8059   0.01063   0.00522  -0.1461   0.8595   1.0000
   2.250   0.8316   0.01059   0.00521  -0.1451   0.8484   1.0000
   2.500   0.8585   0.01036   0.00501  -0.1440   0.8340   1.0000
   2.750   0.8828   0.01000   0.00462  -0.1420   0.8094   1.0000
   3.000   0.9080   0.00984   0.00443  -0.1405   0.7867   1.0000
   3.250   0.9329   0.00979   0.00434  -0.1390   0.7601   1.0000
   3.500   0.9558   0.00985   0.00430  -0.1370   0.7194   1.0000
   3.750   0.9751   0.01005   0.00432  -0.1344   0.6480   1.0000
   4.000   0.9758   0.01189   0.00465  -0.1287   0.3725   1.0000
   4.250   0.9733   0.01540   0.00623  -0.1242   0.0642   1.0000
   4.500   0.9922   0.01663   0.00747  -0.1224   0.0477   1.0000
   4.750   1.0129   0.01757   0.00854  -0.1209   0.0430   1.0000
   5.000   1.0313   0.01881   0.00983  -0.1190   0.0391   1.0000
   5.250   1.0472   0.02078   0.01181  -0.1169   0.0350   1.0000
   5.500   1.0696   0.02207   0.01317  -0.1156   0.0336   1.0000
   5.750   1.0940   0.02374   0.01490  -0.1146   0.0325   1.0000
   6.000   1.1210   0.02577   0.01704  -0.1138   0.0323   1.0000
   6.250   1.1477   0.02769   0.01913  -0.1130   0.0310   1.0000
   6.500   1.1734   0.02976   0.02141  -0.1122   0.0296   1.0000
   6.750   1.2003   0.03327   0.02535  -0.1109   0.0321   1.0000
  12.000   0.9487   0.14182   0.13921  -0.0956   0.0477   1.0000
  12.250   0.9348   0.14922   0.14661  -0.1012   0.0473   1.0000
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