Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 500 AIRFOIL (goe500-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 500 AIRFOIL (goe500-il)
Reynolds number: 200,000
Max Cl/Cd: 91.74 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe500-il-200000-n5.txt
Download as CSV file: xf-goe500-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 500 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2119   0.10115   0.09740  -0.0698   0.9762   0.0241
  -9.750  -0.2017   0.09641   0.09265  -0.0741   0.9740   0.0246
  -9.500  -0.1967   0.09196   0.08820  -0.0772   0.9689   0.0250
  -9.000  -0.1759   0.08249   0.07872  -0.0868   0.9627   0.0273
  -8.750  -0.1617   0.08002   0.07625  -0.0893   0.9584   0.0282
  -8.500  -0.1513   0.07598   0.07220  -0.0933   0.9533   0.0289
  -8.000  -0.1940   0.02973   0.02487  -0.1533   0.9217   0.0326
  -7.750  -0.1649   0.02683   0.02165  -0.1573   0.9186   0.0340
  -7.500  -0.1376   0.02555   0.02020  -0.1588   0.9149   0.0352
  -7.250  -0.1141   0.02435   0.01878  -0.1594   0.9092   0.0368
  -7.000  -0.0853   0.02268   0.01675  -0.1611   0.9055   0.0386
  -6.750  -0.0546   0.02111   0.01479  -0.1627   0.9027   0.0400
  -6.500  -0.0286   0.01991   0.01334  -0.1631   0.8981   0.0416
  -6.250  -0.0031   0.01908   0.01240  -0.1632   0.8930   0.0430
  -6.000   0.0260   0.01828   0.01147  -0.1638   0.8894   0.0444
  -5.750   0.0568   0.01749   0.01051  -0.1646   0.8865   0.0459
  -5.500   0.0814   0.01692   0.00979  -0.1642   0.8806   0.0476
  -5.250   0.1092   0.01635   0.00906  -0.1642   0.8760   0.0496
  -5.000   0.1391   0.01567   0.00827  -0.1647   0.8725   0.0515
  -4.750   0.1662   0.01517   0.00773  -0.1647   0.8676   0.0536
  -4.500   0.1929   0.01478   0.00727  -0.1645   0.8619   0.0562
  -4.250   0.2225   0.01435   0.00675  -0.1648   0.8579   0.0595
  -4.000   0.2500   0.01399   0.00637  -0.1647   0.8524   0.0640
  -3.750   0.2773   0.01372   0.00606  -0.1645   0.8457   0.0711
  -3.500   0.3071   0.01341   0.00574  -0.1647   0.8399   0.0802
  -3.250   0.3331   0.01320   0.00550  -0.1642   0.8305   0.0899
  -3.000   0.3617   0.01300   0.00519  -0.1641   0.8232   0.0999
  -2.750   0.3889   0.01281   0.00498  -0.1639   0.8154   0.1091
  -2.500   0.4176   0.01262   0.00473  -0.1639   0.8095   0.1181
  -2.250   0.4446   0.01248   0.00458  -0.1637   0.8018   0.1275
  -2.000   0.4737   0.01227   0.00436  -0.1638   0.7959   0.1379
  -1.750   0.5005   0.01215   0.00423  -0.1636   0.7877   0.1487
  -1.250   0.5565   0.01185   0.00397  -0.1635   0.7733   0.1737
  -1.000   0.5855   0.01169   0.00382  -0.1635   0.7665   0.1940
  -0.750   0.6123   0.01159   0.00381  -0.1633   0.7578   0.2265
  -0.500   0.6408   0.01149   0.00374  -0.1633   0.7505   0.2619
  -0.250   0.6678   0.01145   0.00375  -0.1631   0.7420   0.2922
   0.000   0.6958   0.01141   0.00372  -0.1630   0.7347   0.3198
   0.250   0.7227   0.01140   0.00375  -0.1627   0.7258   0.3459
   0.500   0.7501   0.01138   0.00377  -0.1625   0.7176   0.3742
   0.750   0.7769   0.01137   0.00382  -0.1622   0.7087   0.4070
   1.000   0.8038   0.01138   0.00390  -0.1619   0.7001   0.4449
   1.250   0.8305   0.01137   0.00399  -0.1616   0.6917   0.4922
   1.500   0.8564   0.01137   0.00410  -0.1611   0.6828   0.5394
   1.750   0.8824   0.01135   0.00419  -0.1606   0.6742   0.5954
   2.000   0.9075   0.01130   0.00430  -0.1600   0.6642   0.6639
   2.500   0.9605   0.01105   0.00441  -0.1590   0.6452   1.0000
   2.750   0.9865   0.01121   0.00454  -0.1586   0.6355   1.0000
   3.000   1.0127   0.01139   0.00466  -0.1582   0.6260   1.0000
   3.250   1.0382   0.01158   0.00483  -0.1577   0.6153   1.0000
   3.500   1.0636   0.01177   0.00500  -0.1572   0.6046   1.0000
   3.750   1.0891   0.01198   0.00517  -0.1566   0.5946   1.0000
   4.000   1.1141   0.01220   0.00537  -0.1560   0.5834   1.0000
   4.250   1.1384   0.01242   0.00558  -0.1553   0.5709   1.0000
   4.500   1.1623   0.01267   0.00580  -0.1545   0.5572   1.0000
   4.750   1.1853   0.01294   0.00603  -0.1535   0.5414   1.0000
   5.000   1.2075   0.01324   0.00628  -0.1524   0.5243   1.0000
   5.250   1.2292   0.01356   0.00654  -0.1512   0.5071   1.0000
   5.500   1.2506   0.01391   0.00684  -0.1500   0.4905   1.0000
   5.750   1.2716   0.01428   0.00717  -0.1488   0.4745   1.0000
   6.000   1.2916   0.01467   0.00752  -0.1473   0.4566   1.0000
   6.250   1.3108   0.01510   0.00790  -0.1458   0.4375   1.0000
   6.500   1.3280   0.01559   0.00833  -0.1439   0.4154   1.0000
   6.750   1.3452   0.01608   0.00877  -0.1420   0.3952   1.0000
   7.000   1.3622   0.01656   0.00924  -0.1402   0.3727   1.0000
   7.250   1.3756   0.01713   0.00974  -0.1377   0.3432   1.0000
   7.500   1.3833   0.01789   0.01034  -0.1343   0.3027   1.0000
   7.750   1.3834   0.01910   0.01123  -0.1299   0.2504   1.0000
   8.000   1.3842   0.02048   0.01235  -0.1260   0.2054   1.0000
   8.250   1.3847   0.02201   0.01361  -0.1222   0.1568   1.0000
   8.500   1.3837   0.02372   0.01504  -0.1185   0.1121   1.0000
   8.750   1.3855   0.02535   0.01648  -0.1153   0.0801   1.0000
   9.000   1.3902   0.02683   0.01786  -0.1126   0.0638   1.0000
   9.250   1.3970   0.02821   0.01923  -0.1102   0.0558   1.0000
   9.500   1.4055   0.02950   0.02058  -0.1081   0.0509   1.0000
   9.750   1.4129   0.03090   0.02205  -0.1060   0.0470   1.0000
  10.000   1.4175   0.03257   0.02377  -0.1037   0.0437   1.0000
  10.250   1.4261   0.03396   0.02527  -0.1019   0.0412   1.0000
  10.500   1.4329   0.03554   0.02695  -0.1001   0.0387   1.0000
  10.750   1.4370   0.03741   0.02889  -0.0983   0.0365   1.0000
  11.000   1.4384   0.03960   0.03115  -0.0964   0.0345   1.0000
  11.250   1.4451   0.04136   0.03305  -0.0949   0.0327   1.0000
  11.500   1.4500   0.04335   0.03517  -0.0935   0.0309   1.0000
  11.750   1.4536   0.04553   0.03743  -0.0922   0.0293   1.0000
  12.000   1.4535   0.04817   0.04012  -0.0908   0.0279   1.0000
  12.250   1.4564   0.05059   0.04267  -0.0897   0.0265   1.0000
  12.500   1.4610   0.05288   0.04509  -0.0888   0.0251   1.0000
  12.750   1.4644   0.05538   0.04770  -0.0880   0.0238   1.0000
  13.000   1.4672   0.05801   0.05042  -0.0873   0.0228   1.0000
  13.250   1.4683   0.06091   0.05339  -0.0867   0.0218   1.0000
  13.500   1.4691   0.06392   0.05650  -0.0861   0.0210   1.0000
  13.750   1.4727   0.06665   0.05939  -0.0857   0.0200   1.0000
  14.000   1.4758   0.06951   0.06238  -0.0854   0.0190   1.0000
  14.250   1.4783   0.07248   0.06547  -0.0853   0.0182   1.0000
  14.500   1.4801   0.07561   0.06869  -0.0853   0.0175   1.0000
  14.750   1.4808   0.07894   0.07211  -0.0854   0.0170   1.0000
  15.000   1.4801   0.08250   0.07575  -0.0855   0.0165   1.0000
  15.250   1.4817   0.08582   0.07923  -0.0856   0.0161   1.0000
  15.500   1.4830   0.08922   0.08280  -0.0858   0.0155   1.0000
  15.750   1.4837   0.09275   0.08651  -0.0862   0.0150   1.0000
  16.000   1.4839   0.09643   0.09033  -0.0868   0.0145   1.0000
  16.250   1.4835   0.10025   0.09429  -0.0877   0.0140   1.0000
  16.500   1.4826   0.10423   0.09839  -0.0888   0.0136   1.0000
  16.750   1.4811   0.10834   0.10260  -0.0901   0.0133   1.0000
  17.000   1.4790   0.11256   0.10693  -0.0915   0.0130   1.0000
  17.250   1.4764   0.11687   0.11133  -0.0929   0.0127   1.0000
  17.500   1.4730   0.12144   0.11609  -0.0945   0.0125   1.0000
  17.750   1.4686   0.12633   0.12120  -0.0964   0.0122   1.0000
  18.000   1.4630   0.13154   0.12662  -0.0987   0.0120   1.0000
  18.250   1.4563   0.13706   0.13235  -0.1013   0.0118   1.0000
  18.500   1.4482   0.14299   0.13850  -0.1045   0.0117   1.0000
  18.750   1.4389   0.14931   0.14503  -0.1080   0.0115   1.0000
  19.000   1.4281   0.15617   0.15210  -0.1122   0.0114   1.0000
  19.250   1.4159   0.16364   0.15977  -0.1169   0.0113   1.0000
<< Back to GOE 500 AIRFOIL (goe500-il)

Polar data table (+)

Polar graphs


<< Back to GOE 500 AIRFOIL (goe500-il)