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=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 500 AIRFOIL (goe500-il)
Reynolds number: 500,000
Max Cl/Cd: 131.78 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe500-il-500000.txt
Download as CSV file: xf-goe500-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.2086   0.10477   0.10238  -0.0677   0.9851   0.0286
 -10.000  -0.1941   0.10109   0.09869  -0.0715   0.9839   0.0297
  -9.000  -0.2688   0.02704   0.02365  -0.1595   0.9491   0.0269
  -8.750  -0.2419   0.02621   0.02283  -0.1608   0.9457   0.0279
  -8.500  -0.2130   0.02521   0.02175  -0.1625   0.9429   0.0289
  -8.250  -0.1885   0.02293   0.01914  -0.1648   0.9390   0.0300
  -8.000  -0.1671   0.02122   0.01711  -0.1654   0.9334   0.0311
  -7.750  -0.1395   0.02050   0.01612  -0.1660   0.9297   0.0321
  -7.500  -0.1151   0.01788   0.01323  -0.1674   0.9263   0.0337
  -7.250  -0.0899   0.01719   0.01246  -0.1673   0.9222   0.0348
  -7.000  -0.0647   0.01647   0.01161  -0.1672   0.9175   0.0359
  -6.750  -0.0374   0.01577   0.01076  -0.1673   0.9137   0.0373
  -6.500  -0.0090   0.01525   0.01007  -0.1675   0.9105   0.0385
  -6.250   0.0164   0.01428   0.00891  -0.1674   0.9060   0.0395
  -6.000   0.0419   0.01327   0.00781  -0.1672   0.9012   0.0410
  -5.750   0.0697   0.01281   0.00730  -0.1673   0.8973   0.0427
  -5.250   0.1238   0.01193   0.00624  -0.1668   0.8876   0.0456
  -5.000   0.1516   0.01153   0.00575  -0.1667   0.8824   0.0468
  -4.750   0.1794   0.01088   0.00501  -0.1666   0.8777   0.0492
  -4.500   0.2057   0.01049   0.00461  -0.1663   0.8708   0.0516
  -4.250   0.2338   0.01016   0.00420  -0.1661   0.8652   0.0544
  -4.000   0.2613   0.00990   0.00388  -0.1659   0.8590   0.0570
  -3.750   0.2888   0.00952   0.00351  -0.1657   0.8527   0.0638
  -3.500   0.3175   0.00928   0.00326  -0.1657   0.8479   0.0746
  -3.250   0.3447   0.00916   0.00319  -0.1655   0.8408   0.0903
  -3.000   0.3728   0.00905   0.00306  -0.1654   0.8348   0.1025
  -2.750   0.4006   0.00895   0.00292  -0.1652   0.8285   0.1108
  -2.500   0.4282   0.00883   0.00279  -0.1651   0.8214   0.1187
  -2.250   0.4562   0.00872   0.00264  -0.1649   0.8147   0.1253
  -2.000   0.4837   0.00858   0.00250  -0.1648   0.8067   0.1325
  -1.750   0.5116   0.00851   0.00237  -0.1646   0.7994   0.1379
  -1.500   0.5391   0.00837   0.00226  -0.1645   0.7913   0.1469
  -1.250   0.5668   0.00828   0.00217  -0.1644   0.7837   0.1589
  -1.000   0.5944   0.00817   0.00210  -0.1642   0.7754   0.1771
  -0.750   0.6219   0.00803   0.00207  -0.1641   0.7671   0.2187
  -0.500   0.6493   0.00793   0.00207  -0.1640   0.7587   0.2768
  -0.250   0.6765   0.00790   0.00209  -0.1637   0.7496   0.3120
   0.000   0.7039   0.00789   0.00210  -0.1635   0.7415   0.3403
   0.250   0.7309   0.00787   0.00214  -0.1633   0.7325   0.3674
   0.500   0.7581   0.00788   0.00218  -0.1631   0.7244   0.3969
   0.750   0.7850   0.00788   0.00224  -0.1628   0.7155   0.4317
   1.000   0.8118   0.00788   0.00232  -0.1625   0.7068   0.4733
   1.250   0.8383   0.00788   0.00240  -0.1622   0.6977   0.5209
   1.500   0.8646   0.00786   0.00250  -0.1618   0.6882   0.5739
   1.750   0.8906   0.00784   0.00260  -0.1614   0.6794   0.6384
   2.000   0.9137   0.00757   0.00271  -0.1603   0.6697   0.7652
   2.250   0.9430   0.00737   0.00274  -0.1604   0.6603   1.0000
   2.500   0.9693   0.00753   0.00282  -0.1600   0.6503   1.0000
   2.750   0.9954   0.00765   0.00291  -0.1596   0.6386   1.0000
   3.000   1.0212   0.00779   0.00301  -0.1591   0.6262   1.0000
   3.250   1.0466   0.00795   0.00312  -0.1585   0.6132   1.0000
   3.500   1.0714   0.00813   0.00324  -0.1578   0.5976   1.0000
   3.750   1.0958   0.00833   0.00337  -0.1571   0.5809   1.0000
   4.000   1.1199   0.00854   0.00351  -0.1563   0.5638   1.0000
   4.250   1.1432   0.00879   0.00368  -0.1554   0.5449   1.0000
   4.500   1.1666   0.00905   0.00387  -0.1545   0.5276   1.0000
   4.750   1.1899   0.00931   0.00406  -0.1536   0.5109   1.0000
   5.000   1.2130   0.00958   0.00429  -0.1527   0.4934   1.0000
   5.250   1.2356   0.00988   0.00452  -0.1517   0.4760   1.0000
   5.500   1.2577   0.01019   0.00477  -0.1506   0.4582   1.0000
   5.750   1.2797   0.01051   0.00504  -0.1495   0.4417   1.0000
   6.000   1.3017   0.01083   0.00533  -0.1485   0.4260   1.0000
   6.250   1.3219   0.01123   0.00564  -0.1471   0.4057   1.0000
   6.500   1.3421   0.01161   0.00596  -0.1457   0.3822   1.0000
   6.750   1.3595   0.01212   0.00634  -0.1439   0.3475   1.0000
   7.000   1.3676   0.01313   0.00696  -0.1405   0.2765   1.0000
   7.250   1.3668   0.01458   0.00792  -0.1358   0.2012   1.0000
   7.500   1.3631   0.01624   0.00906  -0.1307   0.1210   1.0000
   7.750   1.3616   0.01781   0.01025  -0.1262   0.0649   1.0000
   8.000   1.3706   0.01879   0.01114  -0.1233   0.0514   1.0000
   8.250   1.3841   0.01951   0.01190  -0.1211   0.0470   1.0000
   8.500   1.3937   0.02050   0.01289  -0.1185   0.0425   1.0000
   8.750   1.4071   0.02126   0.01372  -0.1164   0.0403   1.0000
   9.000   1.4195   0.02209   0.01460  -0.1143   0.0380   1.0000
   9.250   1.4286   0.02316   0.01571  -0.1119   0.0357   1.0000
   9.500   1.4329   0.02460   0.01721  -0.1090   0.0335   1.0000
   9.750   1.4459   0.02546   0.01815  -0.1072   0.0321   1.0000
  10.000   1.4560   0.02656   0.01931  -0.1052   0.0305   1.0000
  10.250   1.4642   0.02783   0.02062  -0.1031   0.0291   1.0000
  10.500   1.4655   0.02968   0.02253  -0.1004   0.0276   1.0000
  10.750   1.4682   0.03152   0.02445  -0.0980   0.0266   1.0000
  11.000   1.4793   0.03270   0.02572  -0.0965   0.0255   1.0000
  11.250   1.4883   0.03410   0.02719  -0.0949   0.0244   1.0000
  11.500   1.4962   0.03563   0.02877  -0.0934   0.0233   1.0000
  11.750   1.5008   0.03751   0.03070  -0.0918   0.0225   1.0000
  12.000   1.4954   0.04046   0.03372  -0.0895   0.0216   1.0000
  12.250   1.5020   0.04232   0.03568  -0.0882   0.0211   1.0000
  12.500   1.5092   0.04416   0.03761  -0.0871   0.0204   1.0000
  12.750   1.5155   0.04614   0.03969  -0.0860   0.0197   1.0000
  13.000   1.5220   0.04816   0.04179  -0.0851   0.0191   1.0000
  13.250   1.5279   0.05027   0.04396  -0.0843   0.0184   1.0000
  13.500   1.5325   0.05259   0.04632  -0.0836   0.0179   1.0000
  13.750   1.5320   0.05558   0.04937  -0.0826   0.0173   1.0000
  14.000   1.5344   0.05833   0.05223  -0.0817   0.0168   1.0000
  14.250   1.5397   0.06078   0.05479  -0.0813   0.0165   1.0000
  14.500   1.5441   0.06337   0.05750  -0.0808   0.0160   1.0000
  14.750   1.5479   0.06609   0.06034  -0.0804   0.0156   1.0000
  15.000   1.5512   0.06892   0.06328  -0.0802   0.0152   1.0000
  15.250   1.5542   0.07185   0.06630  -0.0801   0.0149   1.0000
  15.500   1.5568   0.07487   0.06941  -0.0802   0.0145   1.0000
  15.750   1.5587   0.07802   0.07263  -0.0804   0.0142   1.0000
  16.000   1.5600   0.08130   0.07598  -0.0806   0.0139   1.0000
  16.250   1.5595   0.08486   0.07964  -0.0802   0.0135   1.0000
  16.500   1.5580   0.08876   0.08371  -0.0808   0.0133   1.0000
  16.750   1.5562   0.09275   0.08787  -0.0817   0.0131   1.0000
  17.000   1.5538   0.09692   0.09221  -0.0827   0.0128   1.0000
  17.250   1.5506   0.10128   0.09673  -0.0839   0.0126   1.0000
  17.500   1.5467   0.10585   0.10145  -0.0853   0.0123   1.0000
  17.750   1.5420   0.11060   0.10637  -0.0869   0.0122   1.0000
  18.000   1.5365   0.11555   0.11147  -0.0888   0.0120   1.0000
  18.250   1.5308   0.12062   0.11669  -0.0909   0.0118   1.0000
  18.500   1.5237   0.12603   0.12225  -0.0934   0.0117   1.0000
  18.750   1.5161   0.13164   0.12802  -0.0962   0.0115   1.0000
  19.000   1.5077   0.13750   0.13404  -0.0993   0.0114   1.0000
  19.250   1.4985   0.14365   0.14033  -0.1027   0.0113   1.0000
<< Back to GOE 500 AIRFOIL (goe500-il)

Polar data table (+)

Polar graphs


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