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

GOE 447 AIRFOIL (goe447-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 447 AIRFOIL (goe447-il)
Reynolds number: 100,000
Max Cl/Cd: 49.64 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe447-il-100000-n5.txt
Download as CSV file: xf-goe447-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 447 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500   0.0654   0.10250   0.09681  -0.1144   0.8723   0.0626
  -9.250   0.0751   0.10003   0.09431  -0.1157   0.8672   0.0642
  -9.000   0.0808   0.09808   0.09234  -0.1176   0.8614   0.0663
  -8.750   0.0800   0.09687   0.09114  -0.1204   0.8541   0.0672
  -8.500   0.0817   0.09511   0.08935  -0.1233   0.8487   0.0674
  -8.250   0.0841   0.09307   0.08734  -0.1246   0.8416   0.0675
  -8.000   0.1032   0.08904   0.08327  -0.1234   0.8379   0.0681
  -7.750   0.1170   0.08614   0.08034  -0.1235   0.8339   0.0687
  -7.500   0.1278   0.08370   0.07786  -0.1240   0.8299   0.0695
  -7.250   0.1331   0.08184   0.07604  -0.1241   0.8238   0.0702
  -7.000   0.1393   0.07989   0.07409  -0.1247   0.8189   0.0710
  -6.750   0.1501   0.07760   0.07175  -0.1266   0.8150   0.0722
  -6.500   0.1574   0.07577   0.06993  -0.1288   0.8093   0.0745
  -6.250   0.1680   0.07383   0.06793  -0.1371   0.8026   0.0764
  -6.000   0.1868   0.07089   0.06489  -0.1425   0.7987   0.0767
  -5.750   0.2048   0.06799   0.06190  -0.1456   0.7952   0.0767
  -5.500   0.2157   0.06574   0.05965  -0.1465   0.7895   0.0767
  -5.000   0.2435   0.05819   0.05200  -0.1470   0.7822   0.0588
  -4.750   0.2662   0.05551   0.04923  -0.1494   0.7796   0.0575
  -4.500   0.2781   0.05343   0.04712  -0.1502   0.7735   0.0565
  -4.250   0.2995   0.05076   0.04434  -0.1529   0.7692   0.0570
  -4.000   0.3274   0.04760   0.04100  -0.1567   0.7660   0.0574
  -3.750   0.3588   0.04421   0.03737  -0.1606   0.7635   0.0568
  -3.500   0.3830   0.04125   0.03420  -0.1627   0.7593   0.0562
  -3.250   0.4038   0.03850   0.03119  -0.1639   0.7539   0.0560
  -3.000   0.4343   0.03533   0.02762  -0.1663   0.7504   0.0563
  -2.750   0.4696   0.03224   0.02396  -0.1689   0.7475   0.0582
  -2.500   0.5058   0.03007   0.02130  -0.1710   0.7450   0.0605
  -2.250   0.5176   0.02979   0.02095  -0.1687   0.7364   0.0616
  -2.000   0.5493   0.02843   0.01927  -0.1695   0.7322   0.0634
  -1.750   0.5848   0.02691   0.01729  -0.1708   0.7293   0.0661
  -1.500   0.6031   0.02657   0.01680  -0.1694   0.7232   0.0687
  -1.250   0.6265   0.02635   0.01652  -0.1688   0.7185   0.0727
  -1.000   0.6570   0.02564   0.01557  -0.1692   0.7153   0.0782
  -0.750   0.6893   0.02499   0.01482  -0.1700   0.7127   0.0839
  -0.500   0.7073   0.02499   0.01476  -0.1685   0.7072   0.0901
  -0.250   0.7293   0.02493   0.01469  -0.1676   0.7021   0.0979
   0.000   0.7588   0.02453   0.01429  -0.1680   0.6987   0.1075
   0.250   0.7916   0.02407   0.01374  -0.1688   0.6960   0.1180
   0.500   0.8059   0.02436   0.01412  -0.1667   0.6895   0.1236
   0.750   0.8297   0.02432   0.01403  -0.1660   0.6847   0.1298
   1.000   0.8597   0.02399   0.01368  -0.1663   0.6815   0.1347
   1.250   0.8936   0.02361   0.01326  -0.1672   0.6789   0.1426
   1.500   0.9024   0.02425   0.01394  -0.1641   0.6709   0.1476
   1.750   0.9302   0.02410   0.01382  -0.1641   0.6666   0.1559
   2.000   0.9638   0.02375   0.01348  -0.1649   0.6634   0.1722
   2.250   0.9773   0.02418   0.01409  -0.1627   0.6564   0.2040
   2.500   1.0010   0.02414   0.01429  -0.1621   0.6512   0.3004
   2.750   1.0291   0.02310   0.01416  -0.1621   0.6479   0.6615
   3.000   1.0475   0.02316   0.01441  -0.1603   0.6418   1.0000
   3.250   1.0650   0.02356   0.01475  -0.1586   0.6352   1.0000
   3.500   1.0970   0.02343   0.01448  -0.1592   0.6313   1.0000
   3.750   1.1110   0.02396   0.01500  -0.1570   0.6241   1.0000
   4.000   1.1310   0.02424   0.01523  -0.1557   0.6178   1.0000
   4.250   1.1650   0.02405   0.01495  -0.1566   0.6140   1.0000
   4.500   1.1679   0.02494   0.01588  -0.1526   0.6054   1.0000
   4.750   1.1932   0.02505   0.01594  -0.1522   0.6001   1.0000
   5.000   1.2296   0.02477   0.01557  -0.1535   0.5962   1.0000
   5.250   1.2257   0.02592   0.01680  -0.1486   0.5858   1.0000
   5.500   1.2599   0.02560   0.01642  -0.1494   0.5810   1.0000
   5.750   1.2603   0.02671   0.01758  -0.1454   0.5710   1.0000
   6.000   1.2903   0.02653   0.01735  -0.1456   0.5654   1.0000
   6.250   1.2947   0.02758   0.01846  -0.1424   0.5561   1.0000
   6.500   1.3211   0.02753   0.01838  -0.1421   0.5497   1.0000
   6.750   1.3272   0.02855   0.01944  -0.1392   0.5403   1.0000
   7.000   1.3519   0.02860   0.01948  -0.1387   0.5337   1.0000
   7.250   1.3595   0.02964   0.02057  -0.1362   0.5249   1.0000
   7.500   1.3820   0.02980   0.02072  -0.1354   0.5179   1.0000
   7.750   1.3903   0.03085   0.02181  -0.1331   0.5089   1.0000
   8.000   1.4133   0.03095   0.02190  -0.1324   0.5015   1.0000
   8.250   1.4198   0.03215   0.02315  -0.1300   0.4921   1.0000
   8.500   1.4441   0.03220   0.02318  -0.1295   0.4853   1.0000
   8.750   1.4517   0.03346   0.02451  -0.1274   0.4776   1.0000
   9.000   1.4700   0.03399   0.02509  -0.1263   0.4714   1.0000
   9.250   1.4895   0.03448   0.02560  -0.1254   0.4656   1.0000
   9.500   1.4971   0.03579   0.02702  -0.1234   0.4582   1.0000
   9.750   1.5223   0.03585   0.02707  -0.1230   0.4527   1.0000
  10.250   1.5424   0.03817   0.02958  -0.1196   0.4387   1.0000
  10.500   1.5587   0.03887   0.03033  -0.1185   0.4323   1.0000
  10.750   1.5630   0.04049   0.03206  -0.1164   0.4244   1.0000
  11.000   1.5874   0.04053   0.03210  -0.1159   0.4180   1.0000
  11.250   1.5811   0.04298   0.03474  -0.1132   0.4089   1.0000
  11.500   1.6030   0.04314   0.03489  -0.1124   0.4013   1.0000
  11.750   1.5925   0.04602   0.03796  -0.1095   0.3914   1.0000
  12.000   1.6011   0.04725   0.03923  -0.1079   0.3820   1.0000
  12.250   1.5975   0.04961   0.04170  -0.1057   0.3708   1.0000
  12.500   1.5903   0.05246   0.04467  -0.1035   0.3588   1.0000
  12.750   1.5861   0.05506   0.04735  -0.1016   0.3458   1.0000
  13.000   1.5807   0.05788   0.05023  -0.0998   0.3319   1.0000
  13.250   1.5740   0.06099   0.05339  -0.0982   0.3171   1.0000
  13.500   1.5675   0.06418   0.05661  -0.0967   0.3017   1.0000
  13.750   1.5629   0.06723   0.05965  -0.0954   0.2865   1.0000
  14.000   1.5599   0.07013   0.06253  -0.0942   0.2724   1.0000
  14.500   1.5530   0.07628   0.06868  -0.0921   0.2476   1.0000
  14.750   1.5490   0.07952   0.07194  -0.0912   0.2367   1.0000
  15.000   1.5452   0.08277   0.07518  -0.0905   0.2262   1.0000
  15.250   1.5408   0.08615   0.07857  -0.0898   0.2163   1.0000
  15.500   1.5360   0.08976   0.08225  -0.0894   0.2070   1.0000
  15.750   1.5309   0.09336   0.08588  -0.0890   0.1975   1.0000
  16.000   1.5242   0.09731   0.08988  -0.0888   0.1877   1.0000
  16.250   1.5175   0.10137   0.09403  -0.0888   0.1779   1.0000
  16.500   1.5091   0.10569   0.09839  -0.0890   0.1672   1.0000
  16.750   1.5003   0.11024   0.10300  -0.0895   0.1555   1.0000
  17.000   1.4915   0.11494   0.10781  -0.0902   0.1410   1.0000
  17.250   1.4810   0.11993   0.11284  -0.0911   0.1066   1.0000
  17.500   1.4577   0.12683   0.11939  -0.0928   0.0849   1.0000
  17.750   1.4417   0.13274   0.12517  -0.0943   0.0729   1.0000
<< Back to GOE 447 AIRFOIL (goe447-il)

Polar data table (+)

Polar graphs


<< Back to GOE 447 AIRFOIL (goe447-il)