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GOE 599 AIRFOIL (goe599-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 599 AIRFOIL (goe599-il)
Reynolds number: 50,000
Max Cl/Cd: 34.17 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe599-il-50000.txt
Download as CSV file: xf-goe599-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 599 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5091   0.09793   0.09091  -0.0061   1.0000   0.3419
  -8.000  -0.4847   0.09255   0.08545  -0.0056   1.0000   0.3473
  -7.750  -0.5965   0.07246   0.06561  -0.0358   1.0000   0.1645
  -7.500  -0.6027   0.06523   0.05823  -0.0380   1.0000   0.1475
  -7.250  -0.6136   0.05794   0.05051  -0.0399   1.0000   0.1352
  -7.000  -0.6095   0.05276   0.04502  -0.0396   1.0000   0.1301
  -6.750  -0.6090   0.04708   0.03821  -0.0392   1.0000   0.1237
  -6.500  -0.5929   0.04374   0.03496  -0.0382   1.0000   0.1276
  -6.250  -0.5777   0.04048   0.03129  -0.0370   1.0000   0.1304
  -6.000  -0.5600   0.03715   0.02744  -0.0359   1.0000   0.1319
  -5.750  -0.5403   0.03428   0.02397  -0.0347   1.0000   0.1362
  -5.500  -0.5194   0.03197   0.02153  -0.0337   1.0000   0.1447
  -5.250  -0.4960   0.02970   0.01877  -0.0325   1.0000   0.1527
  -5.000  -0.4732   0.02791   0.01695  -0.0315   1.0000   0.1672
  -4.750  -0.4499   0.02624   0.01535  -0.0303   1.0000   0.1859
  -4.500  -0.4267   0.02466   0.01387  -0.0291   1.0000   0.2195
  -4.250  -0.4044   0.02272   0.01242  -0.0277   1.0000   0.2982
  -4.000  -0.3873   0.02110   0.01167  -0.0253   1.0000   0.4298
  -3.750  -0.3750   0.02058   0.01182  -0.0211   1.0000   0.5512
  -3.500  -0.3629   0.02028   0.01187  -0.0161   1.0000   0.6441
  -3.250  -0.3495   0.01992   0.01170  -0.0113   1.0000   0.7184
  -3.000  -0.3363   0.01963   0.01151  -0.0064   1.0000   0.7883
  -2.750  -0.3163   0.01952   0.01142  -0.0023   1.0000   0.8651
  -2.500  -0.1414   0.02038   0.01128  -0.0266   1.0000   0.9873
  -2.250  -0.1219   0.01958   0.01030  -0.0288   1.0000   1.0000
  -2.000  -0.1385   0.01895   0.00962  -0.0237   1.0000   1.0000
  -1.750  -0.1406   0.01864   0.00914  -0.0202   1.0000   1.0000
  -1.500  -0.1303   0.01857   0.00885  -0.0184   1.0000   1.0000
  -1.250  -0.1149   0.01863   0.00869  -0.0173   1.0000   1.0000
  -1.000  -0.0972   0.01875   0.00861  -0.0165   1.0000   1.0000
  -0.750  -0.0783   0.01893   0.00862  -0.0158   1.0000   1.0000
  -0.500  -0.0587   0.01916   0.00869  -0.0152   1.0000   1.0000
  -0.250  -0.0387   0.01943   0.00883  -0.0148   1.0000   1.0000
   0.000  -0.0185   0.01973   0.00902  -0.0143   1.0000   1.0000
   0.250   0.0018   0.02008   0.00927  -0.0140   1.0000   1.0000
   0.500   0.0221   0.02047   0.00959  -0.0137   1.0000   1.0000
   0.750   0.0424   0.02091   0.00997  -0.0135   1.0000   1.0000
   1.000   0.0625   0.02139   0.01040  -0.0133   1.0000   1.0000
   1.250   0.0824   0.02193   0.01091  -0.0131   1.0000   1.0000
   1.500   0.1020   0.02252   0.01149  -0.0130   1.0000   1.0000
   1.750   0.1213   0.02317   0.01215  -0.0130   1.0000   1.0000
   2.000   0.1402   0.02390   0.01290  -0.0131   1.0000   1.0000
   2.250   0.1585   0.02470   0.01374  -0.0132   1.0000   1.0000
   2.500   0.1763   0.02559   0.01468  -0.0134   1.0000   1.0000
   2.750   0.2355   0.02704   0.01627  -0.0213   0.9832   1.0000
   3.000   0.2885   0.02820   0.01759  -0.0278   0.9615   1.0000
   3.250   0.3456   0.02932   0.01891  -0.0347   0.9415   1.0000
   3.500   0.3895   0.03014   0.01996  -0.0388   0.9184   1.0000
   3.750   0.4398   0.03090   0.02097  -0.0437   0.8974   1.0000
   4.000   0.4954   0.03095   0.02135  -0.0481   0.8678   1.0000
   4.250   0.5831   0.02848   0.01940  -0.0533   0.8232   1.0000
   4.500   0.6514   0.02520   0.01659  -0.0532   0.7774   1.0000
   4.750   0.6914   0.02294   0.01458  -0.0495   0.7261   1.0000
   5.000   0.7142   0.02164   0.01344  -0.0443   0.6610   1.0000
   5.250   0.7233   0.02117   0.01215  -0.0357   0.4635   1.0000
   5.500   0.7228   0.02388   0.01313  -0.0302   0.2713   1.0000
   5.750   0.7402   0.02571   0.01439  -0.0284   0.2245   1.0000
   6.000   0.7672   0.02733   0.01585  -0.0278   0.2021   1.0000
   6.250   0.8008   0.02912   0.01755  -0.0280   0.1873   1.0000
   6.500   0.8335   0.03112   0.01953  -0.0284   0.1728   1.0000
   6.750   0.8663   0.03351   0.02208  -0.0285   0.1642   1.0000
   7.000   0.8950   0.03624   0.02483  -0.0285   0.1537   1.0000
   7.250   0.9177   0.03872   0.02786  -0.0272   0.1470   1.0000
   7.500   0.9407   0.04199   0.03126  -0.0265   0.1388   1.0000
   7.750   0.9565   0.04507   0.03501  -0.0244   0.1354   1.0000
   8.000   0.9705   0.04832   0.03872  -0.0224   0.1307   1.0000
   8.250   0.9900   0.05251   0.04294  -0.0217   0.1263   1.0000
   8.500   0.9911   0.05626   0.04752  -0.0185   0.1282   1.0000
   8.750   0.9868   0.06089   0.05282  -0.0157   0.1309   1.0000
   9.000   0.9789   0.06585   0.05825  -0.0134   0.1336   1.0000
   9.250   0.9709   0.07092   0.06363  -0.0118   0.1361   1.0000
   9.500   0.9664   0.07617   0.06906  -0.0106   0.1384   1.0000
   9.750   0.9162   0.08277   0.07604  -0.0100   0.1505   1.0000
  10.000   0.8636   0.09171   0.08506  -0.0144   0.1639   1.0000
  10.250   0.8005   0.10975   0.10292  -0.0314   0.2074   1.0000
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