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FX 61-168 AIRFOIL (fx61168-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 61-168 AIRFOIL (fx61168-il)
Reynolds number: 50,000
Max Cl/Cd: 5.07 at α=10.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx61168-il-50000.txt
Download as CSV file: xf-fx61168-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 61-168 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2908   0.10692   0.10095  -0.0297   1.0000   0.3520
  -8.750  -0.4406   0.08774   0.08235  -0.0492   1.0000   0.1709
  -8.500  -0.4710   0.08534   0.08012  -0.0463   1.0000   0.1595
  -8.250  -0.5165   0.08511   0.08009  -0.0417   1.0000   0.1589
  -8.000  -0.5568   0.08449   0.07960  -0.0377   1.0000   0.1597
  -7.750  -0.5962   0.08257   0.07777  -0.0358   1.0000   0.1594
  -7.500  -0.6378   0.07929   0.07443  -0.0353   1.0000   0.1573
  -7.250  -0.6598   0.07496   0.06992  -0.0353   1.0000   0.1463
  -7.000  -0.6803   0.06993   0.06446  -0.0361   1.0000   0.1343
  -6.750  -0.6878   0.06484   0.05870  -0.0370   1.0000   0.1217
  -6.500  -0.6681   0.05971   0.05243  -0.0410   0.9955   0.1088
  -6.250  -0.6399   0.05537   0.04791  -0.0433   0.9899   0.1067
  -6.000  -0.6091   0.05178   0.04377  -0.0461   0.9842   0.1058
  -5.750  -0.5796   0.04885   0.04039  -0.0479   0.9785   0.1051
  -5.500  -0.5452   0.04622   0.03730  -0.0498   0.9732   0.1042
  -5.250  -0.5162   0.04418   0.03487  -0.0503   0.9675   0.1046
  -5.000  -0.4787   0.04275   0.03293  -0.0520   0.9620   0.1086
  -4.750  -0.4576   0.04126   0.03158  -0.0507   0.9567   0.1141
  -4.500  -0.4300   0.04048   0.03071  -0.0496   0.9515   0.1205
  -4.250  -0.4047   0.03988   0.03022  -0.0481   0.9467   0.1308
  -4.000  -0.3879   0.03917   0.02961  -0.0461   0.9415   0.1450
  -3.750  -0.3626   0.03800   0.02867  -0.0464   0.9363   0.1725
  -3.500  -0.3539   0.03523   0.02894  -0.0439   0.9324   0.4808
  -3.250  -0.3710   0.04064   0.03439  -0.0274   0.9279   0.6678
  -3.000  -0.1804   0.05232   0.04496  -0.0195   0.9194   0.8691
  -2.750  -0.1798   0.05187   0.04435  -0.0172   0.9138   0.8802
  -2.500  -0.1686   0.05149   0.04379  -0.0167   0.9084   0.8916
  -2.250  -0.1261   0.05102   0.04300  -0.0215   0.9026   0.9038
  -2.000  -0.1042   0.05065   0.04245  -0.0229   0.8969   0.9139
  -1.750  -0.0718   0.05040   0.04197  -0.0261   0.8912   0.9238
  -1.500  -0.0596   0.05033   0.04176  -0.0260   0.8860   0.9324
  -1.250  -0.0343   0.05021   0.04149  -0.0281   0.8804   0.9404
  -1.000   0.0094   0.05021   0.04126  -0.0332   0.8742   0.9495
  -0.750   0.0187   0.05032   0.04129  -0.0328   0.8685   0.9568
  -0.500   0.0564   0.05045   0.04126  -0.0371   0.8621   0.9651
  -0.250   0.0809   0.05076   0.04145  -0.0391   0.8560   0.9724
   0.000   0.1075   0.05107   0.04167  -0.0418   0.8497   0.9787
   0.250   0.1521   0.05149   0.04195  -0.0471   0.8430   0.9853
   0.500   0.1603   0.05207   0.04250  -0.0469   0.8375   0.9905
   0.750   0.1962   0.05264   0.04298  -0.0509   0.8309   0.9960
   1.000   0.2137   0.05342   0.04370  -0.0521   0.8253   1.0000
   1.250   0.2025   0.05420   0.04448  -0.0484   0.8222   1.0000
   1.500   0.1974   0.05494   0.04521  -0.0456   0.8192   1.0000
   1.750   0.1969   0.05571   0.04595  -0.0436   0.8165   1.0000
   2.000   0.2061   0.05657   0.04677  -0.0428   0.8124   1.0000
   2.250   0.1992   0.05749   0.04768  -0.0400   0.8130   1.0000
   2.500  -0.0455   0.05494   0.04557  -0.0037   0.9745   1.0000
   2.750  -0.0282   0.05570   0.04627  -0.0050   0.9629   1.0000
   3.000  -0.0124   0.05643   0.04694  -0.0059   0.9503   1.0000
   3.250   0.0021   0.05714   0.04761  -0.0065   0.9373   1.0000
   3.500   0.0154   0.05785   0.04827  -0.0069   0.9240   1.0000
   3.750   0.0296   0.05867   0.04904  -0.0073   0.9110   1.0000
   4.000   0.0473   0.05981   0.05013  -0.0084   0.8991   1.0000
   4.250   0.0736   0.06170   0.05198  -0.0111   0.8878   1.0000
   4.500   0.1051   0.06390   0.05414  -0.0146   0.8745   1.0000
   4.750   0.1253   0.06502   0.05525  -0.0162   0.8605   1.0000
   5.000   0.1459   0.06636   0.05657  -0.0179   0.8470   1.0000
   5.250   0.1682   0.06798   0.05820  -0.0199   0.8337   1.0000
   5.500   0.1919   0.06986   0.06007  -0.0222   0.8213   1.0000
   5.750   0.2199   0.07219   0.06241  -0.0252   0.8097   1.0000
   6.000   0.2603   0.07555   0.06579  -0.0300   0.7977   1.0000
   6.250   0.2781   0.07690   0.06717  -0.0313   0.7836   1.0000
   6.500   0.2957   0.07853   0.06885  -0.0328   0.7701   1.0000
   6.750   0.3146   0.08050   0.07086  -0.0345   0.7576   1.0000
   7.000   0.3370   0.08283   0.07324  -0.0368   0.7453   1.0000
   7.250   0.3661   0.08575   0.07622  -0.0398   0.7339   1.0000
   7.500   0.4033   0.08923   0.07976  -0.0437   0.7206   1.0000
   7.750   0.4195   0.09104   0.08164  -0.0450   0.7061   1.0000
   8.000   0.4333   0.09293   0.08361  -0.0462   0.6923   1.0000
   8.250   0.4480   0.09514   0.08591  -0.0475   0.6791   1.0000
   8.500   0.4645   0.09770   0.08855  -0.0492   0.6670   1.0000
   8.750   0.4866   0.10070   0.09163  -0.0514   0.6553   1.0000
   9.000   0.5190   0.10440   0.09543  -0.0546   0.6430   1.0000
   9.250   0.5417   0.10731   0.09847  -0.0566   0.6287   1.0000
   9.500   0.5502   0.10937   0.10062  -0.0575   0.6147   1.0000
   9.750   0.5579   0.11179   0.10313  -0.0585   0.6019   1.0000
  10.000   0.5704   0.11468   0.10612  -0.0600   0.5898   1.0000
  10.250   0.5916   0.11830   0.10986  -0.0622   0.5795   1.0000
  10.500   0.6185   0.12204   0.11373  -0.0646   0.5662   1.0000
  10.750   0.6305   0.12472   0.11651  -0.0659   0.5529   1.0000
  11.000   0.6290   0.12721   0.11909  -0.0668   0.5425   1.0000
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