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OAF128 AIRFOIL (oaf128-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: OAF128 AIRFOIL (oaf128-il)
Reynolds number: 50,000
Max Cl/Cd: 24.66 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf128-il-50000.txt
Download as CSV file: xf-oaf128-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF128 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.7262   0.07824   0.07115  -0.0033   1.0000   0.1648
  -8.500  -0.7421   0.06883   0.06160  -0.0100   1.0000   0.1637
  -8.250  -0.7603   0.05838   0.05072  -0.0174   1.0000   0.1633
  -8.000  -0.7519   0.05299   0.04507  -0.0198   1.0000   0.1682
  -7.750  -0.7371   0.04891   0.04071  -0.0215   1.0000   0.1767
  -7.500  -0.7179   0.04565   0.03729  -0.0223   1.0000   0.1869
  -7.250  -0.6981   0.04194   0.03325  -0.0238   1.0000   0.2007
  -7.000  -0.6745   0.03988   0.03111  -0.0240   1.0000   0.2177
  -6.750  -0.6495   0.03871   0.02998  -0.0235   1.0000   0.2370
  -6.500  -0.6243   0.03737   0.02855  -0.0234   1.0000   0.2587
  -6.250  -0.5986   0.03595   0.02700  -0.0237   1.0000   0.2811
  -6.000  -0.5728   0.03561   0.02679  -0.0225   1.0000   0.3009
  -5.750  -0.5467   0.03498   0.02621  -0.0216   1.0000   0.3204
  -5.500  -0.5202   0.03404   0.02526  -0.0213   1.0000   0.3396
  -5.250  -0.4931   0.03296   0.02413  -0.0212   1.0000   0.3586
  -5.000  -0.4658   0.03187   0.02300  -0.0212   1.0000   0.3778
  -4.750  -0.4383   0.03083   0.02193  -0.0212   1.0000   0.3977
  -4.500  -0.4105   0.02980   0.02085  -0.0213   1.0000   0.4180
  -4.250  -0.3824   0.02881   0.01983  -0.0215   1.0000   0.4397
  -4.000  -0.3534   0.02779   0.01873  -0.0222   1.0000   0.4624
  -3.750  -0.3267   0.02708   0.01809  -0.0217   1.0000   0.4828
  -3.500  -0.3004   0.02641   0.01751  -0.0212   1.0000   0.5037
  -3.250  -0.2746   0.02578   0.01698  -0.0207   1.0000   0.5250
  -3.000  -0.2498   0.02521   0.01650  -0.0202   1.0000   0.5469
  -2.750  -0.2272   0.02470   0.01610  -0.0196   1.0000   0.5693
  -2.500  -0.2096   0.02432   0.01582  -0.0185   1.0000   0.5909
  -2.250  -0.1965   0.02412   0.01570  -0.0173   1.0000   0.6112
  -2.000  -0.1842   0.02406   0.01571  -0.0163   1.0000   0.6310
  -1.750  -0.1700   0.02412   0.01581  -0.0159   1.0000   0.6519
  -1.500  -0.1480   0.02421   0.01596  -0.0168   0.9965   0.6740
  -1.250  -0.0869   0.02404   0.01585  -0.0230   0.9733   0.7061
  -1.000  -0.0272   0.02385   0.01570  -0.0285   0.9511   0.7379
  -0.750   0.0252   0.02368   0.01556  -0.0324   0.9284   0.7702
  -0.500   0.0641   0.02361   0.01552  -0.0333   0.9054   0.8027
  -0.250   0.0895   0.02359   0.01556  -0.0316   0.8817   0.8340
   0.000   0.1134   0.02350   0.01550  -0.0294   0.8599   0.8693
   0.250   0.1404   0.02332   0.01534  -0.0281   0.8392   0.9102
   0.500   0.1912   0.02309   0.01509  -0.0311   0.8170   0.9569
   0.750   0.2638   0.02303   0.01498  -0.0399   0.7885   1.0000
   1.000   0.2752   0.02311   0.01493  -0.0402   0.7691   1.0000
   1.250   0.3034   0.02345   0.01508  -0.0417   0.7489   1.0000
   1.500   0.3331   0.02391   0.01538  -0.0425   0.7296   1.0000
   1.750   0.3619   0.02444   0.01577  -0.0428   0.7109   1.0000
   2.000   0.3901   0.02501   0.01621  -0.0428   0.6926   1.0000
   2.250   0.4179   0.02562   0.01674  -0.0427   0.6749   1.0000
   2.500   0.4454   0.02625   0.01728  -0.0424   0.6573   1.0000
   2.750   0.4725   0.02689   0.01786  -0.0420   0.6402   1.0000
   3.000   0.4995   0.02752   0.01842  -0.0414   0.6235   1.0000
   3.250   0.5260   0.02807   0.01888  -0.0404   0.6074   1.0000
   3.500   0.5531   0.02881   0.01960  -0.0400   0.5898   1.0000
   3.750   0.5799   0.02958   0.02037  -0.0396   0.5718   1.0000
   4.000   0.6065   0.03037   0.02116  -0.0392   0.5537   1.0000
   4.250   0.6327   0.03107   0.02184  -0.0385   0.5353   1.0000
   4.500   0.6587   0.03167   0.02242  -0.0375   0.5165   1.0000
   4.750   0.6845   0.03214   0.02285  -0.0363   0.4974   1.0000
   5.000   0.7103   0.03252   0.02318  -0.0350   0.4781   1.0000
   5.250   0.7363   0.03285   0.02345  -0.0337   0.4587   1.0000
   5.500   0.7620   0.03335   0.02394  -0.0327   0.4381   1.0000
   5.750   0.7870   0.03413   0.02479  -0.0321   0.4153   1.0000
   6.000   0.8125   0.03458   0.02524  -0.0310   0.3933   1.0000
   6.250   0.8385   0.03485   0.02541  -0.0296   0.3717   1.0000
   6.500   0.8649   0.03508   0.02547  -0.0281   0.3510   1.0000
   6.750   0.8889   0.03623   0.02670  -0.0276   0.3280   1.0000
   7.000   0.9131   0.03740   0.02790  -0.0270   0.3075   1.0000
   7.250   0.9371   0.03863   0.02914  -0.0264   0.2895   1.0000
   7.500   0.9603   0.04024   0.03081  -0.0259   0.2740   1.0000
   7.750   0.9837   0.04177   0.03233  -0.0254   0.2608   1.0000
   8.000   1.0088   0.04301   0.03349  -0.0247   0.2492   1.0000
   8.250   1.0213   0.04681   0.03779  -0.0252   0.2385   1.0000
   8.500   1.0439   0.04862   0.03955  -0.0246   0.2301   1.0000
   8.750   1.0438   0.05417   0.04568  -0.0257   0.2232   1.0000
   9.000   1.0770   0.05431   0.04551  -0.0243   0.2154   1.0000
   9.250   1.0524   0.06292   0.05485  -0.0264   0.2128   1.0000
   9.500   1.0044   0.07435   0.06668  -0.0309   0.2135   1.0000
   9.750   0.9501   0.08732   0.07975  -0.0380   0.2161   1.0000
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