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N-14 (n14-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: N-14 (n14-il)
Reynolds number: 50,000
Max Cl/Cd: 29.13 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n14-il-50000.txt
Download as CSV file: xf-n14-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-14                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5056   0.11004   0.10206  -0.0024   1.0000   0.3447
  -9.750  -0.6924   0.07822   0.07062  -0.0341   1.0000   0.1646
  -9.500  -0.7167   0.07322   0.06564  -0.0330   1.0000   0.1608
  -9.250  -0.8022   0.06685   0.05891  -0.0278   1.0000   0.1524
  -9.000  -0.8053   0.06272   0.05466  -0.0256   1.0000   0.1512
  -8.750  -0.8124   0.05879   0.05053  -0.0227   1.0000   0.1501
  -8.500  -0.8196   0.05504   0.04650  -0.0194   1.0000   0.1494
  -8.250  -0.8257   0.05158   0.04268  -0.0158   1.0000   0.1499
  -8.000  -0.8295   0.04842   0.03908  -0.0120   1.0000   0.1512
  -7.750  -0.8296   0.04548   0.03567  -0.0082   1.0000   0.1529
  -7.500  -0.8128   0.04283   0.03303  -0.0068   1.0000   0.1568
  -7.250  -0.8014   0.04033   0.03026  -0.0043   1.0000   0.1600
  -7.000  -0.7908   0.03796   0.02749  -0.0015   1.0000   0.1646
  -6.750  -0.7756   0.03588   0.02527   0.0006   1.0000   0.1719
  -6.500  -0.7603   0.03400   0.02308   0.0029   1.0000   0.1809
  -6.250  -0.7397   0.03208   0.02115   0.0044   1.0000   0.1919
  -6.000  -0.7217   0.03042   0.01942   0.0063   1.0000   0.2091
  -5.750  -0.7033   0.02882   0.01797   0.0083   1.0000   0.2339
  -5.500  -0.6864   0.02738   0.01667   0.0105   1.0000   0.2665
  -5.250  -0.6708   0.02621   0.01560   0.0129   1.0000   0.3022
  -5.000  -0.6551   0.02514   0.01469   0.0152   1.0000   0.3398
  -4.750  -0.6425   0.02420   0.01403   0.0181   1.0000   0.3857
  -4.500  -0.6271   0.02307   0.01341   0.0209   1.0000   0.4439
  -4.250  -0.6101   0.02181   0.01282   0.0239   1.0000   0.5306
  -4.000  -0.5898   0.02136   0.01314   0.0280   1.0000   0.6465
  -3.750  -0.5561   0.02227   0.01426   0.0310   1.0000   0.7476
  -3.500  -0.5024   0.02388   0.01562   0.0305   1.0000   0.8163
  -3.250  -0.3855   0.02634   0.01747   0.0193   1.0000   0.8691
  -3.000  -0.2477   0.02777   0.01829   0.0026   1.0000   0.9234
  -2.750  -0.1251   0.02721   0.01726  -0.0150   1.0000   0.9708
  -2.500  -0.0346   0.02560   0.01535  -0.0290   1.0000   1.0000
  -2.250  -0.0287   0.02505   0.01476  -0.0271   1.0000   1.0000
  -2.000  -0.0231   0.02459   0.01428  -0.0249   1.0000   1.0000
  -1.750  -0.0183   0.02423   0.01390  -0.0224   1.0000   1.0000
  -1.500  -0.0144   0.02395   0.01361  -0.0197   1.0000   1.0000
  -1.250  -0.0115   0.02374   0.01339  -0.0168   1.0000   1.0000
  -1.000  -0.0095   0.02359   0.01323  -0.0136   1.0000   1.0000
  -0.750  -0.0082   0.02348   0.01312  -0.0103   1.0000   1.0000
  -0.500  -0.0072   0.02342   0.01304  -0.0068   1.0000   1.0000
  -0.250  -0.0062   0.02339   0.01300  -0.0034   1.0000   1.0000
   0.000  -0.0053   0.02338   0.01299   0.0000   1.0000   1.0000
   0.250  -0.0041   0.02341   0.01301   0.0034   1.0000   1.0000
   0.500  -0.0026   0.02346   0.01306   0.0068   1.0000   1.0000
   0.750  -0.0007   0.02356   0.01315   0.0101   1.0000   1.0000
   1.000   0.0018   0.02369   0.01328   0.0132   1.0000   1.0000
   1.250   0.0050   0.02388   0.01347   0.0162   1.0000   1.0000
   1.500   0.0092   0.02413   0.01372   0.0190   1.0000   1.0000
   1.750   0.0142   0.02445   0.01404   0.0215   1.0000   1.0000
   2.000   0.0198   0.02485   0.01445   0.0238   1.0000   1.0000
   2.250   0.0259   0.02534   0.01496   0.0259   1.0000   1.0000
   2.500   0.0684   0.02639   0.01615   0.0207   0.9877   1.0000
   2.750   0.1606   0.02779   0.01784   0.0069   0.9565   1.0000
   3.000   0.2427   0.02839   0.01874  -0.0040   0.9202   1.0000
   3.250   0.3594   0.02800   0.01886  -0.0189   0.8786   1.0000
   3.500   0.4930   0.02544   0.01691  -0.0330   0.8222   1.0000
   3.750   0.5683   0.02345   0.01523  -0.0370   0.7655   1.0000
   4.000   0.6044   0.02248   0.01427  -0.0351   0.7038   1.0000
   4.250   0.6284   0.02207   0.01369  -0.0317   0.6383   1.0000
   4.500   0.6458   0.02217   0.01344  -0.0278   0.5689   1.0000
   4.750   0.6587   0.02264   0.01345  -0.0236   0.5002   1.0000
   5.000   0.6707   0.02347   0.01384  -0.0199   0.4367   1.0000
   5.250   0.6840   0.02464   0.01463  -0.0167   0.3768   1.0000
   5.500   0.6982   0.02603   0.01562  -0.0139   0.3215   1.0000
   5.750   0.7126   0.02748   0.01676  -0.0111   0.2745   1.0000
   6.000   0.7281   0.02886   0.01804  -0.0085   0.2407   1.0000
   6.250   0.7459   0.03023   0.01937  -0.0065   0.2177   1.0000
   6.500   0.7660   0.03185   0.02102  -0.0048   0.2023   1.0000
   6.750   0.7838   0.03351   0.02280  -0.0029   0.1897   1.0000
   7.000   0.8022   0.03541   0.02477  -0.0012   0.1805   1.0000
   7.250   0.8177   0.03745   0.02708   0.0010   0.1733   1.0000
   7.500   0.8326   0.03965   0.02939   0.0030   0.1666   1.0000
   7.750   0.8394   0.04208   0.03229   0.0062   0.1616   1.0000
   8.000   0.8489   0.04466   0.03514   0.0089   0.1580   1.0000
   8.250   0.8641   0.04738   0.03781   0.0105   0.1533   1.0000
   8.500   0.8598   0.05045   0.04139   0.0146   0.1517   1.0000
   8.750   0.8521   0.05381   0.04518   0.0185   0.1503   1.0000
   9.000   0.8427   0.05752   0.04921   0.0222   0.1503   1.0000
   9.250   0.8299   0.06147   0.05342   0.0256   0.1509   1.0000
   9.500   0.8151   0.06555   0.05769   0.0287   0.1517   1.0000
   9.750   0.8005   0.06977   0.06203   0.0312   0.1526   1.0000
  10.000   0.7870   0.07414   0.06647   0.0332   0.1534   1.0000
  10.250   0.6379   0.09027   0.08256   0.0260   0.1790   1.0000
  10.500   0.5969   0.10311   0.09526   0.0172   0.2112   1.0000
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