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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 100,000
Max Cl/Cd: 43.28 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf25-il-100000.txt
Download as CSV file: xf-raf25-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 25 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5670   0.09018   0.08530  -0.0097   1.0000   0.1016
  -8.000  -0.5820   0.08650   0.08173  -0.0155   1.0000   0.1047
  -7.750  -0.6032   0.08195   0.07705  -0.0271   1.0000   0.1061
  -7.500  -0.5726   0.07861   0.07390  -0.0162   1.0000   0.1129
  -7.250  -0.5901   0.07380   0.06883  -0.0283   1.0000   0.1194
  -7.000  -0.5680   0.07050   0.06575  -0.0210   1.0000   0.1277
  -6.750  -0.5651   0.06620   0.06142  -0.0233   1.0000   0.1375
  -6.500  -0.5620   0.06197   0.05705  -0.0265   1.0000   0.1490
  -6.250  -0.5537   0.05849   0.05347  -0.0271   1.0000   0.1625
  -6.000  -0.5430   0.05513   0.05005  -0.0268   1.0000   0.1765
  -5.750  -0.5307   0.05190   0.04682  -0.0258   1.0000   0.1912
  -5.500  -0.5173   0.04898   0.04390  -0.0243   1.0000   0.2074
  -5.250  -0.4841   0.03377   0.02649  -0.0298   1.0000   0.0759
  -5.000  -0.4632   0.02953   0.02125  -0.0278   1.0000   0.0668
  -4.750  -0.4429   0.02706   0.01876  -0.0269   1.0000   0.0715
  -4.500  -0.4202   0.02467   0.01592  -0.0255   1.0000   0.0729
  -4.250  -0.3963   0.02244   0.01328  -0.0241   1.0000   0.0740
  -4.000  -0.3719   0.02100   0.01142  -0.0228   1.0000   0.0788
  -3.750  -0.3490   0.01914   0.00958  -0.0218   1.0000   0.0866
  -3.500  -0.3247   0.01792   0.00816  -0.0205   1.0000   0.0927
  -3.250  -0.3026   0.01676   0.00709  -0.0193   1.0000   0.1060
  -3.000  -0.2812   0.01565   0.00610  -0.0179   1.0000   0.1190
  -2.750  -0.2604   0.01458   0.00521  -0.0165   1.0000   0.1473
  -2.500  -0.2499   0.01164   0.00467  -0.0133   1.0000   0.5915
  -2.250  -0.2212   0.01080   0.00460  -0.0111   1.0000   0.8571
  -2.000  -0.1281   0.01073   0.00413  -0.0239   1.0000   1.0000
  -1.750  -0.1114   0.01069   0.00389  -0.0222   1.0000   1.0000
  -1.500  -0.0939   0.01068   0.00368  -0.0205   1.0000   1.0000
  -1.250  -0.0759   0.01070   0.00356  -0.0189   1.0000   1.0000
  -1.000  -0.0577   0.01076   0.00348  -0.0173   1.0000   1.0000
  -0.750  -0.0392   0.01084   0.00345  -0.0158   1.0000   1.0000
  -0.500  -0.0207   0.01095   0.00347  -0.0144   1.0000   1.0000
  -0.250  -0.0021   0.01108   0.00351  -0.0131   1.0000   1.0000
   0.000   0.0165   0.01125   0.00361  -0.0118   1.0000   1.0000
   0.250   0.0352   0.01144   0.00376  -0.0106   1.0000   1.0000
   0.500   0.0538   0.01167   0.00394  -0.0094   1.0000   1.0000
   0.750   0.0724   0.01193   0.00418  -0.0084   1.0000   1.0000
   1.000   0.0909   0.01222   0.00447  -0.0074   1.0000   1.0000
   1.250   0.1091   0.01255   0.00481  -0.0065   1.0000   1.0000
   1.500   0.1272   0.01293   0.00521  -0.0056   1.0000   1.0000
   1.750   0.1843   0.01343   0.00582  -0.0124   0.9859   1.0000
   2.000   0.2387   0.01381   0.00636  -0.0183   0.9720   1.0000
   2.250   0.2925   0.01405   0.00679  -0.0239   0.9576   1.0000
   2.500   0.3704   0.01359   0.00666  -0.0328   0.9311   1.0000
   2.750   0.4501   0.01235   0.00581  -0.0397   0.8840   1.0000
   3.000   0.4868   0.01191   0.00551  -0.0390   0.8300   1.0000
   3.250   0.5090   0.01176   0.00530  -0.0353   0.7393   1.0000
   3.500   0.5206   0.01252   0.00506  -0.0295   0.4921   1.0000
   3.750   0.5230   0.01582   0.00611  -0.0252   0.1352   1.0000
   4.000   0.5425   0.01714   0.00730  -0.0234   0.1074   1.0000
   4.250   0.5630   0.01836   0.00841  -0.0219   0.0886   1.0000
   4.500   0.5855   0.01995   0.00997  -0.0204   0.0801   1.0000
   4.750   0.6096   0.02175   0.01168  -0.0195   0.0708   1.0000
   5.000   0.6361   0.02356   0.01372  -0.0185   0.0662   1.0000
   5.250   0.6628   0.02585   0.01624  -0.0175   0.0644   1.0000
   5.500   0.6876   0.02839   0.01898  -0.0165   0.0621   1.0000
   5.750   0.7089   0.03197   0.02290  -0.0154   0.0596   1.0000
   6.000   0.7329   0.03444   0.02624  -0.0128   0.0653   1.0000
   6.250   0.7529   0.03986   0.03188  -0.0117   0.0724   1.0000
   8.000   0.7249   0.06887   0.06452  -0.0062   0.1521   1.0000
   8.250   0.6899   0.07413   0.06981  -0.0089   0.1509   1.0000
   8.500   0.6608   0.08034   0.07594  -0.0140   0.1487   1.0000
   8.750   0.6830   0.08381   0.07943  -0.0098   0.1387   1.0000
   9.000   0.6504   0.09008   0.08562  -0.0162   0.1370   1.0000
   9.250   0.6398   0.09486   0.09036  -0.0185   0.1309   1.0000
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