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RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 6 AIRFOIL (raf6-il)
Reynolds number: 50,000
Max Cl/Cd: 35.74 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf6-il-50000.txt
Download as CSV file: xf-raf6-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 6 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4220   0.10541   0.09915  -0.0219   1.0000   0.1898
  -7.500  -0.4001   0.10044   0.09418  -0.0187   1.0000   0.2024
  -7.250  -0.4298   0.09990   0.09382  -0.0200   1.0000   0.2055
  -7.000  -0.4148   0.09568   0.08964  -0.0160   1.0000   0.2186
  -6.750  -0.4147   0.09261   0.08666  -0.0139   1.0000   0.2282
  -6.500  -0.4396   0.09171   0.08586  -0.0162   1.0000   0.2369
  -6.250  -0.4363   0.08856   0.08279  -0.0132   1.0000   0.2520
  -6.000  -0.4385   0.08591   0.08021  -0.0112   1.0000   0.2677
  -5.750  -0.4421   0.08326   0.07764  -0.0096   1.0000   0.2847
  -5.500  -0.4319   0.07999   0.07446  -0.0037   1.0000   0.3069
  -5.250  -0.4382   0.07774   0.07227  -0.0015   1.0000   0.3333
  -5.000   0.0298   0.05195   0.04524  -0.0307   1.0000   1.0000
  -4.750   0.0358   0.05010   0.04347  -0.0305   1.0000   1.0000
  -4.500   0.0415   0.04832   0.04176  -0.0301   1.0000   1.0000
  -4.250  -0.0020   0.04958   0.04325  -0.0178   1.0000   0.9768
  -4.000  -0.0629   0.05170   0.04564  -0.0026   1.0000   0.9410
  -3.750  -0.1177   0.05259   0.04679   0.0095   1.0000   0.8975
  -3.500  -0.1717   0.05329   0.04772   0.0205   1.0000   0.8642
  -3.250  -0.2229   0.05338   0.04803   0.0300   1.0000   0.8332
  -2.500  -0.1947   0.04034   0.03200  -0.0336   1.0000   0.2034
  -2.250  -0.1585   0.03829   0.02916  -0.0344   1.0000   0.1736
  -2.000  -0.1275   0.03641   0.02677  -0.0345   1.0000   0.1576
  -1.750  -0.0990   0.03457   0.02462  -0.0344   1.0000   0.1490
  -1.500  -0.0703   0.03360   0.02311  -0.0340   1.0000   0.1444
  -1.250  -0.0439   0.03240   0.02173  -0.0338   1.0000   0.1458
  -1.000  -0.0176   0.03156   0.02064  -0.0333   1.0000   0.1455
  -0.750   0.0077   0.03079   0.01975  -0.0327   1.0000   0.1472
  -0.500   0.0332   0.03032   0.01917  -0.0321   1.0000   0.1549
  -0.250   0.0597   0.02993   0.01869  -0.0318   1.0000   0.1619
   0.000   0.0852   0.02971   0.01841  -0.0316   1.0000   0.1735
   0.250   0.1118   0.02946   0.01828  -0.0319   1.0000   0.1959
   0.750   0.1536   0.02729   0.01843  -0.0297   1.0000   1.0000
   1.000   0.1726   0.02817   0.01886  -0.0291   1.0000   1.0000
   1.250   0.2475   0.02973   0.01988  -0.0391   0.9749   1.0000
   1.500   0.3189   0.03057   0.02043  -0.0477   0.9454   1.0000
   1.750   0.3798   0.03082   0.02052  -0.0539   0.9163   1.0000
   2.000   0.4378   0.03064   0.02026  -0.0588   0.8871   1.0000
   2.250   0.4956   0.03003   0.01961  -0.0630   0.8574   1.0000
   2.500   0.5549   0.02892   0.01852  -0.0666   0.8274   1.0000
   2.750   0.6181   0.02720   0.01684  -0.0700   0.7971   1.0000
   3.000   0.6799   0.02529   0.01497  -0.0725   0.7603   1.0000
   3.250   0.7329   0.02366   0.01324  -0.0736   0.7118   1.0000
   3.500   0.7726   0.02288   0.01218  -0.0731   0.6540   1.0000
   3.750   0.8076   0.02292   0.01182  -0.0728   0.6008   1.0000
   4.000   0.8399   0.02350   0.01200  -0.0727   0.5587   1.0000
   4.250   0.8687   0.02438   0.01261  -0.0724   0.5262   1.0000
   4.500   0.8987   0.02533   0.01334  -0.0726   0.5010   1.0000
   4.750   0.9237   0.02638   0.01432  -0.0721   0.4797   1.0000
   5.000   0.9498   0.02747   0.01535  -0.0718   0.4619   1.0000
   5.250   0.9751   0.02861   0.01647  -0.0714   0.4466   1.0000
   5.500   0.9996   0.02980   0.01769  -0.0710   0.4327   1.0000
   5.750   1.0244   0.03104   0.01894  -0.0706   0.4205   1.0000
   6.000   1.0498   0.03231   0.02020  -0.0703   0.4093   1.0000
   6.250   1.0682   0.03372   0.02182  -0.0690   0.3984   1.0000
   6.500   1.0892   0.03519   0.02344  -0.0681   0.3888   1.0000
   6.750   1.1117   0.03664   0.02497  -0.0674   0.3797   1.0000
   7.000   1.1260   0.03843   0.02704  -0.0657   0.3714   1.0000
   7.250   1.1514   0.03986   0.02846  -0.0654   0.3629   1.0000
   7.500   1.1581   0.04208   0.03112  -0.0629   0.3559   1.0000
   7.750   1.1790   0.04382   0.03296  -0.0622   0.3490   1.0000
   8.000   1.1855   0.04632   0.03579  -0.0598   0.3430   1.0000
   8.250   1.1937   0.04869   0.03843  -0.0578   0.3365   1.0000
   8.500   1.2128   0.05077   0.04061  -0.0569   0.3308   1.0000
   8.750   1.1990   0.05474   0.04502  -0.0532   0.3275   1.0000
   9.000   1.1838   0.05901   0.04963  -0.0498   0.3246   1.0000
   9.250   1.1649   0.06367   0.05453  -0.0467   0.3221   1.0000
   9.500   1.1508   0.06817   0.05919  -0.0443   0.3195   1.0000
   9.750   1.1514   0.07196   0.06308  -0.0429   0.3162   1.0000
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