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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 500,000
Max Cl/Cd: 55.7 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf25-il-500000-n5.txt
Download as CSV file: xf-raf25-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5947   0.08387   0.08165  -0.0106   1.0000   0.0050
  -8.750  -0.5993   0.07902   0.07684  -0.0134   1.0000   0.0048
  -8.500  -0.6068   0.07404   0.07189  -0.0166   1.0000   0.0048
  -8.250  -0.6173   0.06751   0.06539  -0.0227   1.0000   0.0047
  -8.000  -0.6217   0.05991   0.05772  -0.0288   1.0000   0.0046
  -7.750  -0.6263   0.05179   0.04944  -0.0324   1.0000   0.0045
  -7.500  -0.6340   0.04187   0.03917  -0.0338   1.0000   0.0043
  -7.250  -0.6565   0.02569   0.02182  -0.0308   1.0000   0.0041
  -7.000  -0.6441   0.02189   0.01749  -0.0289   1.0000   0.0042
  -6.750  -0.6268   0.01956   0.01478  -0.0274   1.0000   0.0044
  -6.500  -0.6078   0.01770   0.01258  -0.0259   1.0000   0.0047
  -6.250  -0.5877   0.01619   0.01079  -0.0246   1.0000   0.0050
  -6.000  -0.5668   0.01493   0.00930  -0.0233   1.0000   0.0053
  -5.750  -0.5449   0.01408   0.00827  -0.0222   1.0000   0.0057
  -5.500  -0.5241   0.01299   0.00704  -0.0209   1.0000   0.0065
  -5.250  -0.5015   0.01251   0.00649  -0.0200   1.0000   0.0072
  -5.000  -0.4793   0.01187   0.00573  -0.0189   1.0000   0.0080
  -4.750  -0.4571   0.01127   0.00501  -0.0178   1.0000   0.0088
  -4.500  -0.4260   0.01052   0.00413  -0.0186   0.9978   0.0108
  -4.250  -0.3931   0.01011   0.00372  -0.0198   0.9951   0.0148
  -4.000  -0.3611   0.00975   0.00335  -0.0209   0.9917   0.0212
  -3.750  -0.3274   0.00965   0.00319  -0.0222   0.9882   0.0275
  -3.500  -0.2950   0.00932   0.00283  -0.0234   0.9834   0.0322
  -3.250  -0.2616   0.00909   0.00257  -0.0247   0.9783   0.0360
  -3.000  -0.2306   0.00888   0.00231  -0.0254   0.9713   0.0385
  -2.500  -0.1666   0.00838   0.00177  -0.0273   0.9581   0.0491
  -2.250  -0.1329   0.00804   0.00156  -0.0287   0.9515   0.0842
  -2.000  -0.1002   0.00765   0.00137  -0.0300   0.9426   0.1488
  -1.750  -0.0667   0.00723   0.00121  -0.0314   0.9332   0.2239
  -1.500  -0.0333   0.00687   0.00109  -0.0329   0.9226   0.3038
  -1.250  -0.0025   0.00645   0.00098  -0.0337   0.9098   0.4036
  -1.000   0.0262   0.00615   0.00093  -0.0340   0.8952   0.4917
  -0.750   0.0533   0.00597   0.00089  -0.0338   0.8765   0.5528
  -0.500   0.0784   0.00577   0.00086  -0.0331   0.8546   0.6232
  -0.250   0.1015   0.00554   0.00086  -0.0319   0.8346   0.7063
   0.000   0.1228   0.00521   0.00089  -0.0302   0.8190   0.8150
   0.250   0.1599   0.00506   0.00094  -0.0317   0.7915   0.9214
   0.500   0.2080   0.00523   0.00093  -0.0361   0.7385   0.9612
   0.750   0.2428   0.00543   0.00094  -0.0375   0.6897   0.9790
   1.000   0.2745   0.00568   0.00097  -0.0384   0.6323   0.9910
   1.250   0.3078   0.00607   0.00101  -0.0398   0.5472   1.0000
   1.500   0.3298   0.00641   0.00108  -0.0386   0.4815   1.0000
   1.750   0.3515   0.00682   0.00120  -0.0375   0.3999   1.0000
   2.000   0.3731   0.00732   0.00135  -0.0364   0.3123   1.0000
   2.250   0.3948   0.00784   0.00153  -0.0354   0.2213   1.0000
   2.500   0.4158   0.00849   0.00179  -0.0343   0.1219   1.0000
   2.750   0.4383   0.00899   0.00204  -0.0333   0.0615   1.0000
   3.000   0.4628   0.00921   0.00227  -0.0326   0.0537   1.0000
   3.250   0.4872   0.00945   0.00252  -0.0320   0.0475   1.0000
   3.500   0.5116   0.00970   0.00278  -0.0313   0.0411   1.0000
   3.750   0.5362   0.00992   0.00302  -0.0306   0.0348   1.0000
   4.000   0.5609   0.01015   0.00326  -0.0300   0.0269   1.0000
   4.250   0.5848   0.01050   0.00354  -0.0292   0.0162   1.0000
   4.500   0.6083   0.01096   0.00406  -0.0283   0.0117   1.0000
   4.750   0.6319   0.01141   0.00461  -0.0274   0.0100   1.0000
   5.000   0.6552   0.01188   0.00513  -0.0265   0.0084   1.0000
   5.250   0.6767   0.01270   0.00605  -0.0253   0.0070   1.0000
   5.500   0.6993   0.01338   0.00684  -0.0242   0.0065   1.0000
   5.750   0.7214   0.01417   0.00774  -0.0231   0.0060   1.0000
   6.000   0.7433   0.01508   0.00877  -0.0219   0.0055   1.0000
   6.250   0.7652   0.01605   0.00990  -0.0208   0.0052   1.0000
   6.500   0.7871   0.01705   0.01102  -0.0198   0.0049   1.0000
   6.750   0.8065   0.01872   0.01287  -0.0185   0.0044   1.0000
   7.000   0.8291   0.01967   0.01402  -0.0176   0.0041   1.0000
   7.250   0.8488   0.02157   0.01623  -0.0162   0.0038   1.0000
   7.500   0.8658   0.02418   0.01925  -0.0144   0.0036   1.0000
   7.750   0.8787   0.02782   0.02339  -0.0122   0.0034   1.0000
   8.000   0.8801   0.03466   0.03091  -0.0086   0.0032   1.0000
   8.250   0.8709   0.04357   0.04042  -0.0048   0.0033   1.0000
   8.500   0.8639   0.05038   0.04759  -0.0025   0.0033   1.0000
   8.750   0.8538   0.05659   0.05406  -0.0010   0.0033   1.0000
   9.000   0.8419   0.06193   0.05958  -0.0002   0.0034   1.0000
   9.250   0.8218   0.06676   0.06452   0.0007   0.0034   1.0000
   9.500   0.8016   0.07235   0.07020  -0.0018   0.0035   1.0000
   9.750   0.7847   0.08156   0.07944  -0.0107   0.0035   1.0000
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