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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 1,000,000
Max Cl/Cd: 68.14 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf25-il-1000000-n5.txt
Download as CSV file: xf-raf25-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6163   0.08311   0.08154  -0.0097   1.0000   0.0026
  -8.750  -0.7915   0.02629   0.02345  -0.0333   1.0000   0.0020
  -8.500  -0.7810   0.02296   0.01971  -0.0317   1.0000   0.0020
  -8.250  -0.7722   0.01917   0.01536  -0.0296   1.0000   0.0021
  -8.000  -0.7541   0.01761   0.01357  -0.0283   1.0000   0.0023
  -7.750  -0.7333   0.01669   0.01252  -0.0274   1.0000   0.0024
  -7.500  -0.7125   0.01574   0.01142  -0.0263   1.0000   0.0025
  -7.250  -0.6906   0.01504   0.01061  -0.0254   1.0000   0.0027
  -7.000  -0.6693   0.01418   0.00960  -0.0243   1.0000   0.0029
  -6.750  -0.6480   0.01326   0.00849  -0.0232   1.0000   0.0030
  -6.500  -0.6263   0.01245   0.00753  -0.0220   1.0000   0.0032
  -6.250  -0.6041   0.01180   0.00676  -0.0210   1.0000   0.0034
  -6.000  -0.5749   0.01119   0.00603  -0.0215   0.9990   0.0036
  -5.750  -0.5448   0.01040   0.00512  -0.0222   0.9973   0.0041
  -5.500  -0.5136   0.00994   0.00461  -0.0232   0.9956   0.0045
  -5.250  -0.4832   0.00952   0.00413  -0.0239   0.9936   0.0050
  -5.000  -0.4524   0.00922   0.00377  -0.0247   0.9910   0.0057
  -4.750  -0.4212   0.00874   0.00322  -0.0255   0.9883   0.0070
  -4.500  -0.3898   0.00842   0.00286  -0.0264   0.9850   0.0084
  -4.250  -0.3588   0.00813   0.00253  -0.0271   0.9798   0.0101
  -4.000  -0.3279   0.00784   0.00228  -0.0278   0.9738   0.0167
  -3.750  -0.2955   0.00764   0.00210  -0.0289   0.9675   0.0253
  -3.500  -0.2622   0.00753   0.00197  -0.0302   0.9607   0.0274
  -3.250  -0.2263   0.00738   0.00181  -0.0321   0.9530   0.0284
  -3.000  -0.1909   0.00716   0.00153  -0.0338   0.9412   0.0319
  -2.750  -0.1564   0.00699   0.00134  -0.0353   0.9266   0.0346
  -2.500  -0.1258   0.00689   0.00117  -0.0359   0.9090   0.0362
  -2.250  -0.0979   0.00682   0.00103  -0.0358   0.8924   0.0406
  -2.000  -0.0713   0.00667   0.00091  -0.0355   0.8769   0.0621
  -1.750  -0.0456   0.00648   0.00081  -0.0351   0.8608   0.1037
  -1.500  -0.0201   0.00634   0.00072  -0.0346   0.8415   0.1464
  -1.250   0.0050   0.00622   0.00064  -0.0340   0.8182   0.1849
  -1.000   0.0297   0.00598   0.00057  -0.0335   0.7980   0.2618
  -0.750   0.0550   0.00578   0.00053  -0.0330   0.7819   0.3272
  -0.500   0.0799   0.00562   0.00049  -0.0325   0.7605   0.3883
  -0.250   0.1036   0.00554   0.00047  -0.0316   0.7190   0.4574
   0.000   0.1265   0.00560   0.00046  -0.0306   0.6565   0.5118
   0.250   0.1506   0.00560   0.00048  -0.0298   0.6098   0.5648
   0.500   0.1745   0.00564   0.00051  -0.0291   0.5609   0.6172
   0.750   0.1978   0.00563   0.00056  -0.0282   0.5145   0.6880
   1.000   0.2174   0.00544   0.00063  -0.0264   0.4656   0.8101
   1.250   0.2386   0.00553   0.00074  -0.0249   0.3836   0.9012
   1.500   0.2891   0.00597   0.00092  -0.0303   0.2856   0.9646
   1.750   0.3253   0.00661   0.00110  -0.0327   0.1598   0.9814
   2.000   0.3577   0.00696   0.00126  -0.0338   0.1086   0.9920
   2.250   0.3915   0.00732   0.00143  -0.0354   0.0543   0.9984
   2.500   0.4217   0.00748   0.00157  -0.0360   0.0490   1.0000
   2.750   0.4463   0.00766   0.00174  -0.0353   0.0422   1.0000
   3.000   0.4713   0.00778   0.00187  -0.0347   0.0399   1.0000
   3.250   0.4960   0.00796   0.00203  -0.0341   0.0350   1.0000
   3.500   0.5209   0.00812   0.00219  -0.0334   0.0317   1.0000
   3.750   0.5454   0.00833   0.00236  -0.0328   0.0216   1.0000
   4.000   0.5698   0.00858   0.00257  -0.0321   0.0122   1.0000
   4.250   0.5941   0.00885   0.00284  -0.0313   0.0096   1.0000
   4.500   0.6180   0.00919   0.00321  -0.0305   0.0071   1.0000
   4.750   0.6423   0.00946   0.00353  -0.0298   0.0062   1.0000
   5.000   0.6664   0.00978   0.00388  -0.0290   0.0055   1.0000
   5.250   0.6898   0.01019   0.00433  -0.0282   0.0048   1.0000
   5.500   0.7126   0.01074   0.00495  -0.0271   0.0043   1.0000
   5.750   0.7360   0.01118   0.00547  -0.0263   0.0040   1.0000
   6.000   0.7595   0.01161   0.00598  -0.0254   0.0036   1.0000
   6.250   0.7830   0.01204   0.00646  -0.0247   0.0033   1.0000
   6.500   0.8064   0.01250   0.00696  -0.0239   0.0030   1.0000
   6.750   0.8291   0.01309   0.00761  -0.0230   0.0029   1.0000
   7.000   0.8488   0.01428   0.00896  -0.0216   0.0026   1.0000
   7.250   0.8713   0.01497   0.00976  -0.0207   0.0025   1.0000
   7.500   0.8930   0.01585   0.01078  -0.0196   0.0024   1.0000
   7.750   0.9145   0.01678   0.01187  -0.0186   0.0023   1.0000
   8.000   0.9351   0.01795   0.01323  -0.0174   0.0021   1.0000
   8.250   0.9539   0.01949   0.01501  -0.0161   0.0020   1.0000
   8.500   0.9715   0.02128   0.01707  -0.0146   0.0019   1.0000
   8.750   0.9872   0.02342   0.01955  -0.0129   0.0018   1.0000
   9.000   0.9963   0.02696   0.02354  -0.0105   0.0018   1.0000
  10.500   0.7808   0.09566   0.09411  -0.0204   0.0020   1.0000
  10.750   0.7721   0.10246   0.10090  -0.0240   0.0021   1.0000
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