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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 200,000
Max Cl/Cd: 47.82 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf25-il-200000-n5.txt
Download as CSV file: xf-raf25-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5785   0.09800   0.09440  -0.0052   1.0000   0.0107
  -9.250  -0.5835   0.09178   0.08824  -0.0082   1.0000   0.0101
  -9.000  -0.5854   0.08714   0.08364  -0.0105   1.0000   0.0101
  -8.750  -0.5881   0.08180   0.07836  -0.0134   1.0000   0.0097
  -8.500  -0.5899   0.07714   0.07375  -0.0162   1.0000   0.0095
  -8.250  -0.5985   0.07162   0.06829  -0.0204   1.0000   0.0095
  -8.000  -0.5962   0.06612   0.06277  -0.0252   1.0000   0.0093
  -7.750  -0.5968   0.06002   0.05656  -0.0289   1.0000   0.0092
  -7.500  -0.5979   0.05362   0.05001  -0.0315   1.0000   0.0091
  -7.250  -0.5964   0.04749   0.04364  -0.0327   1.0000   0.0091
  -7.000  -0.5927   0.04167   0.03748  -0.0328   1.0000   0.0091
  -6.500  -0.5787   0.03090   0.02569  -0.0309   1.0000   0.0099
  -6.250  -0.5628   0.02863   0.02312  -0.0299   1.0000   0.0108
  -6.000  -0.5459   0.02573   0.01978  -0.0285   1.0000   0.0113
  -5.750  -0.5274   0.02292   0.01648  -0.0270   1.0000   0.0118
  -5.500  -0.5071   0.02068   0.01384  -0.0255   1.0000   0.0125
  -5.250  -0.4857   0.01881   0.01164  -0.0242   1.0000   0.0135
  -5.000  -0.4633   0.01767   0.01022  -0.0230   1.0000   0.0157
  -4.750  -0.4416   0.01627   0.00862  -0.0217   1.0000   0.0174
  -4.500  -0.4204   0.01508   0.00734  -0.0205   1.0000   0.0201
  -4.250  -0.3978   0.01445   0.00662  -0.0194   1.0000   0.0252
  -4.000  -0.3760   0.01364   0.00571  -0.0183   1.0000   0.0302
  -3.750  -0.3533   0.01322   0.00519  -0.0174   1.0000   0.0376
  -3.500  -0.3313   0.01265   0.00459  -0.0164   1.0000   0.0447
  -3.250  -0.3087   0.01229   0.00415  -0.0154   1.0000   0.0506
  -3.000  -0.2865   0.01190   0.00374  -0.0145   1.0000   0.0577
  -2.750  -0.2641   0.01159   0.00337  -0.0135   1.0000   0.0652
  -2.500  -0.2326   0.01111   0.00298  -0.0145   0.9965   0.0958
  -2.250  -0.1987   0.01046   0.00271  -0.0164   0.9907   0.1935
  -2.000  -0.1653   0.00988   0.00254  -0.0181   0.9847   0.3136
  -1.750  -0.1334   0.00930   0.00242  -0.0193   0.9781   0.4432
  -1.500  -0.1042   0.00878   0.00237  -0.0198   0.9702   0.5668
  -1.250  -0.0769   0.00820   0.00236  -0.0194   0.9629   0.7090
  -1.000  -0.0352   0.00774   0.00244  -0.0215   0.9620   0.8947
  -0.750   0.0315   0.00772   0.00241  -0.0298   0.9675   0.9968
  -0.500   0.0682   0.00771   0.00233  -0.0318   0.9586   1.0000
  -0.250   0.1048   0.00769   0.00225  -0.0338   0.9490   1.0000
   0.000   0.1429   0.00765   0.00218  -0.0360   0.9389   1.0000
   0.250   0.1820   0.00759   0.00210  -0.0384   0.9243   1.0000
   0.500   0.2205   0.00754   0.00203  -0.0405   0.9062   1.0000
   0.750   0.2557   0.00752   0.00199  -0.0419   0.8887   1.0000
   1.000   0.2879   0.00753   0.00198  -0.0426   0.8626   1.0000
   1.250   0.3159   0.00760   0.00196  -0.0422   0.8295   1.0000
   1.500   0.3401   0.00772   0.00195  -0.0410   0.7838   1.0000
   1.750   0.3622   0.00792   0.00197  -0.0393   0.7237   1.0000
   2.000   0.3843   0.00816   0.00201  -0.0378   0.6682   1.0000
   2.250   0.4055   0.00848   0.00209  -0.0361   0.6023   1.0000
   2.500   0.4253   0.00894   0.00222  -0.0343   0.5178   1.0000
   2.750   0.4446   0.00955   0.00245  -0.0326   0.4147   1.0000
   3.000   0.4634   0.01036   0.00275  -0.0311   0.2935   1.0000
   3.250   0.4806   0.01157   0.00322  -0.0296   0.1283   1.0000
   3.500   0.5025   0.01225   0.00367  -0.0286   0.0736   1.0000
   3.750   0.5256   0.01279   0.00420  -0.0277   0.0592   1.0000
   4.000   0.5487   0.01332   0.00477  -0.0269   0.0459   1.0000
   4.250   0.5717   0.01390   0.00543  -0.0260   0.0352   1.0000
   4.500   0.5950   0.01446   0.00602  -0.0251   0.0253   1.0000
   4.750   0.6166   0.01535   0.00699  -0.0239   0.0210   1.0000
   5.000   0.6393   0.01608   0.00785  -0.0228   0.0170   1.0000
   5.250   0.6611   0.01710   0.00896  -0.0216   0.0148   1.0000
   5.500   0.6809   0.01896   0.01096  -0.0200   0.0132   1.0000
   5.750   0.7024   0.02107   0.01331  -0.0187   0.0121   1.0000
   6.000   0.7257   0.02230   0.01478  -0.0177   0.0108   1.0000
   6.250   0.7475   0.02446   0.01727  -0.0164   0.0101   1.0000
   6.500   0.7676   0.02711   0.02033  -0.0148   0.0097   1.0000
   6.750   0.7852   0.03025   0.02399  -0.0130   0.0095   1.0000
   7.000   0.7997   0.03404   0.02827  -0.0109   0.0093   1.0000
   7.250   0.8106   0.03836   0.03308  -0.0087   0.0094   1.0000
   7.500   0.8180   0.04302   0.03819  -0.0066   0.0095   1.0000
   7.750   0.8234   0.04750   0.04303  -0.0049   0.0093   1.0000
   8.000   0.8284   0.05126   0.04706  -0.0035   0.0088   1.0000
   8.250   0.8264   0.05627   0.05235  -0.0023   0.0089   1.0000
   8.500   0.8200   0.06136   0.05767  -0.0015   0.0091   1.0000
   8.750   0.8150   0.06504   0.06149  -0.0012   0.0086   1.0000
   9.000   0.7929   0.07092   0.06750  -0.0012   0.0094   1.0000
   9.250   0.7772   0.07650   0.07316  -0.0044   0.0092   1.0000
   9.500   0.7627   0.08523   0.08188  -0.0122   0.0098   1.0000
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