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RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 15 AIRFOIL (raf15-il)
Reynolds number: 50,000
Max Cl/Cd: 32.54 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf15-il-50000.txt
Download as CSV file: xf-raf15-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 15 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5005   0.10479   0.09773  -0.0006   1.0000   0.1943
  -8.000  -0.4977   0.10170   0.09470  -0.0003   1.0000   0.2023
  -7.750  -0.5213   0.10128   0.09444  -0.0021   1.0000   0.2054
  -7.500  -0.4979   0.09529   0.08844  -0.0001   1.0000   0.2122
  -7.250  -0.5109   0.09364   0.08690  -0.0012   1.0000   0.2192
  -7.000  -0.5043   0.08940   0.08271  -0.0005   1.0000   0.2242
  -6.750  -0.5074   0.08692   0.08027  -0.0012   1.0000   0.2337
  -6.500  -0.5005   0.08283   0.07623  -0.0003   1.0000   0.2396
  -6.250  -0.5085   0.08134   0.07476  -0.0028   1.0000   0.2501
  -6.000  -0.4929   0.07630   0.06978   0.0004   1.0000   0.2588
  -5.750  -0.4891   0.07294   0.06646   0.0010   1.0000   0.2714
  -5.500  -0.4845   0.06973   0.06328   0.0019   1.0000   0.2860
  -5.250  -0.4807   0.06681   0.06038   0.0027   1.0000   0.3055
  -5.000  -0.4766   0.06360   0.05719   0.0039   1.0000   0.3251
  -4.750  -0.4681   0.06009   0.05373   0.0063   1.0000   0.3436
  -4.500  -0.4634   0.05694   0.05061   0.0087   1.0000   0.3713
  -4.250  -0.4580   0.05372   0.04746   0.0121   1.0000   0.4027
  -4.000  -0.3853   0.04485   0.03676  -0.0094   1.0000   0.1593
  -3.750  -0.3663   0.04055   0.03176  -0.0086   1.0000   0.1475
  -3.500  -0.3480   0.03745   0.02833  -0.0070   1.0000   0.1466
  -3.250  -0.3286   0.03446   0.02502  -0.0053   1.0000   0.1452
  -3.000  -0.3077   0.03167   0.02163  -0.0035   1.0000   0.1471
  -2.750  -0.2876   0.03006   0.02003  -0.0019   1.0000   0.1564
  -2.500  -0.2640   0.02775   0.01714  -0.0003   1.0000   0.1622
  -2.250  -0.2406   0.02627   0.01542   0.0011   1.0000   0.1759
  -2.000  -0.2173   0.02494   0.01407   0.0024   1.0000   0.1958
  -1.750  -0.1915   0.02341   0.01239   0.0035   1.0000   0.2311
  -1.500  -0.1596   0.02174   0.01068   0.0034   1.0000   0.3088
  -1.250  -0.1264   0.02027   0.00945   0.0024   1.0000   0.3969
  -1.000   0.0111   0.01659   0.00776  -0.0189   1.0000   1.0000
  -0.750   0.0373   0.01671   0.00742  -0.0182   1.0000   1.0000
  -0.500   0.0601   0.01681   0.00716  -0.0170   1.0000   1.0000
  -0.250   0.0813   0.01691   0.00702  -0.0156   1.0000   1.0000
   0.000   0.1015   0.01703   0.00696  -0.0141   1.0000   1.0000
   0.250   0.1211   0.01717   0.00696  -0.0126   1.0000   1.0000
   0.500   0.1401   0.01734   0.00703  -0.0111   1.0000   1.0000
   0.750   0.1586   0.01755   0.00718  -0.0095   1.0000   1.0000
   1.000   0.1763   0.01780   0.00740  -0.0079   1.0000   1.0000
   1.250   0.1929   0.01812   0.00772  -0.0063   1.0000   1.0000
   1.500   0.2079   0.01853   0.00816  -0.0047   1.0000   1.0000
   1.750   0.2202   0.01912   0.00879  -0.0031   1.0000   1.0000
   2.000   0.2270   0.02005   0.00980  -0.0014   1.0000   1.0000
   2.250   0.3376   0.02068   0.01076  -0.0201   0.9439   1.0000
   2.500   0.4843   0.01904   0.00958  -0.0407   0.8598   1.0000
   2.750   0.5748   0.01836   0.00875  -0.0498   0.7545   1.0000
   3.000   0.6126   0.01894   0.00898  -0.0501   0.6867   1.0000
   3.250   0.6417   0.01972   0.00955  -0.0494   0.6371   1.0000
   3.500   0.6683   0.02054   0.01022  -0.0484   0.5959   1.0000
   3.750   0.6927   0.02139   0.01098  -0.0472   0.5602   1.0000
   4.000   0.7154   0.02228   0.01181  -0.0457   0.5272   1.0000
   4.250   0.7375   0.02322   0.01276  -0.0440   0.4965   1.0000
   4.500   0.7586   0.02418   0.01370  -0.0422   0.4665   1.0000
   4.750   0.7788   0.02519   0.01470  -0.0402   0.4372   1.0000
   5.000   0.7983   0.02619   0.01568  -0.0381   0.4073   1.0000
   5.250   0.8165   0.02713   0.01662  -0.0357   0.3757   1.0000
   5.500   0.8340   0.02799   0.01737  -0.0332   0.3424   1.0000
   5.750   0.8519   0.02887   0.01806  -0.0308   0.3090   1.0000
   6.000   0.8667   0.02998   0.01924  -0.0279   0.2755   1.0000
   6.250   0.8827   0.03121   0.02048  -0.0253   0.2446   1.0000
   6.500   0.9000   0.03243   0.02164  -0.0229   0.2187   1.0000
   6.750   0.9167   0.03451   0.02397  -0.0206   0.2010   1.0000
   7.000   0.9334   0.03662   0.02627  -0.0185   0.1874   1.0000
   7.250   0.9497   0.03858   0.02834  -0.0164   0.1751   1.0000
   7.500   0.9645   0.04103   0.03107  -0.0143   0.1664   1.0000
   7.750   0.9752   0.04435   0.03478  -0.0119   0.1618   1.0000
   8.000   0.9765   0.04840   0.03944  -0.0088   0.1587   1.0000
   8.250   0.9793   0.05203   0.04345  -0.0063   0.1546   1.0000
   8.500   0.9919   0.05472   0.04618  -0.0045   0.1483   1.0000
   8.750   0.9826   0.05964   0.05155  -0.0018   0.1478   1.0000
   9.000   0.9719   0.06485   0.05708   0.0002   0.1482   1.0000
   9.250   0.9615   0.07013   0.06256   0.0017   0.1490   1.0000
   9.500   0.8849   0.08075   0.07336   0.0004   0.1598   1.0000
   9.750   0.8341   0.09277   0.08526  -0.0073   0.1751   1.0000
  10.000   0.8282   0.09907   0.09155  -0.0097   0.1789   1.0000
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