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E61 (5.64%) (e61-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: E61 (5.64%) (e61-il)
Reynolds number: 500,000
Max Cl/Cd: 157.06 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e61-il-500000-n5.txt
Download as CSV file: xf-e61-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E61  (5.64%)                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.0727   0.09337   0.09104  -0.0790   0.9659   0.0064
  -9.000  -0.0639   0.09013   0.08780  -0.0804   0.9639   0.0066
  -8.750  -0.0538   0.08670   0.08437  -0.0823   0.9623   0.0069
  -8.500  -0.0427   0.08322   0.08090  -0.0844   0.9611   0.0072
  -8.250  -0.0298   0.07976   0.07744  -0.0870   0.9600   0.0084
  -7.500  -0.0881   0.08858   0.08615  -0.0846   0.9612   0.0068
  -7.000   0.0032   0.06447   0.06220  -0.0911   0.9400   0.0084
  -6.000  -0.0178   0.06900   0.06664  -0.0991   0.9353   0.0053
  -5.750  -0.0096   0.06684   0.06449  -0.0997   0.9273   0.0056
  -5.500   0.0138   0.06339   0.06103  -0.1048   0.9244   0.0058
  -5.250   0.0423   0.05955   0.05717  -0.1113   0.9224   0.0062
  -5.000   0.0591   0.05679   0.05441  -0.1141   0.9144   0.0071
  -4.750   0.1030   0.05211   0.04968  -0.1248   0.9116   0.0088
  -4.500   0.1421   0.04719   0.04471  -0.1339   0.9098   0.0084
  -4.250   0.1804   0.04252   0.03995  -0.1423   0.9039   0.0084
  -4.000   0.2395   0.03634   0.03364  -0.1556   0.9012   0.0089
  -3.750   0.2976   0.03086   0.02800  -0.1682   0.8996   0.0105
  -3.500   0.3438   0.02829   0.02530  -0.1749   0.8981   0.0139
  -3.250   0.4148   0.02166   0.01826  -0.1882   0.8978   0.0130
  -3.000   0.4810   0.01602   0.01197  -0.1985   0.8977   0.0126
  -2.750   0.5307   0.01349   0.00896  -0.2039   0.8968   0.0137
  -2.500   0.5695   0.01221   0.00738  -0.2066   0.8931   0.0144
  -2.250   0.6041   0.01188   0.00689  -0.2081   0.8884   0.0159
  -2.000   0.6471   0.01067   0.00541  -0.2116   0.8856   0.0156
  -1.750   0.6913   0.00971   0.00425  -0.2154   0.8829   0.0153
  -1.500   0.7264   0.00912   0.00355  -0.2171   0.8763   0.0151
  -1.250   0.7658   0.00864   0.00296  -0.2197   0.8707   0.0149
  -1.000   0.8004   0.00831   0.00255  -0.2213   0.8627   0.0149
  -0.750   0.8360   0.00806   0.00221  -0.2230   0.8541   0.0151
  -0.500   0.8688   0.00793   0.00197  -0.2241   0.8436   0.0155
  -0.250   0.8991   0.00787   0.00181  -0.2246   0.8321   0.0161
   0.000   0.9287   0.00784   0.00167  -0.2249   0.8205   0.0205
   0.250   0.9577   0.00781   0.00166  -0.2252   0.8087   0.0502
   0.500   0.9885   0.00754   0.00180  -0.2263   0.7963   0.2469
   0.750   1.0171   0.00745   0.00193  -0.2266   0.7833   0.3755
   1.000   1.0396   0.00672   0.00226  -0.2256   0.7705   0.8902
   1.500   1.0884   0.00693   0.00234  -0.2240   0.7423   1.0000
   1.750   1.1137   0.00710   0.00243  -0.2234   0.7275   1.0000
   2.000   1.1387   0.00727   0.00256  -0.2228   0.7121   1.0000
   2.250   1.1630   0.00748   0.00268  -0.2220   0.6945   1.0000
   2.500   1.1867   0.00769   0.00281  -0.2212   0.6743   1.0000
   2.750   1.2097   0.00795   0.00297  -0.2201   0.6524   1.0000
   3.000   1.2323   0.00822   0.00315  -0.2190   0.6273   1.0000
   3.250   1.2533   0.00858   0.00340  -0.2176   0.5958   1.0000
   3.500   1.2726   0.00904   0.00367  -0.2158   0.5545   1.0000
   3.750   1.2910   0.00959   0.00400  -0.2139   0.5069   1.0000
   4.000   1.3091   0.01022   0.00439  -0.2120   0.4584   1.0000
   4.250   1.3219   0.01132   0.00499  -0.2092   0.3690   1.0000
   4.500   1.3336   0.01266   0.00571  -0.2064   0.2645   1.0000
   4.750   1.3502   0.01361   0.00630  -0.2045   0.2006   1.0000
   5.000   1.3658   0.01467   0.00701  -0.2025   0.1343   1.0000
   5.250   1.3795   0.01591   0.00782  -0.2001   0.0661   1.0000
   5.500   1.3946   0.01695   0.00860  -0.1978   0.0235   1.0000
   5.750   1.4116   0.01779   0.00933  -0.1958   0.0064   1.0000
   6.000   1.4311   0.01834   0.00995  -0.1942   0.0042   1.0000
   6.250   1.4503   0.01888   0.01058  -0.1926   0.0038   1.0000
   6.500   1.4689   0.01949   0.01129  -0.1908   0.0035   1.0000
   6.750   1.4865   0.02020   0.01211  -0.1889   0.0031   1.0000
   7.000   1.5031   0.02100   0.01302  -0.1869   0.0029   1.0000
   7.250   1.5186   0.02189   0.01408  -0.1846   0.0027   1.0000
   7.500   1.5330   0.02289   0.01520  -0.1822   0.0025   1.0000
   7.750   1.5461   0.02401   0.01645  -0.1796   0.0025   1.0000
   8.000   1.5579   0.02524   0.01780  -0.1769   0.0024   1.0000
   8.250   1.5683   0.02662   0.01931  -0.1739   0.0023   1.0000
   8.500   1.5780   0.02811   0.02094  -0.1710   0.0023   1.0000
   8.750   1.5866   0.02974   0.02271  -0.1679   0.0023   1.0000
   9.000   1.5954   0.03144   0.02457  -0.1650   0.0023   1.0000
   9.250   1.6034   0.03334   0.02663  -0.1620   0.0023   1.0000
   9.500   1.6117   0.03536   0.02883  -0.1592   0.0023   1.0000
   9.750   1.6197   0.03757   0.03127  -0.1564   0.0023   1.0000
  10.000   1.6276   0.03995   0.03385  -0.1538   0.0023   1.0000
  10.250   1.6346   0.04258   0.03672  -0.1511   0.0023   1.0000
  10.500   1.6400   0.04507   0.03941  -0.1485   0.0023   1.0000
  10.750   1.6429   0.04778   0.04234  -0.1458   0.0022   1.0000
  11.000   1.6437   0.05077   0.04556  -0.1431   0.0021   1.0000
  11.250   1.6415   0.05396   0.04895  -0.1404   0.0020   1.0000
  11.500   1.6330   0.05820   0.05344  -0.1376   0.0019   1.0000
  11.750   1.6213   0.06302   0.05856  -0.1349   0.0019   1.0000
  12.000   1.5987   0.06970   0.06561  -0.1323   0.0018   1.0000
  12.250   1.5754   0.07636   0.07260  -0.1307   0.0017   1.0000
  13.000   1.5334   0.09312   0.09002  -0.1311   0.0017   1.0000
  13.500   1.3298   0.11123   0.10879  -0.1231   0.0021   1.0000
  13.750   1.2808   0.12376   0.12157  -0.1295   0.0017   1.0000
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