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EPPLER 603 AIRFOIL (e603-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 603 AIRFOIL (e603-il)
Reynolds number: 50,000
Max Cl/Cd: 3.97 at α=11°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e603-il-50000.txt
Download as CSV file: xf-e603-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 603 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3352   0.12458   0.11917  -0.0324   0.9765   0.3469
  -9.000  -0.4651   0.10541   0.10019  -0.0503   0.9749   0.1805
  -8.750  -0.5579   0.09573   0.09055  -0.0540   0.9733   0.1556
  -8.500  -0.5862   0.09172   0.08647  -0.0529   0.9713   0.1463
  -8.250  -0.6286   0.08889   0.08360  -0.0494   0.9708   0.1434
  -8.000  -0.6806   0.08584   0.08038  -0.0444   0.9723   0.1400
  -7.750  -0.7072   0.08237   0.07666  -0.0408   0.9736   0.1335
  -7.500  -0.7214   0.07882   0.07299  -0.0379   0.9749   0.1294
  -7.250  -0.7739   0.07512   0.06844  -0.0311   0.9842   0.1217
  -7.000  -0.8066   0.07353   0.06702  -0.0219   1.0000   0.1223
  -6.750  -0.7995   0.06989   0.06319  -0.0206   1.0000   0.1190
  -6.500  -0.7941   0.06608   0.05912  -0.0193   1.0000   0.1163
  -6.250  -0.7875   0.06238   0.05499  -0.0182   1.0000   0.1144
  -6.000  -0.7775   0.05918   0.05135  -0.0172   1.0000   0.1141
  -5.750  -0.7646   0.05639   0.04814  -0.0163   1.0000   0.1150
  -5.500  -0.7492   0.05379   0.04510  -0.0154   1.0000   0.1160
  -5.250  -0.7318   0.05139   0.04225  -0.0145   1.0000   0.1170
  -5.000  -0.7129   0.04937   0.03976  -0.0136   1.0000   0.1194
  -4.750  -0.6941   0.04751   0.03765  -0.0127   1.0000   0.1236
  -4.500  -0.6755   0.04616   0.03627  -0.0118   1.0000   0.1294
  -4.250  -0.6551   0.04501   0.03474  -0.0106   1.0000   0.1364
  -4.000  -0.6366   0.04386   0.03375  -0.0093   1.0000   0.1465
  -3.750  -0.6174   0.04306   0.03301  -0.0075   1.0000   0.1596
  -3.500  -0.5993   0.04249   0.03256  -0.0053   1.0000   0.1785
  -3.250  -0.5834   0.04178   0.03210  -0.0030   1.0000   0.2071
  -3.000  -0.5706   0.04039   0.03132  -0.0009   1.0000   0.2654
  -2.750  -0.0806   0.06135   0.05307  -0.0483   0.9480   0.9936
  -2.500  -0.1587   0.06138   0.05329  -0.0321   0.9893   0.9975
  -2.250  -0.1801   0.06133   0.05315  -0.0258   1.0000   1.0000
  -2.000  -0.1765   0.06117   0.05284  -0.0243   1.0000   1.0000
  -1.750  -0.1726   0.06107   0.05259  -0.0227   1.0000   1.0000
  -1.500  -0.1683   0.06102   0.05240  -0.0211   1.0000   1.0000
  -1.250  -0.1639   0.06101   0.05226  -0.0195   1.0000   1.0000
  -1.000  -0.1592   0.06105   0.05217  -0.0179   1.0000   1.0000
  -0.750  -0.1544   0.06112   0.05212  -0.0163   1.0000   1.0000
  -0.500  -0.1495   0.06122   0.05209  -0.0147   1.0000   1.0000
  -0.250  -0.1445   0.06135   0.05212  -0.0131   1.0000   1.0000
   0.000  -0.1394   0.06150   0.05216  -0.0114   1.0000   1.0000
   0.250  -0.1342   0.06167   0.05224  -0.0098   1.0000   1.0000
   0.500  -0.1291   0.06186   0.05234  -0.0082   1.0000   1.0000
   0.750  -0.1239   0.06207   0.05245  -0.0066   1.0000   1.0000
   1.000  -0.1187   0.06231   0.05260  -0.0049   1.0000   1.0000
   1.250  -0.1136   0.06254   0.05276  -0.0033   1.0000   1.0000
   1.500  -0.1085   0.06280   0.05295  -0.0017   1.0000   1.0000
   1.750  -0.0791   0.06402   0.05408  -0.0051   0.9909   1.0000
   2.000  -0.0514   0.06532   0.05530  -0.0080   0.9793   1.0000
   2.250  -0.0273   0.06658   0.05649  -0.0101   0.9677   1.0000
   2.500  -0.0014   0.06831   0.05813  -0.0124   0.9561   1.0000
   2.750   0.0180   0.06910   0.05886  -0.0134   0.9430   1.0000
   3.000   0.0329   0.06968   0.05940  -0.0135   0.9297   1.0000
   3.250   0.0461   0.07040   0.06008  -0.0132   0.9175   1.0000
   3.500   0.0623   0.07167   0.06131  -0.0135   0.9073   1.0000
   3.750   0.0835   0.07317   0.06277  -0.0145   0.8950   1.0000
   4.000   0.0896   0.07330   0.06287  -0.0129   0.8818   1.0000
   4.250   0.0985   0.07408   0.06363  -0.0118   0.8709   1.0000
   4.500   0.1224   0.07623   0.06573  -0.0131   0.8609   1.0000
   4.750   0.1234   0.07598   0.06548  -0.0106   0.8474   1.0000
   5.000   0.1294   0.07670   0.06619  -0.0090   0.8374   1.0000
   5.250   0.1509   0.07866   0.06813  -0.0098   0.8271   1.0000
   5.500   0.1480   0.07833   0.06780  -0.0068   0.8145   1.0000
   5.750   0.1580   0.07962   0.06908  -0.0059   0.8060   1.0000
   6.000   0.1709   0.08065   0.07011  -0.0053   0.7940   1.0000
   6.250   0.1691   0.08073   0.07020  -0.0027   0.7827   1.0000
   6.500   0.1978   0.08377   0.07323  -0.0046   0.7749   1.0000
   6.750   0.1944   0.08323   0.07271  -0.0021   0.7618   1.0000
   7.000   0.2059   0.08473   0.07423  -0.0021   0.7532   1.0000
   7.250   0.2307   0.08695   0.07647  -0.0038   0.7422   1.0000
   7.500   0.2350   0.08768   0.07724  -0.0031   0.7309   1.0000
   7.750   0.2749   0.09192   0.08152  -0.0070   0.7233   1.0000
   8.000   0.2727   0.09170   0.08134  -0.0056   0.7103   1.0000
   8.250   0.2865   0.09372   0.08340  -0.0065   0.7013   1.0000
   8.500   0.3163   0.09667   0.08643  -0.0091   0.6907   1.0000
   8.750   0.3193   0.09769   0.08750  -0.0089   0.6793   1.0000
   9.000   0.3548   0.10188   0.09175  -0.0124   0.6717   1.0000
   9.250   0.3591   0.10261   0.09256  -0.0123   0.6587   1.0000
   9.500   0.3687   0.10471   0.09473  -0.0131   0.6496   1.0000
   9.750   0.4045   0.10865   0.09876  -0.0164   0.6393   1.0000
  10.000   0.4016   0.10947   0.09965  -0.0160   0.6280   1.0000
  10.250   0.4264   0.11313   0.10339  -0.0185   0.6205   1.0000
  10.500   0.4411   0.11519   0.10556  -0.0198   0.6078   1.0000
  10.750   0.4452   0.11724   0.10768  -0.0205   0.5986   1.0000
  11.000   0.4839   0.12199   0.11254  -0.0239   0.5891   1.0000
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