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EPPLER 904 AIRFOIL (e904-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 904 AIRFOIL (e904-il)
Reynolds number: 50,000
Max Cl/Cd: 23.95 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e904-il-50000-n5.txt
Download as CSV file: xf-e904-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 904 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4765   0.09849   0.09155  -0.0403   1.0000   0.0309
  -9.000  -0.4784   0.09497   0.08810  -0.0408   1.0000   0.0300
  -8.750  -0.4846   0.09099   0.08421  -0.0423   1.0000   0.0302
  -8.500  -0.4928   0.08710   0.08041  -0.0438   1.0000   0.0303
  -8.250  -0.5000   0.08391   0.07730  -0.0435   1.0000   0.0281
  -8.000  -0.5144   0.08083   0.07431  -0.0425   1.0000   0.0277
  -7.750  -0.5281   0.07788   0.07142  -0.0408   1.0000   0.0268
  -7.500  -0.5403   0.07489   0.06845  -0.0390   1.0000   0.0265
  -7.250  -0.5520   0.07179   0.06533  -0.0368   1.0000   0.0256
  -7.000  -0.5619   0.06881   0.06230  -0.0344   1.0000   0.0254
  -6.750  -0.5703   0.06559   0.05897  -0.0319   1.0000   0.0247
  -6.500  -0.5763   0.06214   0.05535  -0.0293   1.0000   0.0236
  -6.250  -0.5791   0.05908   0.05210  -0.0267   1.0000   0.0238
  -6.000  -0.5792   0.05625   0.04908  -0.0242   1.0000   0.0251
  -5.750  -0.5768   0.05312   0.04567  -0.0217   1.0000   0.0257
  -5.500  -0.5716   0.04999   0.04222  -0.0193   1.0000   0.0260
  -5.250  -0.5638   0.04682   0.03867  -0.0170   1.0000   0.0262
  -5.000  -0.5533   0.04381   0.03523  -0.0147   1.0000   0.0269
  -4.750  -0.5400   0.04078   0.03173  -0.0125   1.0000   0.0275
  -4.500  -0.5238   0.03791   0.02835  -0.0104   1.0000   0.0276
  -4.250  -0.5047   0.03515   0.02488  -0.0085   1.0000   0.0276
  -4.000  -0.4833   0.03280   0.02205  -0.0068   1.0000   0.0278
  -3.750  -0.4597   0.03062   0.01942  -0.0053   1.0000   0.0287
  -3.500  -0.4349   0.02872   0.01712  -0.0037   1.0000   0.0304
  -3.250  -0.4109   0.02731   0.01535  -0.0022   1.0000   0.0325
  -3.000  -0.3875   0.02582   0.01371  -0.0011   1.0000   0.0391
  -2.750  -0.3647   0.02464   0.01251  -0.0001   1.0000   0.0638
  -2.500  -0.1722   0.02173   0.01234  -0.0259   1.0000   1.0000
  -2.250  -0.1630   0.02161   0.01174  -0.0232   1.0000   1.0000
  -2.000  -0.1533   0.02152   0.01138  -0.0205   1.0000   1.0000
  -1.750  -0.1431   0.02146   0.01108  -0.0179   1.0000   1.0000
  -1.500  -0.1326   0.02141   0.01081  -0.0154   1.0000   1.0000
  -1.250  -0.1217   0.02137   0.01057  -0.0129   1.0000   1.0000
  -1.000  -0.1104   0.02136   0.01039  -0.0105   1.0000   1.0000
  -0.750  -0.0987   0.02136   0.01022  -0.0081   1.0000   1.0000
  -0.500  -0.0867   0.02139   0.01009  -0.0058   1.0000   1.0000
  -0.250  -0.0742   0.02143   0.00999  -0.0036   1.0000   1.0000
   0.000  -0.0612   0.02149   0.00983  -0.0015   1.0000   1.0000
   0.250  -0.0479   0.02158   0.00981   0.0005   1.0000   1.0000
   0.500  -0.0178   0.02191   0.01003  -0.0009   0.9950   1.0000
   0.750   0.0121   0.02223   0.01028  -0.0023   0.9893   1.0000
   1.000   0.0431   0.02264   0.01063  -0.0039   0.9835   1.0000
   1.250   0.0725   0.02298   0.01095  -0.0051   0.9772   1.0000
   1.500   0.1040   0.02339   0.01136  -0.0067   0.9707   1.0000
   1.750   0.1329   0.02372   0.01172  -0.0079   0.9633   1.0000
   2.000   0.1670   0.02418   0.01224  -0.0100   0.9565   1.0000
   2.250   0.1951   0.02448   0.01262  -0.0108   0.9475   1.0000
   2.500   0.2269   0.02486   0.01312  -0.0123   0.9390   1.0000
   2.750   0.2637   0.02529   0.01372  -0.0147   0.9305   1.0000
   3.000   0.2939   0.02559   0.01441  -0.0157   0.9195   1.0000
   3.250   0.3267   0.02589   0.01494  -0.0171   0.9079   1.0000
   3.500   0.3616   0.02614   0.01550  -0.0188   0.8952   1.0000
   3.750   0.3986   0.02632   0.01601  -0.0206   0.8813   1.0000
   4.000   0.4387   0.02639   0.01650  -0.0228   0.8664   1.0000
   4.250   0.4737   0.02637   0.01697  -0.0237   0.8492   1.0000
   4.500   0.5074   0.02626   0.01763  -0.0241   0.8301   1.0000
   4.750   0.5449   0.02275   0.01475  -0.0185   0.7144   1.0000
   5.000   0.5830   0.02643   0.01395  -0.0156   0.0512   1.0000
   5.250   0.5986   0.02781   0.01557  -0.0130   0.0433   1.0000
   5.500   0.6184   0.02948   0.01740  -0.0113   0.0368   1.0000
   5.750   0.6447   0.03116   0.01927  -0.0108   0.0286   1.0000
   6.000   0.6843   0.03401   0.02236  -0.0119   0.0260   1.0000
   6.250   0.7172   0.03668   0.02544  -0.0118   0.0250   1.0000
   6.500   0.7427   0.03964   0.02885  -0.0105   0.0243   1.0000
   6.750   0.7610   0.04252   0.03217  -0.0084   0.0240   1.0000
   7.000   0.7743   0.04553   0.03562  -0.0057   0.0239   1.0000
   7.250   0.7834   0.04865   0.03917  -0.0026   0.0240   1.0000
   7.500   0.7893   0.05173   0.04262   0.0005   0.0242   1.0000
   7.750   0.7923   0.05488   0.04612   0.0037   0.0244   1.0000
   8.000   0.7924   0.05812   0.04988   0.0069   0.0247   1.0000
   8.250   0.7903   0.06135   0.05338   0.0099   0.0250   1.0000
   8.500   0.7846   0.06478   0.05705   0.0130   0.0253   1.0000
   8.750   0.7774   0.06808   0.06055   0.0158   0.0256   1.0000
   9.000   0.7672   0.07132   0.06394   0.0185   0.0259   1.0000
   9.250   0.7554   0.07441   0.06715   0.0211   0.0261   1.0000
   9.500   0.7420   0.07796   0.07079   0.0229   0.0263   1.0000
   9.750   0.7305   0.08166   0.07456   0.0237   0.0266   1.0000
  10.250   0.6161   0.08415   0.07745   0.0253   0.0265   1.0000
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