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EPPLER 435 AIRFOIL (e435-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 435 AIRFOIL (e435-il)
Reynolds number: 50,000
Max Cl/Cd: 7.64 at α=10.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e435-il-50000.txt
Download as CSV file: xf-e435-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 435 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
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.4267   0.12025   0.11512  -0.0073   1.0000   0.3174
  -6.750  -0.4707   0.12073   0.11576  -0.0049   1.0000   0.3278
  -6.500  -0.4666   0.11838   0.11343  -0.0024   1.0000   0.3440
  -6.250  -0.4684   0.11632   0.11142   0.0003   1.0000   0.3607
  -6.000  -0.4644   0.11389   0.10903   0.0027   1.0000   0.3770
  -5.750  -0.4525   0.11074   0.10589   0.0046   1.0000   0.3910
  -5.500  -0.4514   0.10829   0.10349   0.0068   1.0000   0.4041
  -5.250  -0.4565   0.10622   0.10148   0.0094   1.0000   0.4185
  -5.000  -0.5718   0.06949   0.06297  -0.0466   1.0000   0.1287
  -4.750  -0.5582   0.06546   0.05911  -0.0459   1.0000   0.1243
  -4.500  -0.5292   0.05958   0.05227  -0.0505   1.0000   0.1118
  -4.250  -0.5067   0.05605   0.04841  -0.0519   1.0000   0.1098
  -4.000  -0.4806   0.05271   0.04457  -0.0536   1.0000   0.1078
  -3.750  -0.4534   0.04988   0.04121  -0.0548   1.0000   0.1057
  -3.500  -0.4268   0.04762   0.03850  -0.0555   1.0000   0.1059
  -3.250  -0.4012   0.04602   0.03647  -0.0559   1.0000   0.1100
  -3.000  -0.3760   0.04452   0.03461  -0.0559   1.0000   0.1153
  -2.750  -0.3541   0.04339   0.03346  -0.0553   1.0000   0.1236
  -2.500  -0.3327   0.04245   0.03249  -0.0542   1.0000   0.1355
  -2.250  -0.3122   0.04172   0.03182  -0.0529   1.0000   0.1540
  -2.000  -0.2913   0.04097   0.03126  -0.0518   1.0000   0.1841
  -1.750  -0.2606   0.03854   0.03192  -0.0518   1.0000   0.5425
  -1.500  -0.2785   0.04052   0.03402  -0.0382   1.0000   0.7126
  -1.250  -0.2918   0.04131   0.03475  -0.0273   1.0000   0.7656
  -1.000  -0.3005   0.04167   0.03501  -0.0183   1.0000   0.8095
  -0.750  -0.3093   0.04162   0.03487  -0.0097   1.0000   0.8471
  -0.500  -0.3148   0.04141   0.03455  -0.0023   1.0000   0.8862
  -0.250  -0.1909   0.04539   0.03788  -0.0163   1.0000   1.0000
   0.000  -0.1905   0.04485   0.03720  -0.0143   1.0000   1.0000
   0.250  -0.1897   0.04431   0.03654  -0.0123   1.0000   1.0000
   0.500  -0.1880   0.04378   0.03589  -0.0105   1.0000   1.0000
   0.750  -0.1841   0.04339   0.03537  -0.0091   1.0000   1.0000
   1.000  -0.1748   0.04333   0.03515  -0.0087   1.0000   1.0000
   1.250  -0.1600   0.04362   0.03529  -0.0093   1.0000   1.0000
   1.500  -0.1413   0.04423   0.03573  -0.0106   1.0000   1.0000
   1.750  -0.1117   0.04568   0.03698  -0.0140   0.9965   1.0000
   2.000  -0.0772   0.04759   0.03869  -0.0184   0.9908   1.0000
   2.250  -0.0363   0.05020   0.04109  -0.0239   0.9836   1.0000
   2.500  -0.0058   0.05169   0.04244  -0.0274   0.9746   1.0000
   2.750   0.0262   0.05359   0.04420  -0.0312   0.9668   1.0000
   3.000   0.0693   0.05665   0.04710  -0.0370   0.9582   1.0000
   3.250   0.0956   0.05784   0.04821  -0.0397   0.9468   1.0000
   3.500   0.1226   0.05945   0.04974  -0.0425   0.9363   1.0000
   3.750   0.1578   0.06204   0.05224  -0.0467   0.9278   1.0000
   4.000   0.1929   0.06438   0.05449  -0.0509   0.9155   1.0000
   4.250   0.2148   0.06567   0.05574  -0.0527   0.9029   1.0000
   4.500   0.2402   0.06761   0.05764  -0.0552   0.8921   1.0000
   4.750   0.2812   0.07101   0.06097  -0.0602   0.8820   1.0000
   5.000   0.3044   0.07249   0.06244  -0.0622   0.8682   1.0000
   5.250   0.3236   0.07403   0.06398  -0.0636   0.8552   1.0000
   5.500   0.3467   0.07616   0.06610  -0.0657   0.8438   1.0000
   5.750   0.3841   0.07948   0.06939  -0.0698   0.8330   1.0000
   6.000   0.4097   0.08152   0.07145  -0.0720   0.8185   1.0000
   6.250   0.4236   0.08299   0.07296  -0.0726   0.8049   1.0000
   6.500   0.4416   0.08505   0.07503  -0.0739   0.7921   1.0000
   6.750   0.4666   0.08775   0.07775  -0.0761   0.7813   1.0000
   7.000   0.5005   0.09086   0.08087  -0.0792   0.7677   1.0000
   7.250   0.5214   0.09292   0.08300  -0.0806   0.7529   1.0000
   7.500   0.5322   0.09465   0.08477  -0.0809   0.7388   1.0000
   7.750   0.5447   0.09685   0.08702  -0.0816   0.7264   1.0000
   8.000   0.5642   0.09948   0.08970  -0.0830   0.7143   1.0000
   8.250   0.5900   0.10232   0.09260  -0.0849   0.7010   1.0000
   8.500   0.6142   0.10492   0.09528  -0.0864   0.6859   1.0000
   8.750   0.6337   0.10722   0.09765  -0.0874   0.6699   1.0000
   9.000   0.7142   0.10125   0.09171  -0.0856   0.5822   1.0000
   9.250   0.7340   0.10302   0.09356  -0.0860   0.5670   1.0000
   9.500   0.7530   0.10486   0.09549  -0.0865   0.5523   1.0000
   9.750   0.7724   0.10674   0.09748  -0.0869   0.5382   1.0000
  10.000   0.7937   0.10847   0.09931  -0.0873   0.5239   1.0000
  10.250   0.8194   0.10998   0.10093  -0.0877   0.5102   1.0000
  10.500   0.8486   0.11102   0.10212  -0.0880   0.4958   1.0000
  10.750   0.8506   0.11414   0.10534  -0.0882   0.4834   1.0000
  11.000   0.8506   0.11749   0.10877  -0.0885   0.4709   1.0000
  11.250   0.8626   0.11999   0.11138  -0.0887   0.4579   1.0000
  11.500   0.8812   0.12193   0.11347  -0.0889   0.4445   1.0000
  11.750   0.9049   0.12321   0.11488  -0.0888   0.4308   1.0000
  12.000   0.9383   0.12330   0.11515  -0.0883   0.4170   1.0000
  12.250   0.9116   0.12999   0.12187  -0.0899   0.4065   1.0000
  12.500   0.9179   0.13329   0.12528  -0.0905   0.3951   1.0000
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