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NASA SC(2)-0404 AIRFOIL (sc20404-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0404 AIRFOIL (sc20404-il)
Reynolds number: 500,000
Max Cl/Cd: 52.03 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20404-il-500000.txt
Download as CSV file: xf-sc20404-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0404 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5576   0.09600   0.09386   0.0081   1.0000   0.0091
  -9.000  -0.5599   0.09109   0.08897   0.0067   1.0000   0.0092
  -8.750  -0.5621   0.08628   0.08417   0.0053   1.0000   0.0094
  -8.500  -0.5640   0.08157   0.07948   0.0037   1.0000   0.0095
  -8.250  -0.5664   0.07690   0.07482   0.0019   1.0000   0.0096
  -8.000  -0.5696   0.07213   0.07008  -0.0003   1.0000   0.0097
  -7.750  -0.5739   0.06727   0.06524  -0.0032   1.0000   0.0097
  -7.500  -0.5748   0.06092   0.05889  -0.0108   1.0000   0.0097
  -7.250  -0.5717   0.05463   0.05253  -0.0162   1.0000   0.0097
  -7.000  -0.5653   0.04891   0.04671  -0.0200   1.0000   0.0099
  -6.750  -0.5559   0.04347   0.04118  -0.0230   1.0000   0.0101
  -6.500  -0.5435   0.03824   0.03581  -0.0256   1.0000   0.0104
  -6.250  -0.5284   0.03340   0.03081  -0.0277   1.0000   0.0108
  -6.000  -0.5106   0.02869   0.02591  -0.0295   1.0000   0.0113
  -5.750  -0.4903   0.02437   0.02137  -0.0311   1.0000   0.0119
  -5.500  -0.4678   0.02035   0.01709  -0.0323   1.0000   0.0128
  -5.250  -0.4432   0.01675   0.01317  -0.0333   1.0000   0.0141
  -5.000  -0.4164   0.01395   0.01003  -0.0338   1.0000   0.0160
  -4.750  -0.3870   0.01377   0.00954  -0.0328   1.0000   0.0182
  -4.500  -0.3653   0.01727   0.01191  -0.0364   1.0000   0.0122
  -4.250  -0.3355   0.01514   0.00947  -0.0365   1.0000   0.0140
  -4.000  -0.3075   0.01469   0.00891  -0.0364   1.0000   0.0173
  -3.750  -0.2782   0.01229   0.00626  -0.0364   1.0000   0.0161
  -3.500  -0.2492   0.01074   0.00458  -0.0365   1.0000   0.0159
  -3.250  -0.2197   0.00968   0.00340  -0.0367   1.0000   0.0171
  -3.000  -0.1897   0.00875   0.00234  -0.0373   1.0000   0.0211
  -2.750  -0.1611   0.00833   0.00182  -0.0375   1.0000   0.0298
  -2.500  -0.1293   0.00687   0.00124  -0.0393   1.0000   0.2730
  -2.250  -0.0991   0.00564   0.00110  -0.0408   1.0000   0.5900
  -2.000  -0.0724   0.00533   0.00115  -0.0405   1.0000   0.7062
  -1.750  -0.0463   0.00524   0.00119  -0.0399   1.0000   0.7577
  -1.500  -0.0201   0.00519   0.00118  -0.0395   1.0000   0.7784
  -1.250   0.0055   0.00514   0.00122  -0.0388   1.0000   0.8065
  -1.000   0.0302   0.00510   0.00126  -0.0379   1.0000   0.8365
  -0.750   0.0552   0.00506   0.00128  -0.0372   1.0000   0.8529
  -0.500   0.0803   0.00502   0.00130  -0.0365   1.0000   0.8675
  -0.250   0.1052   0.00498   0.00133  -0.0358   1.0000   0.8819
   0.000   0.1298   0.00494   0.00137  -0.0350   1.0000   0.8967
   0.250   0.1540   0.00490   0.00142  -0.0341   1.0000   0.9122
   0.500   0.1769   0.00483   0.00147  -0.0329   1.0000   0.9317
   0.750   0.1964   0.00471   0.00148  -0.0310   1.0000   0.9769
   1.000   0.2539   0.00488   0.00120  -0.0363   0.7872   0.9701
   1.250   0.2746   0.00680   0.00139  -0.0354   0.3444   1.0000
   1.500   0.3015   0.00819   0.00172  -0.0362   0.0599   1.0000
   1.750   0.3307   0.00894   0.00242  -0.0364   0.0245   1.0000
   2.000   0.3595   0.00951   0.00308  -0.0366   0.0186   1.0000
   2.250   0.3871   0.01069   0.00437  -0.0364   0.0158   1.0000
   2.500   0.4150   0.01191   0.00571  -0.0361   0.0148   1.0000
   2.750   0.4429   0.01377   0.00773  -0.0357   0.0151   1.0000
   3.000   0.4714   0.01586   0.00999  -0.0352   0.0178   1.0000
   4.250   0.6075   0.03142   0.02757  -0.0312   0.0132   1.0000
   4.500   0.6326   0.03510   0.03156  -0.0307   0.0117   1.0000
   4.750   0.6560   0.03902   0.03575  -0.0305   0.0108   1.0000
   5.000   0.6781   0.04308   0.04005  -0.0305   0.0101   1.0000
   5.250   0.6987   0.04723   0.04442  -0.0309   0.0096   1.0000
   5.500   0.7175   0.05147   0.04883  -0.0316   0.0092   1.0000
   5.750   0.7345   0.05587   0.05339  -0.0325   0.0088   1.0000
   6.000   0.7494   0.06063   0.05830  -0.0338   0.0085   1.0000
   6.250   0.7619   0.06613   0.06393  -0.0357   0.0083   1.0000
   6.500   0.7714   0.07229   0.07020  -0.0384   0.0081   1.0000
   6.750   0.7769   0.07939   0.07740  -0.0420   0.0079   1.0000
   7.000   0.7134   0.07275   0.07097  -0.0365   0.0084   1.0000
   7.250   0.7090   0.07837   0.07662  -0.0399   0.0084   1.0000
   7.500   0.7002   0.08351   0.08175  -0.0438   0.0084   1.0000
   7.750   0.6942   0.08849   0.08671  -0.0463   0.0084   1.0000
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