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NASA SC(2)-0404 AIRFOIL (sc20404-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0404 AIRFOIL (sc20404-il)
Reynolds number: 500,000
Max Cl/Cd: 41.53 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20404-il-500000-n5.txt
Download as CSV file: xf-sc20404-il-500000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6756   0.11283   0.11056   0.0263   1.0000   0.0073
  -9.250  -0.6723   0.10876   0.10650   0.0244   1.0000   0.0076
  -9.000  -0.6691   0.10468   0.10243   0.0223   1.0000   0.0078
  -8.750  -0.6655   0.10059   0.09835   0.0201   1.0000   0.0082
  -8.500  -0.6581   0.09686   0.09465   0.0169   1.0000   0.0092
  -8.250  -0.6525   0.09300   0.09080   0.0142   1.0000   0.0095
  -8.000  -0.6486   0.08883   0.08665   0.0112   1.0000   0.0096
  -7.750  -0.6430   0.08415   0.08199   0.0066   1.0000   0.0096
  -7.500  -0.6320   0.07834   0.07618  -0.0011   1.0000   0.0097
  -7.250  -0.6189   0.07236   0.07016  -0.0083   1.0000   0.0097
  -7.000  -0.6147   0.06478   0.06250  -0.0147   1.0000   0.0086
  -6.750  -0.5421   0.04789   0.04567  -0.0198   1.0000   0.0096
  -6.250  -0.5563   0.04371   0.04081  -0.0277   1.0000   0.0037
  -6.000  -0.5331   0.03747   0.03425  -0.0304   1.0000   0.0035
  -5.750  -0.5069   0.03159   0.02798  -0.0323   1.0000   0.0034
  -5.500  -0.4782   0.02582   0.02172  -0.0337   1.0000   0.0033
  -5.250  -0.4476   0.02037   0.01568  -0.0348   1.0000   0.0033
  -5.000  -0.4169   0.01635   0.01106  -0.0354   1.0000   0.0034
  -4.750  -0.3875   0.01397   0.00828  -0.0356   1.0000   0.0037
  -4.500  -0.3586   0.01198   0.00594  -0.0360   1.0000   0.0051
  -4.250  -0.3312   0.01169   0.00562  -0.0361   1.0000   0.0067
  -4.000  -0.3035   0.01127   0.00514  -0.0362   1.0000   0.0086
  -3.750  -0.2746   0.01024   0.00397  -0.0364   1.0000   0.0094
  -3.500  -0.2457   0.00949   0.00312  -0.0368   1.0000   0.0110
  -3.250  -0.2166   0.00882   0.00233  -0.0372   1.0000   0.0142
  -3.000  -0.1886   0.00852   0.00199  -0.0373   1.0000   0.0177
  -2.750  -0.1605   0.00824   0.00154  -0.0374   1.0000   0.0209
  -2.500  -0.1325   0.00794   0.00132  -0.0375   1.0000   0.0425
  -2.250  -0.1048   0.00764   0.00116  -0.0377   1.0000   0.0828
  -2.000  -0.0765   0.00720   0.00101  -0.0382   1.0000   0.1719
  -1.750  -0.0456   0.00592   0.00085  -0.0400   1.0000   0.4960
  -1.500  -0.0177   0.00544   0.00091  -0.0403   1.0000   0.6540
  -1.250   0.0090   0.00531   0.00094  -0.0400   1.0000   0.7028
  -1.000   0.0353   0.00522   0.00100  -0.0396   1.0000   0.7450
  -0.750   0.0616   0.00518   0.00104  -0.0393   1.0000   0.7676
  -0.500   0.0881   0.00516   0.00108  -0.0389   1.0000   0.7841
   0.000   0.1536   0.00507   0.00114  -0.0412   0.9878   0.8108
   0.250   0.1956   0.00524   0.00106  -0.0435   0.8401   0.8196
   0.500   0.2120   0.00729   0.00121  -0.0414   0.3422   0.8289
   0.750   0.2379   0.00806   0.00136  -0.0415   0.1536   0.8391
   1.000   0.2645   0.00846   0.00151  -0.0414   0.0716   0.8494
   1.250   0.2911   0.00873   0.00168  -0.0412   0.0330   0.8595
   1.500   0.3179   0.00895   0.00191  -0.0409   0.0187   0.8708
   1.750   0.3444   0.00918   0.00233  -0.0404   0.0164   0.8835
   2.000   0.3702   0.00951   0.00281  -0.0398   0.0156   0.8980
   2.250   0.3951   0.00976   0.00315  -0.0390   0.0132   0.9151
   2.500   0.4182   0.01038   0.00392  -0.0378   0.0105   0.9388
   2.750   0.4435   0.01132   0.00502  -0.0370   0.0088   1.0000
   3.000   0.4712   0.01215   0.00596  -0.0369   0.0077   1.0000
   3.250   0.4994   0.01217   0.00598  -0.0372   0.0053   1.0000
   3.500   0.5270   0.01269   0.00654  -0.0372   0.0042   1.0000
   3.750   0.5534   0.01502   0.00929  -0.0366   0.0032   1.0000
   4.000   0.5803   0.01797   0.01273  -0.0358   0.0030   1.0000
   4.250   0.6063   0.02247   0.01786  -0.0345   0.0029   1.0000
   4.500   0.6309   0.02773   0.02367  -0.0333   0.0030   1.0000
   4.750   0.6542   0.03306   0.02944  -0.0324   0.0031   1.0000
   5.000   0.6760   0.03852   0.03528  -0.0319   0.0032   1.0000
   5.250   0.6962   0.04422   0.04130  -0.0319   0.0033   1.0000
   7.250   0.8047   0.09205   0.09018  -0.0556   0.0076   1.0000
   7.500   0.8067   0.09714   0.09527  -0.0607   0.0075   1.0000
   7.750   0.8062   0.10179   0.09991  -0.0643   0.0074   1.0000
   8.000   0.8067   0.10631   0.10441  -0.0672   0.0072   1.0000
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