Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0404 AIRFOIL (sc20404-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0404 AIRFOIL (sc20404-il)
Reynolds number: 200,000
Max Cl/Cd: 35.17 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20404-il-200000-n5.txt
Download as CSV file: xf-sc20404-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.6239   0.07154   0.06807  -0.0091   1.0000   0.0093
  -7.000  -0.6133   0.06550   0.06194  -0.0144   1.0000   0.0090
  -6.750  -0.5993   0.05965   0.05596  -0.0190   1.0000   0.0086
  -6.500  -0.5822   0.05398   0.05010  -0.0229   1.0000   0.0083
  -6.250  -0.5623   0.04844   0.04432  -0.0261   1.0000   0.0080
  -6.000  -0.5396   0.04307   0.03865  -0.0287   1.0000   0.0078
  -5.750  -0.5145   0.03790   0.03311  -0.0308   1.0000   0.0076
  -5.500  -0.4873   0.03295   0.02773  -0.0323   1.0000   0.0074
  -5.250  -0.4583   0.02833   0.02259  -0.0334   1.0000   0.0075
  -5.000  -0.4282   0.02420   0.01790  -0.0340   1.0000   0.0078
  -4.750  -0.3984   0.02169   0.01480  -0.0341   1.0000   0.0096
  -4.500  -0.3701   0.01905   0.01175  -0.0349   1.0000   0.0133
  -4.250  -0.3413   0.01683   0.00918  -0.0347   1.0000   0.0138
  -4.000  -0.3131   0.01510   0.00721  -0.0345   1.0000   0.0147
  -3.750  -0.2851   0.01371   0.00568  -0.0345   1.0000   0.0168
  -3.500  -0.2571   0.01292   0.00476  -0.0345   1.0000   0.0206
  -3.250  -0.2285   0.01194   0.00369  -0.0350   1.0000   0.0250
  -3.000  -0.1999   0.01130   0.00292  -0.0352   1.0000   0.0298
  -2.750  -0.1711   0.01067   0.00230  -0.0355   1.0000   0.0549
  -2.500  -0.1418   0.00962   0.00181  -0.0367   1.0000   0.2199
  -2.250  -0.1153   0.00792   0.00187  -0.0373   1.0000   0.6695
  -2.000  -0.0908   0.00778   0.00185  -0.0363   1.0000   0.7293
  -1.750  -0.0676   0.00768   0.00186  -0.0349   1.0000   0.7803
  -1.500  -0.0458   0.00759   0.00186  -0.0331   1.0000   0.8208
  -1.250  -0.0229   0.00751   0.00182  -0.0318   1.0000   0.8439
  -1.000   0.0009   0.00743   0.00175  -0.0307   1.0000   0.8609
  -0.750   0.0248   0.00737   0.00170  -0.0298   1.0000   0.8769
  -0.500   0.0483   0.00729   0.00166  -0.0288   1.0000   0.8922
  -0.250   0.0714   0.00721   0.00163  -0.0276   1.0000   0.9097
   0.000   0.0945   0.00712   0.00161  -0.0265   1.0000   0.9322
   0.250   0.1185   0.00701   0.00160  -0.0257   1.0000   0.9744
   0.500   0.1451   0.00703   0.00170  -0.0257   1.0000   1.0000
   0.750   0.1738   0.00710   0.00183  -0.0261   1.0000   1.0000
   1.000   0.2167   0.00711   0.00196  -0.0295   0.9846   1.0000
   1.250   0.2719   0.00773   0.00174  -0.0335   0.6390   1.0000
   1.500   0.2892   0.01039   0.00212  -0.0326   0.1060   1.0000
   1.750   0.3170   0.01107   0.00253  -0.0329   0.0390   1.0000
   2.000   0.3454   0.01167   0.00330  -0.0330   0.0263   1.0000
   2.250   0.3734   0.01237   0.00412  -0.0330   0.0232   1.0000
   2.500   0.4003   0.01356   0.00541  -0.0328   0.0212   1.0000
   2.750   0.4280   0.01427   0.00622  -0.0327   0.0170   1.0000
   3.000   0.4554   0.01571   0.00782  -0.0323   0.0146   1.0000
   3.250   0.4834   0.01755   0.00991  -0.0318   0.0135   1.0000
   3.500   0.5116   0.01984   0.01257  -0.0312   0.0129   1.0000
   3.750   0.5371   0.02283   0.01599  -0.0310   0.0092   1.0000
   4.000   0.5654   0.02454   0.01809  -0.0304   0.0075   1.0000
   4.250   0.5921   0.02862   0.02286  -0.0295   0.0070   1.0000
   4.500   0.6174   0.03296   0.02770  -0.0287   0.0069   1.0000
   4.750   0.6414   0.03756   0.03274  -0.0282   0.0070   1.0000
   5.000   0.6639   0.04239   0.03794  -0.0281   0.0071   1.0000
   5.250   0.6850   0.04740   0.04328  -0.0284   0.0073   1.0000
   5.500   0.7044   0.05259   0.04874  -0.0292   0.0076   1.0000
   5.750   0.7218   0.05793   0.05431  -0.0305   0.0078   1.0000
   6.000   0.7368   0.06343   0.06000  -0.0324   0.0081   1.0000
   6.250   0.7489   0.06914   0.06586  -0.0348   0.0085   1.0000
   6.500   0.7491   0.07611   0.07290  -0.0371   0.0094   1.0000
   7.500   0.7957   0.09423   0.09130  -0.0506   0.0104   1.0000
   7.750   0.7974   0.09949   0.09655  -0.0559   0.0106   1.0000
   8.000   0.7977   0.10449   0.10151  -0.0600   0.0107   1.0000
   8.250   0.7983   0.10924   0.10624  -0.0632   0.0108   1.0000
<< Back to NASA SC(2)-0404 AIRFOIL (sc20404-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0404 AIRFOIL (sc20404-il)