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NASA SC(2)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 1,000,000
Max Cl/Cd: 61.52 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20406-il-1000000-n5.txt
Download as CSV file: xf-sc20406-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0406 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.7385   0.09284   0.09128   0.0142   1.0000   0.0044
  -9.750  -0.7468   0.08590   0.08437   0.0097   1.0000   0.0044
  -9.250  -0.8595   0.02997   0.02722  -0.0282   1.0000   0.0044
  -9.000  -0.8396   0.02728   0.02426  -0.0289   1.0000   0.0045
  -8.750  -0.8180   0.02489   0.02161  -0.0294   1.0000   0.0047
  -8.500  -0.7949   0.02288   0.01933  -0.0299   1.0000   0.0049
  -8.250  -0.7707   0.02096   0.01715  -0.0303   1.0000   0.0051
  -8.000  -0.7457   0.01930   0.01524  -0.0306   1.0000   0.0053
  -7.750  -0.7199   0.01777   0.01349  -0.0310   1.0000   0.0054
  -7.500  -0.6936   0.01644   0.01195  -0.0313   1.0000   0.0056
  -7.250  -0.6668   0.01519   0.01051  -0.0315   1.0000   0.0057
  -7.000  -0.6397   0.01411   0.00927  -0.0318   1.0000   0.0058
  -6.750  -0.6123   0.01319   0.00820  -0.0321   1.0000   0.0060
  -6.500  -0.5846   0.01248   0.00737  -0.0323   1.0000   0.0062
  -6.250  -0.5567   0.01175   0.00654  -0.0326   1.0000   0.0063
  -6.000  -0.5286   0.01099   0.00567  -0.0329   1.0000   0.0064
  -5.750  -0.5002   0.01035   0.00494  -0.0332   1.0000   0.0064
  -5.500  -0.4716   0.00981   0.00432  -0.0336   1.0000   0.0065
  -5.250  -0.4430   0.00936   0.00378  -0.0339   1.0000   0.0066
  -5.000  -0.4144   0.00898   0.00334  -0.0342   1.0000   0.0067
  -4.750  -0.3859   0.00866   0.00298  -0.0345   1.0000   0.0069
  -4.500  -0.3571   0.00830   0.00256  -0.0348   1.0000   0.0075
  -4.250  -0.3284   0.00803   0.00226  -0.0350   1.0000   0.0089
  -4.000  -0.2999   0.00780   0.00202  -0.0352   1.0000   0.0104
  -3.750  -0.2715   0.00757   0.00183  -0.0355   1.0000   0.0166
  -3.500  -0.2432   0.00736   0.00168  -0.0357   1.0000   0.0262
  -3.250  -0.2150   0.00719   0.00155  -0.0358   1.0000   0.0335
  -3.000  -0.1870   0.00701   0.00142  -0.0360   1.0000   0.0465
  -2.750  -0.1592   0.00683   0.00133  -0.0361   1.0000   0.0625
  -2.500  -0.1316   0.00663   0.00123  -0.0362   1.0000   0.0884
  -2.250  -0.1018   0.00632   0.00113  -0.0368   0.9988   0.1429
  -2.000  -0.0662   0.00593   0.00101  -0.0389   0.9935   0.2263
  -1.750  -0.0291   0.00553   0.00090  -0.0413   0.9802   0.3156
  -1.500   0.0050   0.00500   0.00083  -0.0430   0.9498   0.4654
  -1.250   0.0317   0.00478   0.00083  -0.0427   0.9158   0.5545
  -1.000   0.0580   0.00475   0.00082  -0.0422   0.8840   0.5997
  -0.750   0.0849   0.00479   0.00082  -0.0419   0.8489   0.6225
  -0.500   0.1112   0.00498   0.00082  -0.0414   0.7783   0.6490
  -0.250   0.1385   0.00521   0.00084  -0.0413   0.7053   0.6682
   0.000   0.1658   0.00562   0.00089  -0.0413   0.5872   0.6865
   0.250   0.1927   0.00636   0.00101  -0.0415   0.3950   0.7087
   0.500   0.2202   0.00694   0.00114  -0.0418   0.2450   0.7283
   0.750   0.2478   0.00745   0.00127  -0.0420   0.1269   0.7422
   1.000   0.2759   0.00768   0.00137  -0.0422   0.0838   0.7538
   1.250   0.3041   0.00785   0.00149  -0.0423   0.0584   0.7642
   1.500   0.3323   0.00799   0.00160  -0.0424   0.0448   0.7735
   1.750   0.3605   0.00814   0.00172  -0.0424   0.0345   0.7818
   2.000   0.3886   0.00831   0.00186  -0.0425   0.0239   0.7902
   2.250   0.4165   0.00853   0.00205  -0.0425   0.0122   0.7975
   2.500   0.4446   0.00870   0.00224  -0.0425   0.0093   0.8050
   2.750   0.4726   0.00887   0.00246  -0.0425   0.0090   0.8127
   3.250   0.5281   0.00928   0.00298  -0.0424   0.0082   0.8324
   3.500   0.5557   0.00952   0.00330  -0.0423   0.0079   0.8418
   3.750   0.5831   0.00980   0.00365  -0.0422   0.0077   0.8511
   4.000   0.6103   0.01010   0.00403  -0.0421   0.0074   0.8604
   4.250   0.6373   0.01044   0.00444  -0.0419   0.0072   0.8698
   4.500   0.6640   0.01081   0.00489  -0.0416   0.0069   0.8806
   4.750   0.6903   0.01122   0.00541  -0.0413   0.0066   0.8924
   5.000   0.7159   0.01178   0.00608  -0.0408   0.0062   0.9045
   5.250   0.7399   0.01302   0.00754  -0.0401   0.0056   0.9174
   5.500   0.7636   0.01401   0.00872  -0.0392   0.0055   0.9313
   5.750   0.7858   0.01470   0.00958  -0.0379   0.0054   0.9547
   6.000   0.8103   0.01565   0.01071  -0.0373   0.0053   1.0000
   6.250   0.8358   0.01693   0.01221  -0.0370   0.0053   1.0000
   6.500   0.8614   0.01801   0.01346  -0.0367   0.0051   1.0000
   6.750   0.8865   0.01916   0.01480  -0.0365   0.0050   1.0000
   7.000   0.9102   0.02084   0.01672  -0.0360   0.0048   1.0000
   7.250   0.9318   0.02325   0.01947  -0.0354   0.0047   1.0000
   7.500   0.9502   0.02670   0.02336  -0.0344   0.0046   1.0000
   7.750   0.9180   0.05185   0.04995  -0.0314   0.0043   1.0000
   8.000   0.9121   0.06273   0.06114  -0.0332   0.0041   1.0000
   8.250   0.9052   0.07216   0.07075  -0.0369   0.0040   1.0000
   8.500   0.8963   0.08103   0.07973  -0.0428   0.0040   1.0000
   8.750   0.8805   0.08946   0.08817  -0.0516   0.0040   1.0000
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