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NASA SC(2)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 100,000
Max Cl/Cd: 32.26 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20406-il-100000-n5.txt
Download as CSV file: xf-sc20406-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6684   0.10426   0.09922   0.0124   1.0000   0.0335
  -9.250  -0.6720   0.09797   0.09300   0.0083   1.0000   0.0304
  -9.000  -0.6755   0.09001   0.08507   0.0011   1.0000   0.0266
  -8.750  -0.6763   0.08498   0.08008  -0.0020   1.0000   0.0261
  -8.500  -0.6779   0.07893   0.07404  -0.0079   1.0000   0.0257
  -8.250  -0.6771   0.07279   0.06784  -0.0134   1.0000   0.0252
  -8.000  -0.6746   0.06683   0.06175  -0.0178   1.0000   0.0247
  -7.750  -0.6692   0.06092   0.05566  -0.0214   1.0000   0.0242
  -7.500  -0.6605   0.05509   0.04955  -0.0243   1.0000   0.0237
  -7.250  -0.6480   0.04943   0.04353  -0.0266   1.0000   0.0233
  -7.000  -0.6317   0.04408   0.03774  -0.0283   1.0000   0.0229
  -6.750  -0.6121   0.03912   0.03225  -0.0295   1.0000   0.0227
  -6.500  -0.5895   0.03470   0.02726  -0.0304   1.0000   0.0226
  -6.250  -0.5648   0.03093   0.02290  -0.0309   1.0000   0.0229
  -6.000  -0.5387   0.02778   0.01921  -0.0311   1.0000   0.0235
  -5.750  -0.5120   0.02521   0.01618  -0.0310   1.0000   0.0246
  -5.500  -0.4848   0.02367   0.01414  -0.0308   1.0000   0.0274
  -5.250  -0.4593   0.02163   0.01203  -0.0308   1.0000   0.0304
  -5.000  -0.4330   0.02009   0.01036  -0.0306   1.0000   0.0331
  -4.750  -0.4065   0.01867   0.00879  -0.0302   1.0000   0.0373
  -4.500  -0.3801   0.01741   0.00751  -0.0302   1.0000   0.0452
  -4.250  -0.3533   0.01671   0.00678  -0.0303   1.0000   0.0631
  -4.000  -0.3262   0.01578   0.00576  -0.0303   1.0000   0.0793
  -3.750  -0.2990   0.01495   0.00500  -0.0306   1.0000   0.1045
  -3.500  -0.2715   0.01403   0.00432  -0.0309   1.0000   0.1516
  -3.250  -0.2453   0.01224   0.00380  -0.0319   1.0000   0.4108
  -3.000  -0.2230   0.01158   0.00384  -0.0305   1.0000   0.6034
  -2.750  -0.2007   0.01142   0.00381  -0.0287   1.0000   0.6838
  -2.500  -0.1790   0.01135   0.00382  -0.0267   1.0000   0.7375
  -2.250  -0.1599   0.01132   0.00385  -0.0241   1.0000   0.7841
  -2.000  -0.1431   0.01128   0.00387  -0.0208   1.0000   0.8247
  -1.750  -0.1266   0.01119   0.00381  -0.0176   1.0000   0.8570
  -1.500  -0.1067   0.01106   0.00364  -0.0156   1.0000   0.8795
  -1.250  -0.0847   0.01091   0.00346  -0.0142   1.0000   0.8971
  -1.000  -0.0615   0.01076   0.00330  -0.0132   1.0000   0.9152
  -0.750  -0.0371   0.01060   0.00314  -0.0125   1.0000   0.9340
  -0.500  -0.0090   0.01047   0.00301  -0.0127   1.0000   0.9553
  -0.250   0.0212   0.01035   0.00291  -0.0136   1.0000   0.9874
   0.000   0.0478   0.01034   0.00291  -0.0140   1.0000   1.0000
   0.250   0.0764   0.01040   0.00298  -0.0148   1.0000   1.0000
   0.500   0.1049   0.01049   0.00311  -0.0155   1.0000   1.0000
   0.750   0.1332   0.01061   0.00328  -0.0162   1.0000   1.0000
   1.000   0.1612   0.01075   0.00350  -0.0168   1.0000   1.0000
   1.250   0.1891   0.01091   0.00376  -0.0174   1.0000   1.0000
   1.500   0.2530   0.01088   0.00396  -0.0247   0.9693   1.0000
   1.750   0.3084   0.01069   0.00393  -0.0292   0.9015   1.0000
   2.000   0.3461   0.01073   0.00385  -0.0296   0.7817   1.0000
   2.250   0.3662   0.01177   0.00377  -0.0264   0.5002   1.0000
   2.500   0.3877   0.01355   0.00430  -0.0259   0.2176   1.0000
   2.750   0.4132   0.01473   0.00493  -0.0260   0.1159   1.0000
   3.000   0.4401   0.01561   0.00570  -0.0259   0.0877   1.0000
   3.250   0.4670   0.01647   0.00655  -0.0258   0.0667   1.0000
   3.500   0.4936   0.01734   0.00739  -0.0258   0.0484   1.0000
   3.750   0.5203   0.01830   0.00843  -0.0257   0.0367   1.0000
   4.000   0.5462   0.01973   0.00993  -0.0254   0.0318   1.0000
   4.250   0.5731   0.02128   0.01167  -0.0250   0.0291   1.0000
   4.500   0.6002   0.02303   0.01362  -0.0246   0.0273   1.0000
   4.750   0.6271   0.02500   0.01587  -0.0243   0.0261   1.0000
   5.000   0.6536   0.02728   0.01849  -0.0239   0.0253   1.0000
   5.250   0.6789   0.02995   0.02163  -0.0235   0.0248   1.0000
   5.500   0.7025   0.03297   0.02509  -0.0231   0.0242   1.0000
   5.750   0.7227   0.03687   0.02947  -0.0228   0.0229   1.0000
   6.000   0.7440   0.04058   0.03387  -0.0221   0.0216   1.0000
   6.250   0.7613   0.04529   0.03912  -0.0217   0.0213   1.0000
   6.500   0.7758   0.05036   0.04464  -0.0217   0.0214   1.0000
   6.750   0.7876   0.05565   0.05031  -0.0220   0.0216   1.0000
   7.000   0.7965   0.06108   0.05604  -0.0227   0.0218   1.0000
   7.250   0.8027   0.06656   0.06175  -0.0238   0.0221   1.0000
   7.500   0.8060   0.07208   0.06745  -0.0254   0.0224   1.0000
   7.750   0.8067   0.07754   0.07304  -0.0273   0.0227   1.0000
   8.000   0.8047   0.08280   0.07835  -0.0289   0.0231   1.0000
   8.500   0.7936   0.10012   0.09578  -0.0462   0.0270   1.0000
   8.750   0.7900   0.10615   0.10177  -0.0502   0.0284   1.0000
   9.000   0.7886   0.11152   0.10711  -0.0531   0.0299   1.0000
   9.250   0.7893   0.11628   0.11184  -0.0545   0.0319   1.0000
   9.500   0.6483   0.11096   0.10670  -0.0399   0.0286   1.0000
   9.750   0.6448   0.11578   0.11149  -0.0413   0.0298   1.0000
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