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S1012 (s1012-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: S1012 (s1012-il)
Reynolds number: 50,000
Max Cl/Cd: 26.21 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1012-il-50000.txt
Download as CSV file: xf-s1012-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S1012                                           
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4897   0.11129   0.10345   0.0007   1.0000   0.3841
 -10.000  -0.6990   0.07928   0.07189  -0.0334   1.0000   0.1880
  -9.750  -0.7166   0.07393   0.06650  -0.0336   1.0000   0.1836
  -9.500  -0.7406   0.06921   0.06175  -0.0322   1.0000   0.1808
  -9.250  -0.8096   0.06314   0.05524  -0.0281   1.0000   0.1720
  -9.000  -0.8252   0.05945   0.05124  -0.0247   1.0000   0.1709
  -8.750  -0.8054   0.05682   0.04878  -0.0238   1.0000   0.1770
  -8.500  -0.8144   0.05394   0.04569  -0.0201   1.0000   0.1788
  -8.250  -0.8233   0.05122   0.04270  -0.0160   1.0000   0.1804
  -8.000  -0.8307   0.04854   0.03968  -0.0118   1.0000   0.1822
  -7.750  -0.8358   0.04597   0.03669  -0.0076   1.0000   0.1843
  -7.500  -0.8331   0.04359   0.03402  -0.0043   1.0000   0.1889
  -7.250  -0.8218   0.04171   0.03208  -0.0020   1.0000   0.1956
  -7.000  -0.8164   0.03952   0.02941   0.0012   1.0000   0.2005
  -6.750  -0.8025   0.03755   0.02737   0.0032   1.0000   0.2078
  -6.500  -0.7913   0.03586   0.02529   0.0058   1.0000   0.2169
  -6.250  -0.7729   0.03407   0.02350   0.0072   1.0000   0.2258
  -6.000  -0.7565   0.03249   0.02170   0.0090   1.0000   0.2372
  -5.750  -0.7387   0.03108   0.02011   0.0107   1.0000   0.2507
  -5.500  -0.7185   0.02970   0.01867   0.0121   1.0000   0.2659
  -5.250  -0.6966   0.02841   0.01750   0.0132   1.0000   0.2850
  -5.000  -0.6737   0.02720   0.01636   0.0143   1.0000   0.3083
  -4.750  -0.6509   0.02603   0.01539   0.0154   1.0000   0.3391
  -4.500  -0.6289   0.02481   0.01447   0.0167   1.0000   0.3810
  -4.250  -0.6122   0.02337   0.01365   0.0191   1.0000   0.4464
  -4.000  -0.5839   0.02340   0.01564   0.0248   1.0000   0.6587
  -3.750  -0.5618   0.02645   0.01854   0.0325   1.0000   0.7914
  -3.500  -0.4092   0.03113   0.02239   0.0201   1.0000   0.8789
  -3.250  -0.2330   0.03161   0.02216  -0.0031   1.0000   0.9421
  -3.000  -0.0578   0.02879   0.01880  -0.0311   1.0000   1.0000
  -2.750  -0.0566   0.02850   0.01844  -0.0285   1.0000   1.0000
  -2.500  -0.0545   0.02826   0.01814  -0.0258   1.0000   1.0000
  -2.250  -0.0515   0.02806   0.01787  -0.0231   1.0000   1.0000
  -2.000  -0.0477   0.02789   0.01764  -0.0205   1.0000   1.0000
  -1.750  -0.0432   0.02774   0.01744  -0.0179   1.0000   1.0000
  -1.500  -0.0380   0.02762   0.01727  -0.0153   1.0000   1.0000
  -1.250  -0.0324   0.02752   0.01713  -0.0127   1.0000   1.0000
  -1.000  -0.0264   0.02743   0.01702  -0.0101   1.0000   1.0000
  -0.750  -0.0200   0.02737   0.01693  -0.0076   1.0000   1.0000
  -0.500  -0.0135   0.02732   0.01686  -0.0051   1.0000   1.0000
  -0.250  -0.0068   0.02730   0.01683  -0.0025   1.0000   1.0000
   0.000   0.0000   0.02729   0.01681   0.0000   1.0000   1.0000
   0.250   0.0068   0.02730   0.01683   0.0025   1.0000   1.0000
   0.500   0.0135   0.02732   0.01686   0.0051   1.0000   1.0000
   0.750   0.0201   0.02737   0.01692   0.0076   1.0000   1.0000
   1.000   0.0264   0.02743   0.01701   0.0101   1.0000   1.0000
   1.250   0.0324   0.02751   0.01712   0.0127   1.0000   1.0000
   1.500   0.0381   0.02761   0.01726   0.0153   1.0000   1.0000
   1.750   0.0432   0.02773   0.01744   0.0179   1.0000   1.0000
   2.000   0.0477   0.02788   0.01763   0.0205   1.0000   1.0000
   2.250   0.0515   0.02805   0.01786   0.0231   1.0000   1.0000
   2.500   0.0545   0.02825   0.01812   0.0258   1.0000   1.0000
   2.750   0.0566   0.02849   0.01842   0.0285   1.0000   1.0000
   3.000   0.0577   0.02877   0.01878   0.0311   1.0000   1.0000
   3.250   0.2308   0.03157   0.02211   0.0035   0.9425   1.0000
   3.500   0.4091   0.03112   0.02238  -0.0201   0.8788   1.0000
   3.750   0.5616   0.02645   0.01854  -0.0325   0.7917   1.0000
   4.000   0.5838   0.02340   0.01564  -0.0248   0.6594   1.0000
   4.250   0.6122   0.02336   0.01365  -0.0191   0.4465   1.0000
   4.500   0.6289   0.02481   0.01447  -0.0167   0.3809   1.0000
   4.750   0.6508   0.02603   0.01538  -0.0154   0.3390   1.0000
   5.000   0.6736   0.02720   0.01636  -0.0142   0.3081   1.0000
   5.250   0.6965   0.02840   0.01750  -0.0132   0.2849   1.0000
   5.500   0.7185   0.02971   0.01867  -0.0121   0.2657   1.0000
   5.750   0.7386   0.03107   0.02011  -0.0107   0.2507   1.0000
   6.000   0.7564   0.03248   0.02169  -0.0090   0.2371   1.0000
   6.250   0.7729   0.03407   0.02350  -0.0072   0.2258   1.0000
   6.500   0.7913   0.03586   0.02529  -0.0057   0.2169   1.0000
   6.750   0.8025   0.03755   0.02737  -0.0032   0.2079   1.0000
   7.000   0.8164   0.03952   0.02941  -0.0012   0.2005   1.0000
   7.250   0.8217   0.04171   0.03207   0.0020   0.1954   1.0000
   7.500   0.8338   0.04356   0.03396   0.0042   0.1886   1.0000
   7.750   0.8358   0.04596   0.03668   0.0076   0.1843   1.0000
   8.000   0.8306   0.04853   0.03968   0.0118   0.1821   1.0000
   8.250   0.8234   0.05120   0.04268   0.0160   0.1805   1.0000
   8.500   0.8143   0.05395   0.04569   0.0201   0.1788   1.0000
   8.750   0.8064   0.05676   0.04870   0.0238   0.1767   1.0000
   9.000   0.8249   0.05944   0.05123   0.0248   0.1709   1.0000
   9.250   0.8087   0.06315   0.05527   0.0281   0.1720   1.0000
   9.500   0.7409   0.06922   0.06176   0.0321   0.1808   1.0000
   9.750   0.7172   0.07396   0.06653   0.0335   0.1837   1.0000
  10.000   0.6499   0.08374   0.07628   0.0297   0.1997   1.0000
  10.250   0.6634   0.08861   0.08118   0.0295   0.2022   1.0000
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