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USA 35 A AIRFOIL (usa35a-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 35 A AIRFOIL (usa35a-il)
Reynolds number: 50,000
Max Cl/Cd: 20.39 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa35a-il-50000-n5.txt
Download as CSV file: xf-usa35a-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 35 A AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250   0.0132   0.11608   0.10894  -0.0828   0.9114   0.1490
  -9.000   0.0084   0.11434   0.10722  -0.0865   0.9001   0.1537
  -8.750   0.0031   0.11224   0.10516  -0.0894   0.8879   0.1546
  -7.750   0.0342   0.09264   0.08534  -0.0989   0.8489   0.0970
  -7.500   0.0452   0.08951   0.08219  -0.1001   0.8399   0.0957
  -7.250   0.0445   0.08683   0.07952  -0.1004   0.8277   0.0949
  -7.000   0.0454   0.08357   0.07624  -0.1019   0.8180   0.0943
  -6.750   0.0399   0.08076   0.07343  -0.1028   0.8053   0.0937
  -6.500   0.0421   0.07667   0.06928  -0.1061   0.7966   0.0930
  -6.250   0.0352   0.07294   0.06549  -0.1081   0.7845   0.0918
  -6.000   0.0222   0.06530   0.05746  -0.1144   0.7754   0.0892
  -5.750   0.0257   0.06216   0.05415  -0.1150   0.7657   0.0889
  -5.500   0.0322   0.05860   0.05034  -0.1160   0.7566   0.0889
  -5.250   0.0517   0.05498   0.04641  -0.1181   0.7513   0.0899
  -5.000   0.0581   0.05399   0.04538  -0.1163   0.7398   0.0911
  -4.750   0.0786   0.05157   0.04271  -0.1172   0.7332   0.0930
  -4.500   0.1050   0.04832   0.03899  -0.1192   0.7290   0.0948
  -4.250   0.1100   0.04696   0.03734  -0.1174   0.7177   0.0956
  -4.000   0.1349   0.04422   0.03393  -0.1183   0.7121   0.0978
  -3.750   0.1673   0.04294   0.03265  -0.1193   0.7081   0.1006
  -3.500   0.1760   0.04262   0.03218  -0.1172   0.6977   0.1028
  -3.250   0.2031   0.04108   0.03021  -0.1175   0.6919   0.1058
  -3.000   0.2363   0.03970   0.02860  -0.1183   0.6880   0.1094
  -2.750   0.2497   0.03965   0.02850  -0.1167   0.6794   0.1124
  -2.500   0.2738   0.03892   0.02746  -0.1162   0.6728   0.1165
  -2.250   0.3063   0.03797   0.02643  -0.1167   0.6685   0.1211
  -2.000   0.3326   0.03747   0.02575  -0.1164   0.6629   0.1266
  -1.750   0.3454   0.03766   0.02590  -0.1145   0.6541   0.1304
  -1.500   0.3768   0.03701   0.02518  -0.1147   0.6492   0.1372
  -1.250   0.4143   0.03608   0.02413  -0.1156   0.6456   0.1452
  -1.000   0.4178   0.03697   0.02505  -0.1125   0.6352   0.1502
  -0.750   0.4480   0.03651   0.02452  -0.1125   0.6297   0.1596
  -0.500   0.4860   0.03571   0.02362  -0.1134   0.6259   0.1734
  -0.250   0.4917   0.03664   0.02457  -0.1107   0.6164   0.1827
   0.000   0.5182   0.03643   0.02436  -0.1103   0.6104   0.2014
   0.250   0.5547   0.03558   0.02363  -0.1112   0.6065   0.2351
   0.500   0.5669   0.03609   0.02437  -0.1094   0.5987   0.2734
   0.750   0.5838   0.03616   0.02493  -0.1083   0.5916   0.3587
   1.000   0.6094   0.03498   0.02470  -0.1069   0.5875   0.5891
   1.250   0.6517   0.03374   0.02383  -0.1070   0.5842   1.0000
   1.500   0.6377   0.03602   0.02611  -0.1024   0.5728   1.0000
   1.750   0.6680   0.03613   0.02593  -0.1025   0.5679   1.0000
   2.000   0.7078   0.03581   0.02529  -0.1036   0.5645   1.0000
   2.250   0.6902   0.03833   0.02785  -0.0987   0.5532   1.0000
   2.500   0.7160   0.03868   0.02800  -0.0983   0.5478   1.0000
   2.750   0.7557   0.03839   0.02745  -0.0994   0.5444   1.0000
   3.250   0.7554   0.04199   0.03096  -0.0937   0.5271   1.0000
   3.500   0.7926   0.04166   0.03043  -0.0943   0.5237   1.0000
   3.750   0.8371   0.04105   0.02959  -0.0957   0.5213   1.0000
   4.000   0.7854   0.04644   0.03518  -0.0893   0.5056   1.0000
   4.250   0.8208   0.04608   0.03466  -0.0895   0.5026   1.0000
   4.500   0.8626   0.04536   0.03375  -0.0903   0.5004   1.0000
   5.000   0.8422   0.05160   0.04004  -0.0853   0.4810   1.0000
   5.250   0.8803   0.05095   0.03925  -0.0855   0.4792   1.0000
   5.750   0.8573   0.05839   0.04676  -0.0819   0.4600   1.0000
   6.750   0.8239   0.07477   0.06324  -0.0789   0.4248   1.0000
   7.000   0.8512   0.07481   0.06320  -0.0783   0.4226   1.0000
   7.500   0.8340   0.08349   0.07196  -0.0776   0.4067   1.0000
   7.750   0.8623   0.08340   0.07179  -0.0770   0.4047   1.0000
   8.000   0.8252   0.09116   0.07967  -0.0771   0.3922   1.0000
   8.250   0.8495   0.09149   0.07995  -0.0766   0.3892   1.0000
   8.750   0.8480   0.09801   0.08651  -0.0761   0.3744   1.0000
   9.000   0.8758   0.09772   0.08616  -0.0753   0.3715   1.0000
   9.500   0.8745   0.10423   0.09272  -0.0750   0.3563   1.0000
   9.750   0.9026   0.10384   0.09228  -0.0742   0.3538   1.0000
  10.000   0.8767   0.11051   0.09904  -0.0750   0.3421   1.0000
  10.250   0.8976   0.11105   0.09957  -0.0744   0.3385   1.0000
  10.500   0.9250   0.11068   0.09917  -0.0736   0.3362   1.0000
  10.750   0.8970   0.11787   0.10646  -0.0748   0.3239   1.0000
  11.000   0.9186   0.11827   0.10684  -0.0742   0.3207   1.0000
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