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RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAE 103 AIRFOIL (rae103-il)
Reynolds number: 50,000
Max Cl/Cd: 27.78 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae103-il-50000.txt
Download as CSV file: xf-rae103-il-50000.csv

 
  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 103 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5581   0.10047   0.09338   0.0121   1.0000   0.3737
  -8.750  -0.6725   0.08000   0.07332  -0.0182   1.0000   0.1794
  -8.500  -0.7360   0.07511   0.06831  -0.0208   1.0000   0.1723
  -8.250  -0.7405   0.06687   0.05975  -0.0222   1.0000   0.1428
  -8.000  -0.7545   0.06154   0.05385  -0.0212   1.0000   0.1301
  -7.750  -0.7414   0.05642   0.04869  -0.0204   1.0000   0.1246
  -7.500  -0.7487   0.05217   0.04347  -0.0176   1.0000   0.1164
  -7.250  -0.7348   0.04793   0.03903  -0.0163   1.0000   0.1146
  -7.000  -0.7240   0.04444   0.03501  -0.0142   1.0000   0.1147
  -6.750  -0.7123   0.04159   0.03151  -0.0118   1.0000   0.1165
  -6.500  -0.6928   0.03814   0.02802  -0.0107   1.0000   0.1204
  -6.250  -0.6720   0.03527   0.02479  -0.0092   1.0000   0.1229
  -6.000  -0.6505   0.03287   0.02191  -0.0075   1.0000   0.1290
  -5.750  -0.6269   0.03054   0.01953  -0.0065   1.0000   0.1407
  -5.500  -0.5981   0.02818   0.01709  -0.0058   1.0000   0.1533
  -5.250  -0.5717   0.02618   0.01519  -0.0048   1.0000   0.1787
  -5.000  -0.5473   0.02406   0.01336  -0.0034   1.0000   0.2185
  -4.750  -0.5377   0.02140   0.01164  -0.0003   1.0000   0.3083
  -4.500  -0.5518   0.01988   0.01256   0.0116   1.0000   0.6567
  -4.250  -0.5319   0.02242   0.01499   0.0219   1.0000   0.7772
  -4.000  -0.3050   0.02792   0.01864   0.0018   1.0000   0.9084
  -3.750  -0.2248   0.02696   0.01713  -0.0080   1.0000   0.9456
  -3.500  -0.1597   0.02562   0.01543  -0.0165   1.0000   0.9728
  -3.250  -0.0973   0.02410   0.01359  -0.0250   1.0000   0.9948
  -3.000  -0.0689   0.02319   0.01255  -0.0271   1.0000   1.0000
  -2.750  -0.0540   0.02263   0.01194  -0.0265   1.0000   1.0000
  -2.500  -0.0393   0.02212   0.01140  -0.0257   1.0000   1.0000
  -2.250  -0.0250   0.02167   0.01090  -0.0248   1.0000   1.0000
  -2.000  -0.0115   0.02127   0.01049  -0.0237   1.0000   1.0000
  -1.750   0.0009   0.02092   0.01015  -0.0224   1.0000   1.0000
  -1.500   0.0117   0.02062   0.00988  -0.0208   1.0000   1.0000
  -1.250   0.0202   0.02039   0.00968  -0.0189   1.0000   1.0000
  -1.000   0.0256   0.02023   0.00957  -0.0164   1.0000   1.0000
  -0.750   0.0266   0.02016   0.00955  -0.0134   1.0000   1.0000
  -0.500   0.0216   0.02016   0.00960  -0.0094   1.0000   1.0000
  -0.250   0.0118   0.02020   0.00966  -0.0048   1.0000   1.0000
   0.000   0.0000   0.02023   0.00969   0.0000   1.0000   1.0000
   0.250  -0.0118   0.02020   0.00966   0.0048   1.0000   1.0000
   0.500  -0.0216   0.02016   0.00960   0.0094   1.0000   1.0000
   0.750  -0.0266   0.02015   0.00955   0.0134   1.0000   1.0000
   1.000  -0.0256   0.02023   0.00957   0.0165   1.0000   1.0000
   1.250  -0.0202   0.02038   0.00968   0.0189   1.0000   1.0000
   1.500  -0.0116   0.02062   0.00987   0.0208   1.0000   1.0000
   1.750  -0.0008   0.02091   0.01014   0.0224   1.0000   1.0000
   2.000   0.0116   0.02126   0.01048   0.0237   1.0000   1.0000
   2.250   0.0252   0.02166   0.01089   0.0248   1.0000   1.0000
   2.500   0.0394   0.02212   0.01139   0.0257   1.0000   1.0000
   2.750   0.0541   0.02262   0.01193   0.0265   1.0000   1.0000
   3.000   0.0691   0.02318   0.01255   0.0271   1.0000   1.0000
   3.250   0.0974   0.02409   0.01357   0.0250   0.9948   1.0000
   3.500   0.1605   0.02563   0.01544   0.0163   0.9726   1.0000
   3.750   0.2253   0.02696   0.01712   0.0080   0.9455   1.0000
   4.000   0.3053   0.02791   0.01864  -0.0018   0.9082   1.0000
   4.250   0.5319   0.02242   0.01499  -0.0219   0.7774   1.0000
   4.500   0.5518   0.01986   0.01255  -0.0115   0.6561   1.0000
   4.750   0.5376   0.02141   0.01164   0.0003   0.3077   1.0000
   5.000   0.5472   0.02407   0.01336   0.0035   0.2181   1.0000
   5.250   0.5717   0.02618   0.01519   0.0049   0.1785   1.0000
   5.500   0.5981   0.02818   0.01709   0.0058   0.1533   1.0000
   5.750   0.6269   0.03055   0.01953   0.0065   0.1406   1.0000
   6.000   0.6505   0.03287   0.02191   0.0075   0.1290   1.0000
   6.250   0.6720   0.03527   0.02479   0.0092   0.1230   1.0000
   6.500   0.6928   0.03815   0.02803   0.0107   0.1203   1.0000
   6.750   0.7125   0.04161   0.03152   0.0118   0.1166   1.0000
   7.000   0.7239   0.04444   0.03502   0.0143   0.1146   1.0000
   7.250   0.7350   0.04791   0.03900   0.0163   0.1146   1.0000
   7.500   0.7488   0.05217   0.04346   0.0176   0.1164   1.0000
   7.750   0.7414   0.05646   0.04872   0.0204   0.1247   1.0000
   8.000   0.7546   0.06152   0.05383   0.0212   0.1301   1.0000
   8.250   0.7417   0.06694   0.05982   0.0222   0.1430   1.0000
   8.500   0.7131   0.07315   0.06644   0.0212   0.1616   1.0000
   8.750   0.6726   0.08005   0.07337   0.0181   0.1796   1.0000
   9.000   0.5586   0.10051   0.09342  -0.0122   0.3737   1.0000
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