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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 200,000
Max Cl/Cd: 54.94 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m10-il-200000-n5.txt
Download as CSV file: xf-m10-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M10 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5816   0.09642   0.09302   0.0099   1.0000   0.0131
  -8.250  -0.5820   0.09164   0.08828   0.0069   1.0000   0.0128
  -8.000  -0.5838   0.08661   0.08330   0.0031   1.0000   0.0125
  -7.750  -0.5817   0.08068   0.07738  -0.0029   1.0000   0.0123
  -7.500  -0.5756   0.07455   0.07123  -0.0085   1.0000   0.0122
  -7.250  -0.5664   0.06887   0.06550  -0.0132   1.0000   0.0122
  -7.000  -0.5543   0.06369   0.06023  -0.0170   1.0000   0.0124
  -6.750  -0.5387   0.06037   0.05683  -0.0192   1.0000   0.0133
  -6.500  -0.5221   0.05583   0.05217  -0.0219   1.0000   0.0143
  -6.250  -0.5047   0.05000   0.04612  -0.0246   1.0000   0.0150
  -6.000  -0.4864   0.04348   0.03929  -0.0266   1.0000   0.0153
  -5.750  -0.4670   0.03634   0.03168  -0.0277   1.0000   0.0159
  -5.500  -0.4476   0.02739   0.02185  -0.0276   1.0000   0.0176
  -5.250  -0.4248   0.02340   0.01722  -0.0272   1.0000   0.0199
  -5.000  -0.4001   0.02259   0.01629  -0.0271   1.0000   0.0217
  -4.750  -0.3749   0.02111   0.01447  -0.0267   1.0000   0.0246
  -4.500  -0.3489   0.01960   0.01252  -0.0262   1.0000   0.0290
  -4.250  -0.3233   0.01813   0.01068  -0.0257   1.0000   0.0314
  -4.000  -0.2984   0.01740   0.00993  -0.0255   1.0000   0.0348
  -3.750  -0.2729   0.01654   0.00888  -0.0251   1.0000   0.0373
  -3.500  -0.2476   0.01587   0.00803  -0.0245   1.0000   0.0402
  -3.250  -0.2226   0.01504   0.00704  -0.0239   1.0000   0.0414
  -3.000  -0.1983   0.01419   0.00610  -0.0232   1.0000   0.0426
  -2.750  -0.1746   0.01335   0.00525  -0.0225   1.0000   0.0445
  -2.500  -0.1416   0.01279   0.00464  -0.0238   0.9933   0.0464
  -2.250  -0.1063   0.01231   0.00413  -0.0256   0.9825   0.0482
  -2.000  -0.0712   0.01196   0.00376  -0.0273   0.9698   0.0516
  -1.750  -0.0367   0.01161   0.00335  -0.0287   0.9554   0.0539
  -1.500  -0.0029   0.01129   0.00297  -0.0300   0.9394   0.0551
  -1.250   0.0290   0.01095   0.00262  -0.0308   0.9211   0.0589
  -1.000   0.0590   0.01073   0.00237  -0.0310   0.9013   0.0671
  -0.750   0.0869   0.01045   0.00218  -0.0309   0.8811   0.0985
  -0.500   0.1128   0.00998   0.00206  -0.0306   0.8605   0.2095
  -0.250   0.1259   0.00783   0.00211  -0.0269   0.8421   0.8762
   0.250   0.2052   0.00785   0.00197  -0.0314   0.8037   1.0000
   0.500   0.2303   0.00794   0.00192  -0.0307   0.7846   1.0000
   0.750   0.2559   0.00803   0.00190  -0.0300   0.7657   1.0000
   1.000   0.2816   0.00813   0.00190  -0.0295   0.7474   1.0000
   1.250   0.3075   0.00823   0.00192  -0.0289   0.7298   1.0000
   1.500   0.3336   0.00835   0.00196  -0.0284   0.7124   1.0000
   1.750   0.3598   0.00846   0.00202  -0.0280   0.6948   1.0000
   2.000   0.3861   0.00859   0.00209  -0.0276   0.6772   1.0000
   2.250   0.4122   0.00873   0.00218  -0.0271   0.6550   1.0000
   2.500   0.4379   0.00891   0.00226  -0.0266   0.6247   1.0000
   2.750   0.4636   0.00911   0.00234  -0.0261   0.5893   1.0000
   3.000   0.4895   0.00932   0.00248  -0.0256   0.5568   1.0000
   3.250   0.5154   0.00956   0.00263  -0.0252   0.5208   1.0000
   3.500   0.5409   0.00987   0.00279  -0.0247   0.4729   1.0000
   3.750   0.5659   0.01030   0.00301  -0.0243   0.4080   1.0000
   4.000   0.5904   0.01088   0.00329  -0.0239   0.3301   1.0000
   4.250   0.6130   0.01194   0.00376  -0.0236   0.1968   1.0000
   4.500   0.6350   0.01330   0.00449  -0.0233   0.0743   1.0000
   4.750   0.6601   0.01398   0.00511  -0.0230   0.0546   1.0000
   5.000   0.6854   0.01455   0.00576  -0.0227   0.0444   1.0000
   5.250   0.7102   0.01523   0.00650  -0.0223   0.0358   1.0000
   5.500   0.7338   0.01616   0.00749  -0.0218   0.0295   1.0000
   5.750   0.7587   0.01679   0.00826  -0.0214   0.0242   1.0000
   6.000   0.7821   0.01775   0.00932  -0.0208   0.0207   1.0000
   6.250   0.8037   0.01911   0.01076  -0.0199   0.0179   1.0000
   6.500   0.8270   0.02017   0.01197  -0.0192   0.0160   1.0000
   6.750   0.8494   0.02163   0.01358  -0.0183   0.0146   1.0000
   7.000   0.8719   0.02322   0.01535  -0.0174   0.0135   1.0000
   7.250   0.8945   0.02440   0.01666  -0.0168   0.0122   1.0000
   7.500   0.9137   0.02684   0.01936  -0.0159   0.0108   1.0000
   7.750   0.9329   0.02966   0.02261  -0.0148   0.0103   1.0000
   8.000   0.9510   0.03242   0.02580  -0.0135   0.0100   1.0000
   8.250   0.9658   0.03576   0.02961  -0.0121   0.0097   1.0000
   8.500   0.9765   0.03964   0.03398  -0.0106   0.0096   1.0000
   8.750   0.9825   0.04399   0.03880  -0.0090   0.0095   1.0000
   9.000   0.9836   0.04859   0.04382  -0.0074   0.0094   1.0000
   9.250   0.9806   0.05318   0.04878  -0.0060   0.0093   1.0000
   9.500   0.9727   0.05774   0.05364  -0.0049   0.0092   1.0000
   9.750   0.9578   0.06202   0.05815  -0.0039   0.0092   1.0000
  10.000   0.9395   0.06677   0.06308  -0.0048   0.0092   1.0000
  10.250   0.9205   0.07286   0.06932  -0.0087   0.0094   1.0000
  10.500   0.9015   0.08071   0.07729  -0.0152   0.0096   1.0000
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