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NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 200,000
Max Cl/Cd: 73.39 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m13-il-200000-n5.txt
Download as CSV file: xf-m13-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M13 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4491   0.10418   0.10085   0.0024   1.0000   0.0202
  -8.000  -0.4431   0.10101   0.09771  -0.0001   1.0000   0.0213
  -7.500  -0.4390   0.09188   0.08866  -0.0063   1.0000   0.0144
  -7.250  -0.4312   0.08800   0.08479  -0.0088   1.0000   0.0142
  -7.000  -0.4195   0.08351   0.08032  -0.0148   1.0000   0.0145
  -6.750  -0.4040   0.08254   0.07934  -0.0161   1.0000   0.0217
  -6.500  -0.3902   0.07834   0.07513  -0.0201   1.0000   0.0217
  -6.250  -0.3818   0.07403   0.07084  -0.0210   1.0000   0.0190
  -6.000  -0.3655   0.06918   0.06596  -0.0255   1.0000   0.0175
  -5.750  -0.3466   0.06420   0.06093  -0.0300   1.0000   0.0167
  -5.500  -0.3263   0.05960   0.05626  -0.0338   1.0000   0.0165
  -5.250  -0.3025   0.05523   0.05180  -0.0376   0.9942   0.0168
  -5.000  -0.2682   0.05112   0.04755  -0.0429   0.9703   0.0190
  -4.750  -0.2316   0.04585   0.04202  -0.0483   0.9483   0.0202
  -4.500  -0.1962   0.04035   0.03619  -0.0524   0.9277   0.0210
  -4.250  -0.1647   0.03558   0.03106  -0.0548   0.9083   0.0236
  -4.000  -0.1409   0.03434   0.02965  -0.0551   0.8875   0.0261
  -3.750  -0.1131   0.03053   0.02545  -0.0556   0.8700   0.0272
  -3.500  -0.0837   0.02626   0.02057  -0.0556   0.8544   0.0304
  -3.250  -0.0572   0.02246   0.01614  -0.0554   0.8395   0.0318
  -3.000  -0.0316   0.02080   0.01416  -0.0552   0.8240   0.0331
  -2.750  -0.0058   0.02012   0.01330  -0.0550   0.8090   0.0356
  -2.500   0.0214   0.01819   0.01090  -0.0546   0.7956   0.0363
  -2.250   0.0487   0.01668   0.00897  -0.0542   0.7824   0.0369
  -2.000   0.0760   0.01559   0.00757  -0.0538   0.7697   0.0376
  -1.750   0.1031   0.01477   0.00650  -0.0535   0.7574   0.0382
  -1.500   0.1303   0.01421   0.00573  -0.0532   0.7457   0.0394
  -1.250   0.1573   0.01370   0.00506  -0.0530   0.7346   0.0403
  -1.000   0.1843   0.01323   0.00446  -0.0527   0.7240   0.0401
  -0.750   0.2114   0.01279   0.00392  -0.0525   0.7131   0.0401
  -0.500   0.2385   0.01245   0.00351  -0.0523   0.7028   0.0400
  -0.250   0.2657   0.01219   0.00318  -0.0521   0.6929   0.0400
   0.000   0.2930   0.01200   0.00291  -0.0520   0.6828   0.0401
   0.500   0.3480   0.01176   0.00254  -0.0518   0.6634   0.0408
   0.750   0.3755   0.01169   0.00243  -0.0517   0.6536   0.0418
   1.000   0.4031   0.01164   0.00237  -0.0516   0.6438   0.0435
   1.250   0.4306   0.01161   0.00237  -0.0515   0.6344   0.0479
   1.500   0.4582   0.01162   0.00238  -0.0515   0.6246   0.0667
   1.750   0.4858   0.01164   0.00241  -0.0514   0.6146   0.0793
   2.000   0.5133   0.01167   0.00243  -0.0514   0.6050   0.0890
   2.500   0.5623   0.01028   0.00264  -0.0506   0.5853   0.7268
   2.750   0.5972   0.00990   0.00272  -0.0517   0.5745   1.0000
   3.000   0.6242   0.01005   0.00281  -0.0515   0.5643   1.0000
   3.250   0.6513   0.01018   0.00293  -0.0514   0.5530   1.0000
   3.500   0.6784   0.01033   0.00309  -0.0513   0.5415   1.0000
   3.750   0.7054   0.01048   0.00324  -0.0512   0.5301   1.0000
   4.000   0.7323   0.01065   0.00342  -0.0510   0.5194   1.0000
   4.250   0.7592   0.01083   0.00364  -0.0509   0.5086   1.0000
   4.500   0.7862   0.01101   0.00388  -0.0508   0.4979   1.0000
   4.750   0.8130   0.01120   0.00413  -0.0507   0.4861   1.0000
   5.000   0.8388   0.01143   0.00433  -0.0503   0.4593   1.0000
   5.250   0.8635   0.01178   0.00456  -0.0498   0.4152   1.0000
   5.500   0.8864   0.01240   0.00486  -0.0493   0.3355   1.0000
   5.750   0.9005   0.01469   0.00592  -0.0484   0.1205   1.0000
   6.000   0.9183   0.01650   0.00721  -0.0475   0.0293   1.0000
   6.250   0.9411   0.01738   0.00817  -0.0468   0.0225   1.0000
   6.500   0.9642   0.01813   0.00910  -0.0462   0.0187   1.0000
   6.750   0.9858   0.01909   0.01023  -0.0454   0.0167   1.0000
   7.000   1.0050   0.02035   0.01163  -0.0444   0.0151   1.0000
   7.250   1.0181   0.02246   0.01385  -0.0425   0.0131   1.0000
   7.500   1.0383   0.02345   0.01496  -0.0416   0.0123   1.0000
   7.750   1.0562   0.02491   0.01654  -0.0402   0.0117   1.0000
   8.000   1.0741   0.02653   0.01833  -0.0388   0.0111   1.0000
   8.250   1.0926   0.02832   0.02026  -0.0375   0.0107   1.0000
   8.500   1.1115   0.03024   0.02235  -0.0363   0.0102   1.0000
   8.750   1.1298   0.03180   0.02405  -0.0352   0.0095   1.0000
   9.000   1.1463   0.03328   0.02566  -0.0343   0.0087   1.0000
   9.250   1.1603   0.03597   0.02850  -0.0330   0.0081   1.0000
   9.500   1.1719   0.03962   0.03247  -0.0314   0.0079   1.0000
   9.750   1.1820   0.04225   0.03545  -0.0297   0.0078   1.0000
  10.000   1.1882   0.04518   0.03874  -0.0277   0.0077   1.0000
  10.250   1.1898   0.04833   0.04223  -0.0255   0.0076   1.0000
  10.500   1.1861   0.05149   0.04570  -0.0230   0.0076   1.0000
  10.750   1.1768   0.05461   0.04908  -0.0203   0.0076   1.0000
  11.000   1.1654   0.05806   0.05279  -0.0185   0.0076   1.0000
  11.250   1.1525   0.06192   0.05694  -0.0177   0.0076   1.0000
  11.500   1.1384   0.06625   0.06149  -0.0179   0.0076   1.0000
  11.750   1.1235   0.07105   0.06650  -0.0190   0.0077   1.0000
  12.000   1.1078   0.07635   0.07198  -0.0208   0.0077   1.0000
  12.250   1.0914   0.08218   0.07799  -0.0234   0.0077   1.0000
  12.500   1.0746   0.08851   0.08443  -0.0267   0.0077   1.0000
  12.750   1.0579   0.09534   0.09139  -0.0305   0.0078   1.0000
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