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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 1,000,000
Max Cl/Cd: 80.94 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m10-il-1000000.txt
Download as CSV file: xf-m10-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5992   0.08650   0.08499   0.0080   1.0000   0.0119
  -7.750  -0.5938   0.08138   0.07987   0.0022   1.0000   0.0125
  -7.250  -0.6386   0.02449   0.02149  -0.0251   1.0000   0.0092
  -7.000  -0.6266   0.01606   0.01189  -0.0241   1.0000   0.0098
  -6.750  -0.6031   0.01389   0.00934  -0.0237   1.0000   0.0108
  -6.500  -0.5761   0.01365   0.00907  -0.0237   1.0000   0.0115
  -6.250  -0.5490   0.01332   0.00869  -0.0236   1.0000   0.0123
  -6.000  -0.5222   0.01283   0.00810  -0.0234   1.0000   0.0132
  -5.750  -0.4948   0.01265   0.00786  -0.0233   1.0000   0.0138
  -5.500  -0.4687   0.01179   0.00691  -0.0232   1.0000   0.0155
  -5.250  -0.4409   0.01202   0.00718  -0.0232   1.0000   0.0166
  -5.000  -0.4136   0.01185   0.00698  -0.0231   1.0000   0.0179
  -4.750  -0.3861   0.01183   0.00693  -0.0230   1.0000   0.0190
  -4.500  -0.3605   0.01069   0.00565  -0.0227   1.0000   0.0208
  -4.250  -0.3336   0.01062   0.00560  -0.0226   1.0000   0.0223
  -4.000  -0.3069   0.01055   0.00553  -0.0224   1.0000   0.0239
  -3.750  -0.2809   0.01026   0.00521  -0.0220   1.0000   0.0253
  -3.500  -0.2532   0.01000   0.00492  -0.0220   0.9989   0.0265
  -3.250  -0.2172   0.01010   0.00501  -0.0239   0.9931   0.0275
  -3.000  -0.1845   0.00869   0.00348  -0.0251   0.9837   0.0300
  -2.750  -0.1519   0.00830   0.00307  -0.0262   0.9688   0.0317
  -2.500  -0.1238   0.00803   0.00276  -0.0261   0.9456   0.0328
  -2.250  -0.0988   0.00781   0.00244  -0.0252   0.9196   0.0335
  -2.000  -0.0736   0.00762   0.00215  -0.0244   0.8941   0.0342
  -1.750  -0.0476   0.00747   0.00189  -0.0239   0.8702   0.0351
  -1.500  -0.0208   0.00736   0.00169  -0.0235   0.8473   0.0366
  -1.250   0.0063   0.00727   0.00150  -0.0232   0.8263   0.0380
  -1.000   0.0337   0.00721   0.00135  -0.0231   0.8060   0.0390
  -0.750   0.0613   0.00710   0.00114  -0.0229   0.7869   0.0428
  -0.500   0.0891   0.00703   0.00104  -0.0228   0.7686   0.0480
  -0.250   0.1168   0.00686   0.00097  -0.0228   0.7509   0.0907
   0.000   0.1441   0.00647   0.00093  -0.0229   0.7338   0.2291
   0.250   0.1664   0.00479   0.00090  -0.0226   0.7177   0.7847
   0.500   0.1863   0.00436   0.00094  -0.0202   0.7020   0.9690
   0.750   0.2353   0.00443   0.00095  -0.0249   0.6844   1.0000
   1.000   0.2616   0.00453   0.00093  -0.0246   0.6597   1.0000
   1.250   0.2880   0.00464   0.00093  -0.0242   0.6320   1.0000
   1.500   0.3146   0.00474   0.00094  -0.0240   0.6084   1.0000
   1.750   0.3415   0.00485   0.00097  -0.0237   0.5883   1.0000
   2.000   0.3685   0.00497   0.00100  -0.0236   0.5638   1.0000
   2.500   0.4223   0.00534   0.00111  -0.0233   0.4926   1.0000
   2.750   0.4492   0.00557   0.00119  -0.0232   0.4498   1.0000
   3.000   0.4759   0.00588   0.00130  -0.0231   0.3951   1.0000
   3.250   0.5021   0.00636   0.00147  -0.0230   0.3158   1.0000
   3.500   0.5270   0.00725   0.00178  -0.0230   0.1804   1.0000
   3.750   0.5517   0.00828   0.00224  -0.0229   0.0490   1.0000
   4.000   0.5788   0.00858   0.00250  -0.0228   0.0384   1.0000
   4.250   0.6061   0.00880   0.00274  -0.0228   0.0347   1.0000
   4.500   0.6330   0.00918   0.00312  -0.0226   0.0290   1.0000
   4.750   0.6601   0.00944   0.00343  -0.0226   0.0266   1.0000
   5.000   0.6872   0.00966   0.00365  -0.0225   0.0235   1.0000
   5.250   0.7131   0.01029   0.00432  -0.0223   0.0189   1.0000
   5.500   0.7403   0.01047   0.00451  -0.0222   0.0177   1.0000
   5.750   0.7671   0.01076   0.00483  -0.0221   0.0159   1.0000
   6.000   0.7934   0.01118   0.00526  -0.0220   0.0142   1.0000
   6.250   0.8177   0.01210   0.00628  -0.0215   0.0127   1.0000
   6.500   0.8442   0.01238   0.00660  -0.0214   0.0119   1.0000
   6.750   0.8698   0.01287   0.00713  -0.0211   0.0110   1.0000
   7.000   0.8952   0.01337   0.00769  -0.0208   0.0102   1.0000
   7.250   0.9200   0.01401   0.00836  -0.0205   0.0096   1.0000
   7.500   0.9383   0.01626   0.01083  -0.0192   0.0087   1.0000
   7.750   0.9626   0.01701   0.01169  -0.0187   0.0085   1.0000
   8.000   0.9866   0.01782   0.01260  -0.0183   0.0081   1.0000
   8.250   1.0097   0.01884   0.01375  -0.0177   0.0077   1.0000
   8.500   1.0331   0.01967   0.01470  -0.0172   0.0072   1.0000
   8.750   1.0559   0.02055   0.01568  -0.0167   0.0068   1.0000
   9.000   1.0773   0.02173   0.01701  -0.0160   0.0065   1.0000
   9.250   1.0978   0.02301   0.01844  -0.0153   0.0063   1.0000
   9.500   1.1160   0.02474   0.02036  -0.0144   0.0061   1.0000
   9.750   1.1294   0.02739   0.02330  -0.0130   0.0060   1.0000
  10.000   1.1338   0.03160   0.02796  -0.0110   0.0058   1.0000
  10.250   1.1262   0.03729   0.03419  -0.0083   0.0058   1.0000
  10.500   1.1106   0.04312   0.04046  -0.0057   0.0057   1.0000
  10.750   1.0912   0.04776   0.04539  -0.0031   0.0057   1.0000
  11.000   1.0713   0.05199   0.04981  -0.0021   0.0058   1.0000
  11.250   1.0507   0.05758   0.05559  -0.0044   0.0058   1.0000
  11.500   1.0345   0.06394   0.06211  -0.0089   0.0058   1.0000
  11.750   1.0180   0.07188   0.07019  -0.0152   0.0059   1.0000
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