Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 1,000,000
Max Cl/Cd: 75.27 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m10-il-1000000-n5.txt
Download as CSV file: xf-m10-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6382   0.09655   0.09505   0.0155   1.0000   0.0033
  -9.000  -0.6412   0.09138   0.08991   0.0125   1.0000   0.0033
  -8.750  -0.7898   0.02400   0.02080  -0.0265   1.0000   0.0030
  -8.500  -0.7726   0.01976   0.01607  -0.0261   1.0000   0.0032
  -8.250  -0.7500   0.01782   0.01382  -0.0259   1.0000   0.0033
  -8.000  -0.7257   0.01651   0.01231  -0.0257   1.0000   0.0035
  -7.750  -0.7006   0.01547   0.01112  -0.0255   1.0000   0.0037
  -7.500  -0.6751   0.01455   0.01004  -0.0253   1.0000   0.0039
  -7.250  -0.6492   0.01375   0.00910  -0.0252   1.0000   0.0042
  -7.000  -0.6231   0.01305   0.00827  -0.0250   1.0000   0.0045
  -6.750  -0.5966   0.01248   0.00759  -0.0249   1.0000   0.0047
  -6.500  -0.5705   0.01164   0.00662  -0.0247   1.0000   0.0053
  -6.250  -0.5439   0.01106   0.00597  -0.0246   1.0000   0.0058
  -6.000  -0.5171   0.01058   0.00541  -0.0245   1.0000   0.0064
  -5.750  -0.4902   0.01015   0.00488  -0.0243   1.0000   0.0070
  -5.500  -0.4633   0.00973   0.00439  -0.0242   1.0000   0.0080
  -5.250  -0.4365   0.00932   0.00399  -0.0240   1.0000   0.0099
  -5.000  -0.4067   0.00903   0.00371  -0.0246   0.9941   0.0134
  -4.750  -0.3742   0.00892   0.00365  -0.0257   0.9803   0.0174
  -4.500  -0.3425   0.00880   0.00348  -0.0266   0.9620   0.0187
  -4.250  -0.3150   0.00862   0.00323  -0.0264   0.9370   0.0200
  -4.000  -0.2897   0.00853   0.00307  -0.0257   0.9098   0.0222
  -3.750  -0.2638   0.00848   0.00294  -0.0252   0.8847   0.0240
  -3.500  -0.2372   0.00844   0.00281  -0.0248   0.8618   0.0263
  -3.250  -0.2102   0.00843   0.00270  -0.0246   0.8403   0.0282
  -3.000  -0.1829   0.00839   0.00258  -0.0244   0.8196   0.0293
  -2.750  -0.1555   0.00834   0.00244  -0.0243   0.8003   0.0299
  -2.500  -0.1279   0.00827   0.00228  -0.0242   0.7822   0.0303
  -2.250  -0.1003   0.00820   0.00213  -0.0242   0.7642   0.0305
  -2.000  -0.0725   0.00811   0.00198  -0.0241   0.7469   0.0306
  -1.750  -0.0449   0.00795   0.00172  -0.0241   0.7299   0.0310
  -1.500  -0.0171   0.00782   0.00150  -0.0241   0.7136   0.0314
  -1.250   0.0109   0.00772   0.00132  -0.0241   0.6973   0.0318
  -1.000   0.0389   0.00764   0.00117  -0.0241   0.6820   0.0322
  -0.750   0.0669   0.00758   0.00105  -0.0242   0.6671   0.0329
  -0.500   0.0950   0.00753   0.00095  -0.0242   0.6521   0.0340
  -0.250   0.1231   0.00750   0.00087  -0.0243   0.6376   0.0346
   0.000   0.1513   0.00748   0.00081  -0.0243   0.6232   0.0362
   0.250   0.1793   0.00747   0.00077  -0.0244   0.6048   0.0399
   0.500   0.2071   0.00745   0.00076  -0.0244   0.5769   0.0700
   0.750   0.2349   0.00738   0.00076  -0.0245   0.5499   0.1246
   1.250   0.2895   0.00669   0.00082  -0.0250   0.5039   0.4563
   1.500   0.3051   0.00533   0.00089  -0.0223   0.4756   0.9606
   1.750   0.3465   0.00548   0.00093  -0.0255   0.4428   1.0000
   2.000   0.3735   0.00564   0.00098  -0.0253   0.4151   1.0000
   2.250   0.4006   0.00582   0.00106  -0.0253   0.3854   1.0000
   2.500   0.4270   0.00621   0.00118  -0.0252   0.3188   1.0000
   2.750   0.4535   0.00661   0.00133  -0.0252   0.2561   1.0000
   3.000   0.4800   0.00704   0.00152  -0.0251   0.1942   1.0000
   3.250   0.5059   0.00765   0.00178  -0.0251   0.1094   1.0000
   3.500   0.5322   0.00817   0.00205  -0.0251   0.0489   1.0000
   3.750   0.5594   0.00839   0.00223  -0.0250   0.0389   1.0000
   4.000   0.5867   0.00861   0.00243  -0.0250   0.0344   1.0000
   4.250   0.6139   0.00883   0.00266  -0.0250   0.0305   1.0000
   4.500   0.6413   0.00899   0.00284  -0.0250   0.0293   1.0000
   4.750   0.6686   0.00917   0.00304  -0.0250   0.0268   1.0000
   5.000   0.6956   0.00943   0.00327  -0.0249   0.0188   1.0000
   5.250   0.7225   0.00970   0.00352  -0.0249   0.0146   1.0000
   5.500   0.7493   0.01003   0.00384  -0.0248   0.0104   1.0000
   5.750   0.7760   0.01031   0.00416  -0.0247   0.0090   1.0000
   6.000   0.8024   0.01069   0.00455  -0.0246   0.0073   1.0000
   6.250   0.8287   0.01109   0.00500  -0.0245   0.0065   1.0000
   6.500   0.8549   0.01147   0.00543  -0.0243   0.0059   1.0000
   6.750   0.8809   0.01188   0.00588  -0.0242   0.0054   1.0000
   7.000   0.9066   0.01233   0.00637  -0.0240   0.0049   1.0000
   7.250   0.9311   0.01303   0.00717  -0.0236   0.0043   1.0000
   7.500   0.9565   0.01349   0.00769  -0.0234   0.0041   1.0000
   7.750   0.9816   0.01400   0.00827  -0.0231   0.0038   1.0000
   8.000   1.0062   0.01459   0.00894  -0.0228   0.0036   1.0000
   8.250   1.0304   0.01522   0.00964  -0.0224   0.0033   1.0000
   8.500   1.0544   0.01585   0.01035  -0.0221   0.0031   1.0000
   8.750   1.0780   0.01654   0.01111  -0.0216   0.0030   1.0000
   9.000   1.1007   0.01739   0.01205  -0.0211   0.0028   1.0000
   9.250   1.1203   0.01880   0.01364  -0.0202   0.0026   1.0000
   9.500   1.1417   0.01982   0.01480  -0.0196   0.0026   1.0000
   9.750   1.1627   0.02086   0.01604  -0.0189   0.0025   1.0000
  10.000   1.1825   0.02207   0.01742  -0.0180   0.0024   1.0000
  10.250   1.2008   0.02348   0.01903  -0.0171   0.0023   1.0000
  10.500   1.2173   0.02511   0.02087  -0.0160   0.0022   1.0000
  10.750   1.2317   0.02698   0.02299  -0.0147   0.0022   1.0000
  11.000   1.2433   0.02916   0.02542  -0.0133   0.0021   1.0000
  11.250   1.2512   0.03164   0.02819  -0.0116   0.0020   1.0000
  11.500   1.2553   0.03437   0.03119  -0.0098   0.0020   1.0000
  11.750   1.2539   0.03740   0.03450  -0.0077   0.0019   1.0000
  12.000   1.2416   0.04060   0.03795  -0.0046   0.0019   1.0000
  12.250   1.2264   0.04434   0.04191  -0.0031   0.0019   1.0000
  12.500   1.2029   0.05006   0.04788  -0.0042   0.0019   1.0000
  12.750   1.1687   0.05901   0.05712  -0.0092   0.0019   1.0000
  13.000   1.1296   0.07126   0.06962  -0.0181   0.0020   1.0000
  13.250   1.0640   0.09480   0.09341  -0.0336   0.0021   1.0000
<< Back to NACA M10 AIRFOIL (m10-il)

Polar data table (+)

Polar graphs


<< Back to NACA M10 AIRFOIL (m10-il)