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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 500,000
Max Cl/Cd: 64.93 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m10-il-500000-n5.txt
Download as CSV file: xf-m10-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6095   0.09411   0.09197   0.0122   1.0000   0.0053
  -8.500  -0.6086   0.08991   0.08779   0.0095   1.0000   0.0054
  -8.250  -0.6084   0.08563   0.08354   0.0063   1.0000   0.0056
  -8.000  -0.6061   0.08048   0.07840   0.0010   1.0000   0.0058
  -7.750  -0.6003   0.07441   0.07231  -0.0052   1.0000   0.0060
  -7.500  -0.5927   0.06752   0.06536  -0.0114   1.0000   0.0062
  -7.250  -0.5840   0.05919   0.05691  -0.0177   1.0000   0.0063
  -7.000  -0.6117   0.02016   0.01572  -0.0267   1.0000   0.0065
  -6.750  -0.5887   0.01762   0.01271  -0.0263   1.0000   0.0069
  -6.500  -0.5639   0.01605   0.01081  -0.0260   1.0000   0.0072
  -6.250  -0.5388   0.01458   0.00908  -0.0257   1.0000   0.0081
  -6.000  -0.5126   0.01386   0.00821  -0.0255   1.0000   0.0090
  -5.750  -0.4862   0.01329   0.00753  -0.0254   1.0000   0.0103
  -5.500  -0.4600   0.01251   0.00659  -0.0251   1.0000   0.0119
  -5.250  -0.4331   0.01224   0.00634  -0.0249   1.0000   0.0145
  -5.000  -0.4060   0.01207   0.00613  -0.0248   1.0000   0.0172
  -4.750  -0.3790   0.01203   0.00603  -0.0247   1.0000   0.0190
  -4.500  -0.3526   0.01172   0.00570  -0.0245   1.0000   0.0216
  -4.250  -0.3260   0.01162   0.00558  -0.0243   1.0000   0.0238
  -4.000  -0.2979   0.01141   0.00532  -0.0244   0.9976   0.0263
  -3.750  -0.2642   0.01111   0.00494  -0.0258   0.9871   0.0284
  -3.500  -0.2308   0.01095   0.00471  -0.0271   0.9738   0.0299
  -3.250  -0.1981   0.01076   0.00446  -0.0281   0.9569   0.0305
  -3.000  -0.1687   0.00990   0.00350  -0.0285   0.9369   0.0327
  -2.750  -0.1418   0.00957   0.00309  -0.0282   0.9134   0.0343
  -2.500  -0.1160   0.00936   0.00279  -0.0276   0.8899   0.0354
  -2.250  -0.0903   0.00919   0.00252  -0.0270   0.8675   0.0363
  -2.000  -0.0640   0.00904   0.00227  -0.0265   0.8460   0.0370
  -1.750  -0.0375   0.00891   0.00203  -0.0261   0.8259   0.0377
  -1.500  -0.0106   0.00880   0.00182  -0.0258   0.8062   0.0383
  -1.250   0.0165   0.00872   0.00166  -0.0256   0.7875   0.0396
  -1.000   0.0438   0.00865   0.00149  -0.0254   0.7695   0.0406
  -0.750   0.0713   0.00859   0.00135  -0.0253   0.7517   0.0412
  -0.500   0.0989   0.00854   0.00124  -0.0252   0.7343   0.0424
  -0.250   0.1266   0.00848   0.00114  -0.0251   0.7177   0.0501
   0.000   0.1541   0.00835   0.00109  -0.0251   0.7016   0.0869
   0.250   0.1815   0.00814   0.00107  -0.0251   0.6860   0.1653
   0.500   0.2079   0.00748   0.00106  -0.0253   0.6707   0.3999
   0.750   0.2230   0.00587   0.00112  -0.0222   0.6569   0.9407
   1.000   0.2703   0.00593   0.00111  -0.0266   0.6353   1.0000
   1.250   0.2966   0.00606   0.00111  -0.0262   0.6075   1.0000
   1.500   0.3231   0.00620   0.00113  -0.0259   0.5793   1.0000
   1.750   0.3498   0.00634   0.00117  -0.0257   0.5563   1.0000
   2.000   0.3765   0.00650   0.00121  -0.0255   0.5263   1.0000
   2.250   0.4033   0.00667   0.00129  -0.0253   0.4983   1.0000
   2.500   0.4302   0.00685   0.00137  -0.0252   0.4708   1.0000
   2.750   0.4568   0.00710   0.00147  -0.0251   0.4296   1.0000
   3.000   0.4831   0.00744   0.00162  -0.0249   0.3765   1.0000
   3.250   0.5091   0.00789   0.00180  -0.0248   0.3090   1.0000
   3.500   0.5347   0.00847   0.00206  -0.0248   0.2286   1.0000
   3.750   0.5594   0.00933   0.00244  -0.0247   0.1195   1.0000
   4.000   0.5848   0.01001   0.00285  -0.0246   0.0547   1.0000
   4.250   0.6114   0.01035   0.00315  -0.0245   0.0425   1.0000
   4.500   0.6382   0.01062   0.00345  -0.0244   0.0373   1.0000
   4.750   0.6648   0.01096   0.00380  -0.0242   0.0324   1.0000
   5.000   0.6915   0.01124   0.00412  -0.0241   0.0283   1.0000
   5.250   0.7181   0.01153   0.00445  -0.0240   0.0224   1.0000
   5.500   0.7443   0.01194   0.00487  -0.0239   0.0170   1.0000
   5.750   0.7704   0.01235   0.00528  -0.0237   0.0135   1.0000
   6.000   0.7951   0.01314   0.00618  -0.0232   0.0105   1.0000
   6.250   0.8207   0.01364   0.00676  -0.0230   0.0097   1.0000
   6.500   0.8461   0.01412   0.00730  -0.0227   0.0085   1.0000
   6.750   0.8710   0.01468   0.00791  -0.0224   0.0076   1.0000
   7.000   0.8937   0.01576   0.00910  -0.0218   0.0068   1.0000
   7.250   0.9171   0.01666   0.01012  -0.0212   0.0064   1.0000
   7.500   0.9400   0.01765   0.01130  -0.0206   0.0061   1.0000
   7.750   0.9623   0.01880   0.01260  -0.0198   0.0057   1.0000
   8.000   0.9849   0.01979   0.01372  -0.0192   0.0053   1.0000
   8.250   1.0081   0.02052   0.01453  -0.0188   0.0049   1.0000
   8.500   1.0302   0.02141   0.01551  -0.0184   0.0045   1.0000
   8.750   1.0479   0.02344   0.01777  -0.0173   0.0043   1.0000
   9.000   1.0668   0.02518   0.01979  -0.0163   0.0041   1.0000
   9.250   1.0833   0.02740   0.02234  -0.0150   0.0040   1.0000
   9.500   1.0960   0.03031   0.02566  -0.0135   0.0038   1.0000
   9.750   1.1020   0.03431   0.03015  -0.0116   0.0037   1.0000
  10.000   1.0957   0.04015   0.03658  -0.0090   0.0036   1.0000
  10.250   1.0666   0.04891   0.04594  -0.0059   0.0035   1.0000
  10.500   1.0355   0.05533   0.05267  -0.0036   0.0035   1.0000
  10.750   1.0102   0.06141   0.05896  -0.0051   0.0036   1.0000
  11.000   0.9869   0.06879   0.06651  -0.0101   0.0036   1.0000
  11.250   0.9653   0.07804   0.07590  -0.0178   0.0037   1.0000
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