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NACA M8 AIRFOIL (m8-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M8 AIRFOIL (m8-il)
Reynolds number: 50,000
Max Cl/Cd: 26.46 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m8-il-50000-n5.txt
Download as CSV file: xf-m8-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M8 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.1265   0.11210   0.10611  -0.0270   0.8297   0.1012
  -8.750  -0.1247   0.11117   0.10516  -0.0289   0.8164   0.1022
  -8.500  -0.1228   0.10982   0.10381  -0.0304   0.8050   0.1027
  -8.250  -0.1012   0.10421   0.09815  -0.0287   0.7943   0.1065
  -8.000  -0.0931   0.10209   0.09598  -0.0291   0.7841   0.1109
  -7.750  -0.0913   0.10113   0.09504  -0.0310   0.7734   0.1142
  -7.500  -0.0933   0.10065   0.09456  -0.0330   0.7638   0.1150
  -7.250  -0.0733   0.09569   0.08956  -0.0317   0.7545   0.1177
  -7.000  -0.0650   0.09353   0.08737  -0.0318   0.7453   0.1206
  -6.750  -0.0596   0.09201   0.08584  -0.0332   0.7356   0.1245
  -6.500  -0.0535   0.09068   0.08450  -0.0361   0.7266   0.1265
  -6.250  -0.0390   0.08747   0.08125  -0.0346   0.7171   0.1317
  -6.000  -0.0303   0.08653   0.08027  -0.0385   0.7084   0.1361
  -5.750  -0.0165   0.08330   0.07704  -0.0376   0.6994   0.1389
  -5.500  -0.0032   0.08120   0.07489  -0.0389   0.6915   0.1426
  -5.250   0.0119   0.07957   0.07321  -0.0427   0.6828   0.1450
  -5.000   0.0262   0.07682   0.07040  -0.0428   0.6762   0.1468
  -4.500   0.0697   0.06888   0.06212  -0.0550   0.6614   0.1021
  -4.250   0.0863   0.06620   0.05944  -0.0547   0.6536   0.0985
  -4.000   0.1086   0.06331   0.05644  -0.0575   0.6462   0.0947
  -3.750   0.1320   0.06073   0.05369  -0.0598   0.6403   0.0945
  -3.500   0.1563   0.05853   0.05142  -0.0624   0.6314   0.0958
  -3.250   0.1821   0.05594   0.04863  -0.0646   0.6253   0.0953
  -3.000   0.2093   0.05338   0.04590  -0.0675   0.6177   0.0943
  -2.750   0.2362   0.05106   0.04341  -0.0694   0.6106   0.0948
  -2.500   0.2635   0.04892   0.04102  -0.0707   0.6057   0.0976
  -2.250   0.2917   0.04682   0.03876  -0.0730   0.5970   0.0997
  -2.000   0.3209   0.04425   0.03588  -0.0746   0.5910   0.1008
  -1.750   0.3497   0.04192   0.03324  -0.0758   0.5851   0.1035
  -1.500   0.3729   0.04165   0.03295  -0.0755   0.5765   0.1077
  -1.250   0.4018   0.03976   0.03073  -0.0760   0.5711   0.1125
  -1.000   0.4285   0.03867   0.02944  -0.0764   0.5638   0.1171
  -0.750   0.4543   0.03818   0.02881  -0.0763   0.5566   0.1244
  -0.500   0.4810   0.03760   0.02805  -0.0757   0.5518   0.1307
  -0.250   0.5060   0.03760   0.02797  -0.0757   0.5445   0.1399
   0.000   0.5321   0.03740   0.02762  -0.0755   0.5378   0.1500
   0.250   0.5610   0.03699   0.02692  -0.0750   0.5331   0.1644
   0.500   0.5865   0.03711   0.02685  -0.0750   0.5256   0.1796
   0.750   0.6121   0.03702   0.02656  -0.0748   0.5189   0.1969
   1.000   0.6390   0.03655   0.02583  -0.0743   0.5144   0.2179
   1.250   0.6620   0.03664   0.02579  -0.0740   0.5083   0.2371
   1.500   0.6840   0.03695   0.02598  -0.0736   0.5008   0.2547
   1.750   0.7112   0.03652   0.02533  -0.0729   0.4955   0.2706
   2.000   0.7333   0.03681   0.02558  -0.0727   0.4866   0.2857
   2.250   0.7597   0.03662   0.02524  -0.0722   0.4789   0.3038
   2.500   0.7888   0.03625   0.02468  -0.0719   0.4725   0.3147
   2.750   0.8109   0.03685   0.02521  -0.0718   0.4629   0.3173
   3.000   0.8421   0.03648   0.02460  -0.0714   0.4578   0.3172
   3.250   0.8614   0.03749   0.02563  -0.0713   0.4490   0.3169
   3.500   0.8882   0.03758   0.02560  -0.0710   0.4427   0.3174
   3.750   0.9176   0.03741   0.02526  -0.0706   0.4379   0.3181
   4.000   0.9325   0.03875   0.02670  -0.0703   0.4286   0.3181
   4.250   0.9621   0.03847   0.02629  -0.0698   0.4235   0.3174
   4.500   0.9776   0.03966   0.02753  -0.0693   0.4152   0.3167
   4.750   1.0010   0.03994   0.02779  -0.0687   0.4087   0.3159
   5.000   1.0332   0.03937   0.02706  -0.0681   0.4046   0.3151
   5.250   1.0372   0.04153   0.02939  -0.0672   0.3945   0.3146
   5.500   1.0660   0.04117   0.02893  -0.0664   0.3894   0.3139
   5.750   1.0747   0.04272   0.03059  -0.0654   0.3812   0.3135
   6.000   1.0930   0.04329   0.03115  -0.0644   0.3745   0.3136
   6.250   1.1258   0.04255   0.03028  -0.0636   0.3704   0.3143
   6.500   1.1130   0.04598   0.03393  -0.0621   0.3604   0.3143
   6.750   1.1363   0.04609   0.03402  -0.0611   0.3553   0.3150
   7.000   1.1724   0.04514   0.03295  -0.0603   0.3518   0.3157
   7.250   1.1213   0.05172   0.03981  -0.0582   0.3400   0.3151
   7.500   1.1513   0.05114   0.03918  -0.0571   0.3361   0.3156
   7.750   1.1952   0.04954   0.03748  -0.0563   0.3333   0.3160
   8.000   1.1042   0.06159   0.04983  -0.0568   0.3185   0.3153
   8.250   1.1369   0.06039   0.04860  -0.0553   0.3162   0.3155
   8.500   1.1751   0.05875   0.04690  -0.0537   0.3145   0.3157
   9.000   1.0941   0.07587   0.06428  -0.0572   0.2965   0.3152
   9.500   1.0328   0.09308   0.08165  -0.0622   0.2800   0.3149
  10.000   0.9817   0.10962   0.09833  -0.0676   0.2671   0.3147
  10.250   0.9928   0.11181   0.10055  -0.0675   0.2652   0.3148
  10.500   1.0078   0.11337   0.10214  -0.0671   0.2637   0.3149
  10.750   0.9689   0.12424   0.11310  -0.0716   0.2581   0.3147
  11.000   0.9673   0.12855   0.11748  -0.0728   0.2546   0.3148
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