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M6 (65%) (m665-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 500,000
Max Cl/Cd: 74.71 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m665-il-500000.txt
Download as CSV file: xf-m665-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: M6 (65%)                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5469   0.11337   0.11096   0.0015   1.0000   0.0176
 -10.250  -0.5449   0.10946   0.10705  -0.0002   1.0000   0.0181
  -7.750  -0.5898   0.05986   0.05741  -0.0248   1.0000   0.0204
  -7.500  -0.5868   0.05677   0.05427  -0.0240   1.0000   0.0207
  -7.250  -0.5809   0.05424   0.05166  -0.0231   1.0000   0.0211
  -7.000  -0.5734   0.05163   0.04898  -0.0221   1.0000   0.0216
  -6.750  -0.5663   0.04841   0.04565  -0.0208   1.0000   0.0221
  -6.500  -0.5572   0.04552   0.04263  -0.0194   1.0000   0.0233
  -6.250  -0.5485   0.04185   0.03877  -0.0174   1.0000   0.0245
  -6.000  -0.5298   0.03994   0.03645  -0.0147   1.0000   0.0267
  -5.750  -0.5515   0.02776   0.02354  -0.0075   1.0000   0.0190
  -5.500  -0.5472   0.02352   0.01885  -0.0032   1.0000   0.0179
  -5.250  -0.5374   0.02029   0.01516   0.0005   1.0000   0.0179
  -5.000  -0.5229   0.01825   0.01278   0.0034   1.0000   0.0184
  -4.500  -0.4672   0.01479   0.00875   0.0034   0.9960   0.0209
  -4.250  -0.4305   0.01444   0.00843   0.0012   0.9918   0.0232
  -4.000  -0.3938   0.01372   0.00761  -0.0006   0.9871   0.0254
  -3.750  -0.3555   0.01308   0.00685  -0.0028   0.9831   0.0271
  -3.500  -0.3242   0.01171   0.00544  -0.0037   0.9756   0.0305
  -3.250  -0.2865   0.01110   0.00481  -0.0058   0.9705   0.0331
  -3.000  -0.2522   0.01066   0.00433  -0.0071   0.9607   0.0356
  -2.750  -0.2178   0.00986   0.00347  -0.0085   0.9500   0.0385
  -2.500  -0.1824   0.00935   0.00294  -0.0100   0.9363   0.0431
  -2.250  -0.1486   0.00903   0.00256  -0.0112   0.9183   0.0481
  -2.000  -0.1190   0.00871   0.00221  -0.0114   0.8976   0.0597
  -1.750  -0.0950   0.00824   0.00195  -0.0104   0.8742   0.1219
  -1.500  -0.0734   0.00772   0.00178  -0.0092   0.8523   0.2431
  -1.250  -0.0595   0.00672   0.00160  -0.0066   0.8308   0.4950
  -1.000  -0.0516   0.00546   0.00159  -0.0019   0.8121   0.8483
  -0.750   0.0297   0.00570   0.00189  -0.0130   0.7987   0.9352
  -0.500   0.0694   0.00606   0.00213  -0.0150   0.7819   0.9534
  -0.250   0.1139   0.00640   0.00234  -0.0182   0.7655   0.9643
   0.000   0.1664   0.00675   0.00258  -0.0232   0.7490   0.9756
   0.250   0.2269   0.00698   0.00270  -0.0301   0.7328   0.9865
   0.500   0.2790   0.00701   0.00265  -0.0353   0.7162   0.9942
   0.750   0.3232   0.00689   0.00244  -0.0389   0.6982   0.9987
   1.000   0.3518   0.00689   0.00235  -0.0391   0.6796   1.0000
   1.250   0.3743   0.00691   0.00230  -0.0381   0.6623   1.0000
   1.500   0.3971   0.00694   0.00228  -0.0371   0.6469   1.0000
   1.750   0.4200   0.00697   0.00227  -0.0361   0.6319   1.0000
   2.000   0.4429   0.00701   0.00227  -0.0351   0.6156   1.0000
   2.250   0.4657   0.00706   0.00227  -0.0341   0.5970   1.0000
   2.500   0.4884   0.00713   0.00229  -0.0331   0.5782   1.0000
   2.750   0.5112   0.00720   0.00231  -0.0321   0.5554   1.0000
   3.000   0.5334   0.00731   0.00234  -0.0310   0.5238   1.0000
   3.250   0.5547   0.00753   0.00236  -0.0297   0.4760   1.0000
   3.500   0.5766   0.00775   0.00243  -0.0286   0.4325   1.0000
   3.750   0.5984   0.00801   0.00254  -0.0275   0.3870   1.0000
   4.000   0.6193   0.00840   0.00270  -0.0263   0.3241   1.0000
   4.250   0.6385   0.00905   0.00295  -0.0249   0.2329   1.0000
   4.500   0.6557   0.01001   0.00335  -0.0234   0.1172   1.0000
   4.750   0.6734   0.01097   0.00391  -0.0217   0.0429   1.0000
   5.000   0.6941   0.01152   0.00444  -0.0203   0.0317   1.0000
   5.250   0.7155   0.01193   0.00491  -0.0191   0.0285   1.0000
   5.500   0.7347   0.01264   0.00567  -0.0175   0.0256   1.0000
   5.750   0.7537   0.01336   0.00647  -0.0158   0.0246   1.0000
   6.000   0.7736   0.01392   0.00710  -0.0143   0.0240   1.0000
   6.250   0.7928   0.01459   0.00784  -0.0127   0.0233   1.0000
   6.500   0.8119   0.01527   0.00859  -0.0111   0.0226   1.0000
   6.750   0.8309   0.01597   0.00935  -0.0095   0.0219   1.0000
   7.000   0.8493   0.01682   0.01026  -0.0078   0.0215   1.0000
   7.250   0.8678   0.01772   0.01124  -0.0062   0.0211   1.0000
   7.500   0.8863   0.01872   0.01233  -0.0045   0.0207   1.0000
   7.750   0.9048   0.01983   0.01354  -0.0029   0.0204   1.0000
   8.000   0.9232   0.02114   0.01496  -0.0013   0.0202   1.0000
   8.250   0.9413   0.02246   0.01639   0.0002   0.0198   1.0000
   8.500   0.9583   0.02393   0.01793   0.0017   0.0190   1.0000
   8.750   0.9727   0.02659   0.02079   0.0036   0.0186   1.0000
   9.000   0.9857   0.02920   0.02367   0.0056   0.0185   1.0000
   9.250   0.9943   0.03240   0.02718   0.0081   0.0184   1.0000
   9.500   0.9957   0.03656   0.03172   0.0111   0.0182   1.0000
   9.750   1.0122   0.03516   0.03051   0.0133   0.0170   1.0000
  10.000   1.0129   0.03821   0.03389   0.0164   0.0168   1.0000
  10.250   1.0067   0.04184   0.03783   0.0199   0.0170   1.0000
  10.500   0.9970   0.04525   0.04153   0.0234   0.0170   1.0000
  10.750   0.9775   0.04904   0.04554   0.0275   0.0173   1.0000
  11.000   0.9546   0.05243   0.04908   0.0312   0.0176   1.0000
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