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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 50,000
Max Cl/Cd: 18.4 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m1-il-50000.txt
Download as CSV file: xf-m1-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-M1 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5470   0.09904   0.09251   0.0242   1.0000   0.3213
  -8.250  -0.5506   0.09553   0.08905   0.0246   1.0000   0.3367
  -8.000  -0.5589   0.09224   0.08582   0.0250   1.0000   0.3518
  -7.750  -0.5374   0.08751   0.08107   0.0270   1.0000   0.3776
  -7.500  -0.5255   0.08332   0.07689   0.0285   1.0000   0.4024
  -7.000  -0.6488   0.08654   0.07991   0.0432   1.0000   0.4258
  -6.750  -0.6477   0.08405   0.07747   0.0461   1.0000   0.4632
  -6.250  -0.6138   0.07614   0.06953   0.0504   1.0000   0.5246
  -6.000  -0.6477   0.05051   0.04223  -0.0088   1.0000   0.1545
  -5.750  -0.6256   0.04508   0.03618  -0.0095   1.0000   0.1393
  -5.500  -0.6035   0.04108   0.03168  -0.0094   1.0000   0.1379
  -5.250  -0.5793   0.03730   0.02731  -0.0091   1.0000   0.1354
  -5.000  -0.5527   0.03382   0.02306  -0.0086   1.0000   0.1319
  -4.750  -0.5259   0.03103   0.01985  -0.0080   1.0000   0.1335
  -4.500  -0.4988   0.02897   0.01728  -0.0073   1.0000   0.1418
  -4.250  -0.4710   0.02661   0.01477  -0.0067   1.0000   0.1467
  -4.000  -0.4423   0.02474   0.01269  -0.0059   1.0000   0.1540
  -3.750  -0.4143   0.02308   0.01095  -0.0051   1.0000   0.1686
  -3.500  -0.3865   0.02147   0.00947  -0.0045   1.0000   0.1919
  -3.250  -0.3613   0.01963   0.00794  -0.0035   1.0000   0.2352
  -3.000  -0.2520   0.01663   0.00742  -0.0102   1.0000   1.0000
  -2.750  -0.2320   0.01620   0.00662  -0.0097   1.0000   1.0000
  -2.500  -0.2115   0.01584   0.00600  -0.0091   1.0000   1.0000
  -2.250  -0.1907   0.01556   0.00547  -0.0085   1.0000   1.0000
  -2.000  -0.1698   0.01532   0.00502  -0.0077   1.0000   1.0000
  -1.750  -0.1488   0.01513   0.00465  -0.0069   1.0000   1.0000
  -1.500  -0.1276   0.01497   0.00431  -0.0060   1.0000   1.0000
  -1.250  -0.1063   0.01485   0.00406  -0.0051   1.0000   1.0000
  -1.000  -0.0850   0.01475   0.00386  -0.0041   1.0000   1.0000
  -0.750  -0.0638   0.01467   0.00371  -0.0031   1.0000   1.0000
  -0.500  -0.0426   0.01462   0.00359  -0.0021   1.0000   1.0000
  -0.250  -0.0213   0.01459   0.00352  -0.0010   1.0000   1.0000
   0.000   0.0000   0.01458   0.00350   0.0000   1.0000   1.0000
   0.250   0.0213   0.01459   0.00352   0.0010   1.0000   1.0000
   0.500   0.0425   0.01462   0.00359   0.0021   1.0000   1.0000
   0.750   0.0638   0.01467   0.00371   0.0031   1.0000   1.0000
   1.000   0.0850   0.01475   0.00386   0.0041   1.0000   1.0000
   1.250   0.1063   0.01485   0.00406   0.0051   1.0000   1.0000
   1.500   0.1276   0.01497   0.00431   0.0060   1.0000   1.0000
   1.750   0.1488   0.01513   0.00465   0.0069   1.0000   1.0000
   2.000   0.1698   0.01532   0.00502   0.0077   1.0000   1.0000
   2.250   0.1908   0.01555   0.00547   0.0085   1.0000   1.0000
   2.500   0.2115   0.01584   0.00599   0.0091   1.0000   1.0000
   2.750   0.2320   0.01620   0.00662   0.0097   1.0000   1.0000
   3.000   0.2521   0.01663   0.00742   0.0102   1.0000   1.0000
   3.250   0.3613   0.01964   0.00794   0.0035   0.2352   1.0000
   3.500   0.3865   0.02147   0.00947   0.0045   0.1919   1.0000
   3.750   0.4143   0.02308   0.01095   0.0051   0.1686   1.0000
   4.000   0.4423   0.02474   0.01269   0.0059   0.1539   1.0000
   4.250   0.4710   0.02661   0.01477   0.0067   0.1467   1.0000
   4.500   0.4988   0.02898   0.01728   0.0073   0.1418   1.0000
   4.750   0.5259   0.03103   0.01985   0.0080   0.1336   1.0000
   5.000   0.5527   0.03382   0.02306   0.0086   0.1319   1.0000
   5.250   0.5793   0.03730   0.02731   0.0091   0.1354   1.0000
   5.500   0.6035   0.04108   0.03168   0.0094   0.1379   1.0000
   5.750   0.6256   0.04509   0.03619   0.0095   0.1394   1.0000
   6.000   0.6477   0.05051   0.04224   0.0088   0.1546   1.0000
   6.750   0.6477   0.08404   0.07745  -0.0462   0.4631   1.0000
   7.000   0.6512   0.08665   0.08002  -0.0431   0.4251   1.0000
   7.500   0.6619   0.09338   0.08668  -0.0396   0.3714   1.0000
   7.750   0.6788   0.09793   0.09129  -0.0378   0.3508   1.0000
   8.000   0.6707   0.10053   0.09381  -0.0378   0.3309   1.0000
   8.250   0.6829   0.10461   0.09789  -0.0364   0.3121   1.0000
   8.500   0.6919   0.10897   0.10223  -0.0354   0.2950   1.0000
   8.750   0.6874   0.11200   0.10519  -0.0358   0.2795   1.0000
   9.000   0.6867   0.11531   0.10845  -0.0360   0.2640   1.0000
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