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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 500,000
Max Cl/Cd: 38.22 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m1-il-500000.txt
Download as CSV file: xf-m1-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.7270   0.09127   0.08911   0.0135   1.0000   0.0174
  -9.000  -0.7362   0.08327   0.08114   0.0067   1.0000   0.0175
  -8.750  -0.7420   0.07567   0.07349  -0.0012   1.0000   0.0175
  -8.500  -0.7527   0.06820   0.06589  -0.0052   1.0000   0.0177
  -8.250  -0.7582   0.06189   0.05944  -0.0074   1.0000   0.0179
  -8.000  -0.7551   0.05764   0.05508  -0.0082   1.0000   0.0183
  -7.750  -0.7470   0.05431   0.05164  -0.0086   1.0000   0.0187
  -7.500  -0.7365   0.05091   0.04810  -0.0090   1.0000   0.0193
  -7.250  -0.7332   0.04063   0.03724  -0.0086   1.0000   0.0160
  -7.000  -0.7243   0.03406   0.03018  -0.0077   1.0000   0.0153
  -6.750  -0.7080   0.02921   0.02481  -0.0065   1.0000   0.0160
  -6.500  -0.6889   0.02522   0.02030  -0.0054   1.0000   0.0164
  -6.250  -0.6657   0.02304   0.01773  -0.0045   1.0000   0.0170
  -6.000  -0.6454   0.01879   0.01294  -0.0036   1.0000   0.0181
  -5.750  -0.6207   0.01756   0.01160  -0.0033   1.0000   0.0193
  -5.500  -0.5947   0.01696   0.01094  -0.0031   1.0000   0.0209
  -5.250  -0.5683   0.01638   0.01025  -0.0028   1.0000   0.0231
  -5.000  -0.5416   0.01578   0.00952  -0.0025   1.0000   0.0246
  -4.750  -0.5171   0.01372   0.00728  -0.0018   1.0000   0.0268
  -4.500  -0.4915   0.01286   0.00640  -0.0015   1.0000   0.0297
  -4.250  -0.4652   0.01234   0.00585  -0.0013   1.0000   0.0326
  -4.000  -0.4389   0.01178   0.00523  -0.0009   1.0000   0.0348
  -3.750  -0.4121   0.01147   0.00487  -0.0007   1.0000   0.0361
  -3.500  -0.3874   0.01037   0.00371  -0.0002   1.0000   0.0395
  -3.250  -0.3611   0.00989   0.00321   0.0001   1.0000   0.0418
  -3.000  -0.3347   0.00953   0.00282   0.0004   1.0000   0.0445
  -2.750  -0.3082   0.00924   0.00251   0.0007   1.0000   0.0474
  -2.500  -0.2819   0.00894   0.00218   0.0011   1.0000   0.0501
  -2.250  -0.2557   0.00866   0.00192   0.0014   1.0000   0.0568
  -2.000  -0.2303   0.00812   0.00167   0.0018   1.0000   0.1226
  -1.750  -0.2076   0.00681   0.00143   0.0020   1.0000   0.4012
  -1.500  -0.1851   0.00586   0.00135   0.0026   1.0000   0.6206
  -1.250  -0.1626   0.00533   0.00136   0.0039   1.0000   0.7594
  -1.000  -0.1407   0.00506   0.00143   0.0056   1.0000   0.8502
  -0.750  -0.1188   0.00497   0.00152   0.0074   1.0000   0.9143
  -0.500  -0.0846   0.00501   0.00163   0.0064   1.0000   0.9700
  -0.250  -0.0360   0.00507   0.00168   0.0019   1.0000   0.9945
   0.000   0.0000   0.00509   0.00170   0.0000   1.0000   1.0000
   0.250   0.0358   0.00507   0.00168  -0.0019   0.9946   1.0000
   0.500   0.0847   0.00501   0.00163  -0.0065   0.9699   1.0000
   0.750   0.1189   0.00497   0.00152  -0.0074   0.9149   1.0000
   1.000   0.1408   0.00506   0.00143  -0.0056   0.8504   1.0000
   1.250   0.1627   0.00533   0.00136  -0.0039   0.7597   1.0000
   1.500   0.1851   0.00586   0.00135  -0.0027   0.6212   1.0000
   1.750   0.2076   0.00681   0.00143  -0.0020   0.4007   1.0000
   2.000   0.2304   0.00811   0.00167  -0.0018   0.1256   1.0000
   2.250   0.2557   0.00866   0.00192  -0.0014   0.0568   1.0000
   2.500   0.2819   0.00894   0.00218  -0.0011   0.0501   1.0000
   2.750   0.3082   0.00924   0.00251  -0.0007   0.0474   1.0000
   3.000   0.3347   0.00953   0.00282  -0.0004   0.0445   1.0000
   3.250   0.3612   0.00990   0.00321  -0.0001   0.0418   1.0000
   3.500   0.3874   0.01037   0.00371   0.0002   0.0395   1.0000
   3.750   0.4121   0.01146   0.00487   0.0007   0.0361   1.0000
   4.000   0.4389   0.01177   0.00522   0.0009   0.0348   1.0000
   4.250   0.4652   0.01235   0.00585   0.0013   0.0327   1.0000
   4.500   0.4915   0.01286   0.00640   0.0015   0.0297   1.0000
   4.750   0.5172   0.01370   0.00726   0.0018   0.0269   1.0000
   5.000   0.5416   0.01578   0.00952   0.0024   0.0246   1.0000
   5.250   0.5683   0.01639   0.01026   0.0028   0.0231   1.0000
   5.500   0.5948   0.01694   0.01091   0.0031   0.0209   1.0000
   5.750   0.6207   0.01755   0.01159   0.0033   0.0193   1.0000
   6.000   0.6454   0.01880   0.01294   0.0036   0.0180   1.0000
   6.250   0.6658   0.02301   0.01769   0.0045   0.0170   1.0000
   6.500   0.6889   0.02524   0.02031   0.0054   0.0164   1.0000
   6.750   0.7080   0.02926   0.02486   0.0065   0.0160   1.0000
   7.000   0.7244   0.03405   0.03016   0.0077   0.0153   1.0000
   7.250   0.7332   0.04064   0.03726   0.0085   0.0160   1.0000
   9.000   0.7372   0.08322   0.08109  -0.0068   0.0176   1.0000
  12.250   0.5778   0.13904   0.13685  -0.0223   0.0171   1.0000
  12.500   0.5782   0.14262   0.14042  -0.0236   0.0171   1.0000
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