Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA M17 AIRFOIL (m17-il)
Reynolds number: 500,000
Max Cl/Cd: 98.27 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m17-il-500000-n5.txt
Download as CSV file: xf-m17-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M17 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4895   0.09685   0.09435   0.0099   0.7980   0.0090
  -8.250  -0.4880   0.09281   0.09027   0.0075   0.7841   0.0092
  -8.000  -0.4878   0.08856   0.08598   0.0042   0.7713   0.0093
  -7.750  -0.4853   0.08380   0.08117   0.0001   0.7598   0.0096
  -7.500  -0.4801   0.07789   0.07519  -0.0051   0.7494   0.0101
  -7.250  -0.4736   0.07076   0.06797  -0.0109   0.7405   0.0108
  -7.000  -0.4569   0.06837   0.06549  -0.0129   0.7296   0.0111
  -6.750  -0.4396   0.06569   0.06273  -0.0150   0.7191   0.0116
  -6.500  -0.4222   0.06199   0.05891  -0.0176   0.7098   0.0122
  -6.250  -0.4047   0.05662   0.05338  -0.0208   0.7015   0.0131
  -6.000  -0.3878   0.04746   0.04392  -0.0248   0.6950   0.0148
  -5.500  -0.3420   0.04439   0.04063  -0.0261   0.6771   0.0160
  -5.250  -0.3193   0.04074   0.03675  -0.0269   0.6698   0.0171
  -5.000  -0.2998   0.02937   0.02469  -0.0272   0.6652   0.0204
  -4.750  -0.2751   0.02851   0.02368  -0.0273   0.6574   0.0209
  -4.500  -0.2502   0.02728   0.02230  -0.0273   0.6501   0.0215
  -4.250  -0.2250   0.02603   0.02087  -0.0272   0.6429   0.0223
  -4.000  -0.1998   0.02394   0.01854  -0.0269   0.6364   0.0236
  -3.750  -0.1751   0.02074   0.01489  -0.0263   0.6304   0.0246
  -3.500  -0.1492   0.01875   0.01254  -0.0259   0.6245   0.0256
  -3.250  -0.1222   0.01770   0.01125  -0.0256   0.6179   0.0265
  -2.750  -0.0677   0.01599   0.00912  -0.0253   0.6062   0.0273
  -2.500  -0.0401   0.01542   0.00838  -0.0253   0.6002   0.0276
  -2.250  -0.0124   0.01494   0.00776  -0.0252   0.5946   0.0277
  -2.000   0.0150   0.01388   0.00651  -0.0251   0.5887   0.0280
  -1.750   0.0425   0.01306   0.00553  -0.0250   0.5834   0.0283
  -1.500   0.0700   0.01226   0.00463  -0.0249   0.5781   0.0291
  -1.250   0.0976   0.01178   0.00408  -0.0249   0.5723   0.0294
  -1.000   0.1252   0.01141   0.00362  -0.0248   0.5670   0.0295
  -0.750   0.1528   0.01105   0.00323  -0.0247   0.5616   0.0297
  -0.500   0.1804   0.01074   0.00288  -0.0246   0.5564   0.0300
  -0.250   0.2079   0.01048   0.00259  -0.0245   0.5513   0.0306
   0.000   0.2356   0.01026   0.00235  -0.0245   0.5457   0.0313
   0.250   0.2634   0.01011   0.00217  -0.0245   0.5403   0.0323
   0.500   0.2913   0.00998   0.00203  -0.0245   0.5355   0.0334
   0.750   0.3192   0.00988   0.00191  -0.0246   0.5302   0.0345
   1.000   0.3472   0.00981   0.00182  -0.0246   0.5249   0.0355
   1.250   0.3753   0.00975   0.00175  -0.0247   0.5197   0.0365
   1.500   0.4034   0.00971   0.00170  -0.0247   0.5145   0.0374
   1.750   0.4315   0.00971   0.00167  -0.0248   0.5095   0.0387
   2.000   0.4596   0.00967   0.00166  -0.0249   0.5043   0.0431
   2.250   0.4875   0.00960   0.00169  -0.0250   0.4989   0.0721
   2.500   0.5151   0.00952   0.00175  -0.0251   0.4941   0.1313
   2.750   0.5398   0.00875   0.00185  -0.0250   0.4889   0.5046
   3.250   0.6351   0.00786   0.00219  -0.0336   0.4725   0.9977
   3.500   0.6703   0.00798   0.00227  -0.0354   0.4630   1.0000
   3.750   0.6964   0.00807   0.00233  -0.0352   0.4529   1.0000
   4.000   0.7226   0.00815   0.00241  -0.0349   0.4455   1.0000
   4.250   0.7487   0.00825   0.00252  -0.0346   0.4394   1.0000
   4.500   0.7748   0.00836   0.00262  -0.0344   0.4329   1.0000
   4.750   0.8008   0.00846   0.00274  -0.0341   0.4256   1.0000
   5.000   0.8268   0.00859   0.00288  -0.0339   0.4168   1.0000
   5.250   0.8524   0.00875   0.00302  -0.0336   0.4018   1.0000
   5.500   0.8779   0.00894   0.00318  -0.0333   0.3843   1.0000
   5.750   0.9031   0.00919   0.00337  -0.0330   0.3624   1.0000
   6.000   0.9266   0.00973   0.00368  -0.0326   0.3069   1.0000
   6.250   0.9474   0.01075   0.00427  -0.0321   0.2236   1.0000
   6.750   0.9841   0.01335   0.00595  -0.0306   0.0575   1.0000
   7.000   1.0039   0.01426   0.00670  -0.0298   0.0241   1.0000
   7.250   1.0263   0.01477   0.00722  -0.0291   0.0181   1.0000
   7.500   1.0487   0.01526   0.00774  -0.0285   0.0156   1.0000
   7.750   1.0700   0.01588   0.00841  -0.0277   0.0130   1.0000
   8.000   1.0915   0.01644   0.00904  -0.0270   0.0119   1.0000
   8.250   1.1128   0.01701   0.00970  -0.0263   0.0111   1.0000
   8.500   1.1332   0.01765   0.01041  -0.0255   0.0103   1.0000
   8.750   1.1529   0.01834   0.01116  -0.0246   0.0096   1.0000
   9.000   1.1700   0.01925   0.01212  -0.0235   0.0088   1.0000
   9.250   1.1864   0.02018   0.01313  -0.0224   0.0082   1.0000
   9.500   1.2033   0.02099   0.01403  -0.0213   0.0078   1.0000
   9.750   1.2180   0.02193   0.01508  -0.0200   0.0074   1.0000
  10.000   1.2308   0.02296   0.01619  -0.0185   0.0071   1.0000
  10.250   1.2405   0.02403   0.01735  -0.0167   0.0068   1.0000
  10.500   1.2468   0.02527   0.01867  -0.0148   0.0066   1.0000
  10.750   1.2533   0.02676   0.02023  -0.0135   0.0064   1.0000
  11.000   1.2595   0.02846   0.02202  -0.0126   0.0062   1.0000
  11.250   1.2639   0.03047   0.02411  -0.0119   0.0060   1.0000
  11.500   1.2646   0.03294   0.02668  -0.0111   0.0058   1.0000
  11.750   1.2672   0.03530   0.02914  -0.0105   0.0057   1.0000
  12.000   1.2733   0.03733   0.03132  -0.0101   0.0056   1.0000
  12.250   1.2785   0.03950   0.03359  -0.0098   0.0054   1.0000
  12.500   1.2826   0.04178   0.03599  -0.0095   0.0052   1.0000
  12.750   1.2862   0.04416   0.03849  -0.0092   0.0050   1.0000
  13.000   1.2886   0.04668   0.04113  -0.0090   0.0049   1.0000
  13.250   1.2898   0.04936   0.04392  -0.0088   0.0048   1.0000
  13.500   1.2905   0.05215   0.04683  -0.0088   0.0046   1.0000
  13.750   1.2904   0.05505   0.04984  -0.0088   0.0045   1.0000
  14.000   1.2895   0.05810   0.05302  -0.0089   0.0045   1.0000
  14.250   1.2879   0.06129   0.05632  -0.0092   0.0044   1.0000
  14.500   1.2857   0.06467   0.05982  -0.0097   0.0043   1.0000
  14.750   1.2830   0.06826   0.06353  -0.0104   0.0043   1.0000
  15.000   1.2796   0.07206   0.06745  -0.0113   0.0042   1.0000
  15.250   1.2752   0.07607   0.07158  -0.0123   0.0041   1.0000
  15.500   1.2698   0.08037   0.07600  -0.0136   0.0041   1.0000
  15.750   1.2636   0.08493   0.08072  -0.0150   0.0041   1.0000
  16.000   1.2560   0.08978   0.08570  -0.0167   0.0040   1.0000
  16.250   1.2471   0.09500   0.09105  -0.0186   0.0040   1.0000
  16.500   1.2368   0.10062   0.09682  -0.0208   0.0039   1.0000
  16.750   1.2251   0.10665   0.10299  -0.0232   0.0039   1.0000
  17.000   1.2125   0.11302   0.10951  -0.0260   0.0039   1.0000
  17.250   1.1989   0.11982   0.11646  -0.0292   0.0039   1.0000
  17.500   1.1848   0.12699   0.12378  -0.0327   0.0039   1.0000
  17.750   1.1702   0.13458   0.13153  -0.0366   0.0039   1.0000
  18.000   1.1552   0.14254   0.13963  -0.0409   0.0039   1.0000
  18.250   1.1398   0.15103   0.14827  -0.0456   0.0039   1.0000
  18.500   1.1232   0.16033   0.15773  -0.0509   0.0039   1.0000
  18.750   1.1014   0.17187   0.16942  -0.0574   0.0039   1.0000
<< Back to NACA M17 AIRFOIL (m17-il)

Polar data table (+)

Polar graphs


<< Back to NACA M17 AIRFOIL (m17-il)