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

MH 116 9.84% (mh116-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: MH 116 9.84% (mh116-il)
Reynolds number: 500,000
Max Cl/Cd: 125.59 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh116-il-500000.txt
Download as CSV file: xf-mh116-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 116  9.84%                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3310   0.10266   0.10049  -0.0388   1.0000   0.0158
  -9.250  -0.3346   0.10053   0.09840  -0.0376   1.0000   0.0160
  -9.000  -0.3404   0.09849   0.09640  -0.0362   0.9999   0.0163
  -8.750  -0.3255   0.09417   0.09208  -0.0409   0.9979   0.0168
  -8.500  -0.3104   0.08963   0.08755  -0.0460   0.9954   0.0174
  -8.250  -0.2961   0.08489   0.08281  -0.0515   0.9921   0.0183
  -8.000  -0.2817   0.07982   0.07775  -0.0576   0.9875   0.0191
  -7.750  -0.2618   0.07376   0.07170  -0.0668   0.9832   0.0204
  -7.500  -0.2454   0.06845   0.06640  -0.0745   0.9745   0.0208
  -7.250  -0.2171   0.06010   0.05803  -0.0891   0.9700   0.0209
  -7.000  -0.1925   0.04855   0.04625  -0.1059   0.9606   0.0210
  -6.750  -0.1862   0.03679   0.03411  -0.1175   0.9502   0.0223
  -6.500  -0.1612   0.03377   0.03095  -0.1204   0.9445   0.0231
  -6.000  -0.1258   0.01897   0.01451  -0.1230   0.9251   0.0115
  -5.750  -0.1008   0.01571   0.01055  -0.1224   0.9175   0.0097
  -5.500  -0.0743   0.01467   0.00930  -0.1221   0.9103   0.0094
  -5.250  -0.0476   0.01364   0.00808  -0.1219   0.9034   0.0093
  -5.000  -0.0219   0.01258   0.00686  -0.1215   0.8959   0.0093
  -4.750   0.0042   0.01142   0.00552  -0.1212   0.8891   0.0095
  -4.500   0.0304   0.01064   0.00461  -0.1209   0.8816   0.0101
  -4.250   0.0580   0.01012   0.00395  -0.1208   0.8750   0.0117
  -4.000   0.0846   0.00932   0.00320  -0.1204   0.8676   0.0403
  -3.750   0.1121   0.00903   0.00295  -0.1204   0.8611   0.0681
  -3.500   0.1389   0.00875   0.00281  -0.1204   0.8538   0.1055
  -3.250   0.1665   0.00852   0.00264  -0.1203   0.8473   0.1379
  -3.000   0.1937   0.00832   0.00249  -0.1203   0.8402   0.1712
  -2.750   0.2213   0.00814   0.00236  -0.1203   0.8336   0.2068
  -2.500   0.2485   0.00795   0.00226  -0.1202   0.8265   0.2498
  -2.250   0.2761   0.00778   0.00217  -0.1202   0.8200   0.2925
  -2.000   0.3034   0.00761   0.00210  -0.1201   0.8129   0.3404
  -1.750   0.3308   0.00745   0.00203  -0.1201   0.8062   0.3931
  -1.500   0.3580   0.00728   0.00200  -0.1200   0.7992   0.4508
  -1.250   0.3851   0.00710   0.00198  -0.1199   0.7924   0.5209
  -1.000   0.4117   0.00691   0.00198  -0.1196   0.7851   0.5974
  -0.750   0.4379   0.00674   0.00199  -0.1192   0.7781   0.6765
  -0.500   0.4628   0.00656   0.00203  -0.1185   0.7704   0.7568
  -0.250   0.4861   0.00642   0.00205  -0.1171   0.7633   0.8344
   0.000   0.5067   0.00628   0.00206  -0.1151   0.7550   0.9103
   0.250   0.5459   0.00620   0.00195  -0.1173   0.7480   0.9983
   0.500   0.5731   0.00625   0.00195  -0.1172   0.7395   1.0000
   0.750   0.6005   0.00633   0.00196  -0.1171   0.7318   1.0000
   1.000   0.6276   0.00639   0.00197  -0.1169   0.7234   1.0000
   1.250   0.6548   0.00647   0.00200  -0.1168   0.7148   1.0000
   1.500   0.6820   0.00656   0.00201  -0.1166   0.7067   1.0000
   1.750   0.7089   0.00663   0.00207  -0.1164   0.6974   1.0000
   2.000   0.7359   0.00671   0.00212  -0.1163   0.6884   1.0000
   2.500   0.7895   0.00690   0.00223  -0.1158   0.6689   1.0000
   2.750   0.8162   0.00699   0.00232  -0.1156   0.6585   1.0000
   3.000   0.8427   0.00710   0.00240  -0.1153   0.6475   1.0000
   3.250   0.8690   0.00722   0.00248  -0.1150   0.6360   1.0000
   3.500   0.8950   0.00735   0.00258  -0.1146   0.6234   1.0000
   4.250   0.9720   0.00777   0.00296  -0.1133   0.5802   1.0000
   4.500   0.9972   0.00794   0.00310  -0.1128   0.5627   1.0000
   4.750   1.0217   0.00815   0.00328  -0.1122   0.5431   1.0000
   5.000   1.0457   0.00838   0.00347  -0.1115   0.5200   1.0000
   5.250   1.0689   0.00866   0.00368  -0.1106   0.4943   1.0000
   5.500   1.0913   0.00900   0.00392  -0.1097   0.4639   1.0000
   6.000   1.1325   0.00994   0.00458  -0.1072   0.3895   1.0000
   6.250   1.1524   0.01047   0.00497  -0.1059   0.3520   1.0000
   6.500   1.1717   0.01105   0.00539  -0.1046   0.3161   1.0000
   6.750   1.1897   0.01172   0.00587  -0.1030   0.2742   1.0000
   7.000   1.2071   0.01243   0.00638  -0.1014   0.2337   1.0000
   7.250   1.2244   0.01314   0.00690  -0.0998   0.1967   1.0000
   7.500   1.2411   0.01386   0.00746  -0.0981   0.1639   1.0000
   7.750   1.2570   0.01463   0.00808  -0.0963   0.1316   1.0000
   8.000   1.2727   0.01537   0.00868  -0.0945   0.1060   1.0000
   8.250   1.2865   0.01617   0.00933  -0.0924   0.0804   1.0000
   8.500   1.2988   0.01694   0.00998  -0.0900   0.0608   1.0000
   8.750   1.3117   0.01768   0.01064  -0.0877   0.0460   1.0000
   9.000   1.3236   0.01847   0.01139  -0.0853   0.0336   1.0000
   9.250   1.3324   0.01949   0.01231  -0.0825   0.0197   1.0000
   9.500   1.3308   0.02128   0.01403  -0.0782   0.0065   1.0000
   9.750   1.3372   0.02259   0.01546  -0.0752   0.0052   1.0000
  10.000   1.3466   0.02368   0.01666  -0.0729   0.0046   1.0000
  10.250   1.3549   0.02488   0.01800  -0.0705   0.0043   1.0000
  10.500   1.3618   0.02621   0.01943  -0.0681   0.0040   1.0000
  10.750   1.3675   0.02769   0.02102  -0.0658   0.0038   1.0000
  11.000   1.3700   0.02948   0.02292  -0.0634   0.0036   1.0000
  11.250   1.3729   0.03130   0.02487  -0.0612   0.0036   1.0000
  11.500   1.3718   0.03358   0.02728  -0.0589   0.0035   1.0000
  11.750   1.3712   0.03593   0.02976  -0.0569   0.0034   1.0000
  12.000   1.3686   0.03863   0.03260  -0.0549   0.0034   1.0000
  12.250   1.3697   0.04104   0.03515  -0.0535   0.0034   1.0000
  12.500   1.3652   0.04427   0.03853  -0.0519   0.0033   1.0000
  12.750   1.3642   0.04719   0.04161  -0.0508   0.0033   1.0000
  13.000   1.3633   0.05022   0.04479  -0.0499   0.0033   1.0000
  13.250   1.3621   0.05337   0.04810  -0.0492   0.0033   1.0000
  13.500   1.3603   0.05672   0.05160  -0.0487   0.0033   1.0000
  13.750   1.3570   0.06038   0.05546  -0.0484   0.0034   1.0000
  14.000   1.3537   0.06412   0.05937  -0.0485   0.0034   1.0000
  14.250   1.3487   0.06824   0.06367  -0.0488   0.0034   1.0000
  14.500   1.3434   0.07254   0.06815  -0.0495   0.0034   1.0000
  14.750   1.3360   0.07737   0.07317  -0.0507   0.0034   1.0000
  15.000   1.3275   0.08256   0.07855  -0.0522   0.0035   1.0000
  15.250   1.3161   0.08854   0.08475  -0.0542   0.0035   1.0000
  15.500   1.3035   0.09503   0.09146  -0.0569   0.0036   1.0000
  15.750   1.2870   0.10263   0.09931  -0.0603   0.0037   1.0000
  16.000   1.2672   0.11133   0.10826  -0.0647   0.0038   1.0000
  16.250   1.2459   0.12090   0.11807  -0.0700   0.0040   1.0000
<< Back to MH 116 9.84% (mh116-il)

Polar data table (+)

Polar graphs


<< Back to MH 116 9.84% (mh116-il)