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=5


Details Polar file
Airfoil: MH 116 9.84% (mh116-il)
Reynolds number: 500,000
Max Cl/Cd: 113.57 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh116-il-500000-n5.txt
Download as CSV file: xf-mh116-il-500000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3148   0.09774   0.09549  -0.0456   0.9970   0.0081
  -9.250  -0.2135   0.07398   0.07184  -0.0617   0.9743   0.0055
  -9.000  -0.2098   0.06850   0.06637  -0.0651   0.9701   0.0054
  -8.500  -0.2887   0.07563   0.07343  -0.0631   0.9751   0.0054
  -8.250  -0.2828   0.06968   0.06749  -0.0693   0.9660   0.0052
  -8.000  -0.2735   0.06222   0.06004  -0.0791   0.9555   0.0051
  -7.750  -0.2803   0.03331   0.03061  -0.1143   0.9286   0.0045
  -7.500  -0.2751   0.02497   0.02141  -0.1186   0.9128   0.0043
  -7.250  -0.2579   0.02111   0.01694  -0.1193   0.9022   0.0042
  -7.000  -0.2370   0.01868   0.01406  -0.1193   0.8929   0.0042
  -6.750  -0.2136   0.01691   0.01193  -0.1192   0.8848   0.0042
  -6.500  -0.1893   0.01554   0.01026  -0.1190   0.8767   0.0042
  -6.250  -0.1642   0.01447   0.00897  -0.1187   0.8692   0.0042
  -6.000  -0.1386   0.01355   0.00784  -0.1185   0.8616   0.0043
  -5.750  -0.1127   0.01277   0.00690  -0.1183   0.8543   0.0043
  -5.500  -0.0863   0.01212   0.00607  -0.1182   0.8471   0.0044
  -5.250  -0.0598   0.01155   0.00537  -0.1180   0.8399   0.0045
  -5.000  -0.0329   0.01108   0.00475  -0.1178   0.8330   0.0047
  -4.750  -0.0058   0.01065   0.00421  -0.1177   0.8260   0.0050
  -4.500   0.0215   0.01032   0.00375  -0.1176   0.8193   0.0054
  -4.250   0.0489   0.00996   0.00329  -0.1175   0.8124   0.0065
  -4.000   0.0762   0.00961   0.00289  -0.1173   0.8057   0.0131
  -3.750   0.1034   0.00931   0.00262  -0.1173   0.7990   0.0271
  -3.500   0.1305   0.00897   0.00241  -0.1173   0.7923   0.0603
  -3.250   0.1579   0.00874   0.00225  -0.1173   0.7857   0.0865
  -3.000   0.1854   0.00857   0.00210  -0.1173   0.7789   0.1091
  -2.750   0.2130   0.00842   0.00196  -0.1173   0.7725   0.1318
  -2.500   0.2405   0.00827   0.00185  -0.1173   0.7656   0.1604
  -2.250   0.2680   0.00814   0.00176  -0.1173   0.7591   0.1874
  -2.000   0.2956   0.00802   0.00168  -0.1173   0.7522   0.2154
  -1.750   0.3232   0.00792   0.00160  -0.1173   0.7457   0.2454
  -1.500   0.3508   0.00783   0.00155  -0.1173   0.7386   0.2749
  -1.250   0.3783   0.00772   0.00151  -0.1172   0.7320   0.3125
  -1.000   0.4057   0.00761   0.00148  -0.1172   0.7245   0.3543
  -0.750   0.4331   0.00751   0.00147  -0.1172   0.7173   0.3993
  -0.500   0.4604   0.00741   0.00147  -0.1171   0.7093   0.4484
  -0.250   0.4876   0.00731   0.00148  -0.1170   0.7012   0.5019
   0.000   0.5144   0.00722   0.00150  -0.1169   0.6929   0.5586
   0.250   0.5412   0.00710   0.00154  -0.1167   0.6840   0.6217
   0.500   0.5674   0.00700   0.00160  -0.1163   0.6756   0.6866
   0.750   0.5928   0.00690   0.00166  -0.1157   0.6667   0.7538
   1.000   0.6167   0.00679   0.00173  -0.1147   0.6575   0.8235
   1.250   0.6380   0.00670   0.00178  -0.1130   0.6485   0.8949
   1.500   0.6708   0.00666   0.00176  -0.1139   0.6383   0.9876
   1.750   0.6984   0.00675   0.00181  -0.1139   0.6277   1.0000
   2.000   0.7252   0.00686   0.00187  -0.1137   0.6170   1.0000
   2.250   0.7518   0.00698   0.00194  -0.1134   0.6055   1.0000
   2.500   0.7782   0.00711   0.00202  -0.1132   0.5930   1.0000
   2.750   0.8044   0.00725   0.00210  -0.1129   0.5796   1.0000
   3.000   0.8304   0.00740   0.00222  -0.1125   0.5642   1.0000
   3.250   0.8562   0.00757   0.00233  -0.1122   0.5476   1.0000
   3.500   0.8814   0.00777   0.00246  -0.1117   0.5288   1.0000
   3.750   0.9063   0.00798   0.00262  -0.1112   0.5081   1.0000
   4.000   0.9307   0.00823   0.00279  -0.1106   0.4853   1.0000
   4.250   0.9545   0.00853   0.00298  -0.1099   0.4583   1.0000
   4.500   0.9777   0.00887   0.00320  -0.1091   0.4291   1.0000
   4.750   1.0004   0.00925   0.00346  -0.1083   0.3975   1.0000
   5.000   1.0223   0.00969   0.00377  -0.1073   0.3631   1.0000
   5.250   1.0442   0.01014   0.00409  -0.1064   0.3318   1.0000
   5.500   1.0661   0.01059   0.00441  -0.1055   0.3026   1.0000
   5.750   1.0875   0.01108   0.00477  -0.1045   0.2713   1.0000
   6.000   1.1087   0.01157   0.00514  -0.1035   0.2425   1.0000
   6.250   1.1295   0.01207   0.00556  -0.1024   0.2149   1.0000
   6.500   1.1503   0.01257   0.00596  -0.1014   0.1911   1.0000
   6.750   1.1700   0.01314   0.00641  -0.1001   0.1646   1.0000
   7.000   1.1887   0.01376   0.00690  -0.0988   0.1375   1.0000
   7.250   1.2068   0.01440   0.00741  -0.0973   0.1130   1.0000
   7.500   1.2256   0.01497   0.00791  -0.0960   0.0945   1.0000
   7.750   1.2420   0.01568   0.00850  -0.0943   0.0727   1.0000
   8.000   1.2584   0.01633   0.00910  -0.0926   0.0561   1.0000
   8.250   1.2709   0.01711   0.00976  -0.0902   0.0379   1.0000
   8.500   1.2828   0.01791   0.01047  -0.0877   0.0235   1.0000
   8.750   1.2947   0.01872   0.01122  -0.0853   0.0144   1.0000
   9.000   1.3093   0.01936   0.01190  -0.0834   0.0112   1.0000
   9.250   1.3222   0.02013   0.01269  -0.0813   0.0075   1.0000
   9.500   1.3332   0.02104   0.01361  -0.0789   0.0040   1.0000
   9.750   1.3449   0.02192   0.01455  -0.0768   0.0029   1.0000
  10.000   1.3566   0.02282   0.01555  -0.0748   0.0023   1.0000
  10.250   1.3667   0.02387   0.01668  -0.0726   0.0019   1.0000
  10.500   1.3759   0.02501   0.01793  -0.0704   0.0017   1.0000
  10.750   1.3860   0.02611   0.01914  -0.0685   0.0016   1.0000
  11.000   1.3949   0.02733   0.02048  -0.0665   0.0015   1.0000
  11.250   1.4036   0.02860   0.02186  -0.0647   0.0015   1.0000
  11.500   1.4107   0.03005   0.02342  -0.0628   0.0014   1.0000
  11.750   1.4168   0.03162   0.02512  -0.0611   0.0013   1.0000
  12.000   1.4219   0.03334   0.02696  -0.0593   0.0013   1.0000
  12.250   1.4268   0.03514   0.02888  -0.0578   0.0013   1.0000
  12.500   1.4292   0.03723   0.03110  -0.0562   0.0012   1.0000
  12.750   1.4311   0.03945   0.03349  -0.0549   0.0012   1.0000
  13.000   1.4320   0.04185   0.03602  -0.0537   0.0012   1.0000
  13.250   1.4322   0.04442   0.03873  -0.0527   0.0012   1.0000
  13.500   1.4307   0.04729   0.04174  -0.0519   0.0011   1.0000
  13.750   1.4280   0.05041   0.04501  -0.0513   0.0011   1.0000
  14.000   1.4255   0.05366   0.04840  -0.0509   0.0011   1.0000
  14.250   1.4210   0.05729   0.05218  -0.0509   0.0011   1.0000
  14.500   1.4167   0.06107   0.05611  -0.0511   0.0011   1.0000
  14.750   1.4104   0.06533   0.06052  -0.0517   0.0011   1.0000
  15.000   1.4035   0.06985   0.06520  -0.0526   0.0011   1.0000
  15.250   1.3943   0.07495   0.07047  -0.0539   0.0011   1.0000
  15.500   1.3859   0.08017   0.07585  -0.0556   0.0011   1.0000
  15.750   1.3763   0.08580   0.08165  -0.0576   0.0011   1.0000
  16.000   1.3653   0.09192   0.08793  -0.0600   0.0011   1.0000
  16.250   1.3537   0.09843   0.09460  -0.0629   0.0011   1.0000
  16.500   1.3413   0.10532   0.10166  -0.0662   0.0011   1.0000
  16.750   1.3285   0.11251   0.10901  -0.0698   0.0011   1.0000
  17.000   1.3152   0.12005   0.11672  -0.0738   0.0011   1.0000
  17.250   1.3020   0.12782   0.12464  -0.0781   0.0011   1.0000
<< Back to MH 116 9.84% (mh116-il)

Polar data table (+)

Polar graphs


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