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MH 122 9.32% (mh122-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 122 9.32% (mh122-il)
Reynolds number: 200,000
Max Cl/Cd: 90.1 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh122-il-200000.txt
Download as CSV file: xf-mh122-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 122  9.32%                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4426   0.10210   0.09890  -0.0275   1.0000   0.0312
  -8.000  -0.4517   0.09986   0.09671  -0.0267   1.0000   0.0316
  -7.750  -0.4616   0.09772   0.09461  -0.0258   1.0000   0.0320
  -7.500  -0.4581   0.09355   0.09047  -0.0319   0.9964   0.0329
  -7.250  -0.3612   0.07306   0.07016  -0.0477   0.9833   0.0353
  -7.000  -0.3604   0.06976   0.06688  -0.0477   0.9798   0.0360
  -6.750  -0.3637   0.06616   0.06329  -0.0489   0.9750   0.0366
  -6.500  -0.3574   0.06076   0.05790  -0.0544   0.9716   0.0376
  -6.250  -0.3564   0.05479   0.05188  -0.0614   0.9670   0.0381
  -6.000  -0.3523   0.04956   0.04653  -0.0661   0.9624   0.0392
  -5.750  -0.3374   0.04359   0.04037  -0.0724   0.9593   0.0413
  -5.500  -0.3157   0.04054   0.03661  -0.0787   0.9550   0.0465
  -5.000  -0.2988   0.04196   0.03744  -0.0868   0.9583   0.0497
  -4.750  -0.2759   0.03921   0.03456  -0.0880   0.9555   0.0530
  -4.500  -0.2498   0.03629   0.03106  -0.0897   0.9522   0.0622
  -4.250  -0.2219   0.03362   0.02830  -0.0911   0.9500   0.0663
  -4.000  -0.1737   0.02705   0.02052  -0.0909   0.9498   0.0210
  -3.750  -0.1392   0.02439   0.01743  -0.0922   0.9486   0.0190
  -3.500  -0.1048   0.02249   0.01520  -0.0932   0.9473   0.0179
  -3.250  -0.0703   0.02102   0.01350  -0.0942   0.9462   0.0180
  -3.000  -0.0342   0.01990   0.01227  -0.0958   0.9450   0.0196
  -2.750  -0.0118   0.01928   0.01154  -0.0950   0.9412   0.0216
  -2.500   0.0240   0.01656   0.01029  -0.0982   0.9402   0.3746
  -2.250   0.0506   0.01645   0.01064  -0.0981   0.9371   0.5504
  -2.000   0.0815   0.01657   0.01074  -0.0986   0.9346   0.5929
  -1.750   0.1150   0.01670   0.01084  -0.0997   0.9327   0.6287
  -1.500   0.1369   0.01686   0.01100  -0.0986   0.9286   0.6596
  -1.250   0.1601   0.01698   0.01117  -0.0977   0.9244   0.6896
  -1.000   0.1915   0.01704   0.01120  -0.0985   0.9217   0.7104
  -0.750   0.2263   0.01707   0.01121  -0.1000   0.9197   0.7279
  -0.500   0.2482   0.01718   0.01133  -0.0990   0.9149   0.7454
  -0.250   0.2744   0.01724   0.01141  -0.0989   0.9105   0.7662
   0.000   0.3068   0.01720   0.01144  -0.0998   0.9078   0.7888
   0.250   0.3416   0.01713   0.01145  -0.1011   0.9057   0.8166
   0.500   0.3534   0.01722   0.01165  -0.0980   0.8985   0.8539
   0.750   0.3875   0.01698   0.01158  -0.0991   0.8954   0.9412
   1.000   0.4321   0.01692   0.01146  -0.1028   0.8931   1.0000
   1.250   0.4563   0.01716   0.01165  -0.1027   0.8863   1.0000
   1.500   0.4931   0.01716   0.01163  -0.1046   0.8824   1.0000
   1.750   0.5361   0.01701   0.01149  -0.1076   0.8800   1.0000
   2.000   0.5610   0.01716   0.01164  -0.1072   0.8723   1.0000
   2.250   0.6029   0.01687   0.01137  -0.1097   0.8685   1.0000
   2.500   0.6557   0.01619   0.01076  -0.1140   0.8664   1.0000
   2.750   0.6857   0.01594   0.01061  -0.1140   0.8573   1.0000
   3.000   0.7556   0.01436   0.00915  -0.1207   0.8536   1.0000
   3.250   0.7961   0.01351   0.00839  -0.1221   0.8427   1.0000
   3.500   0.8344   0.01287   0.00788  -0.1233   0.8325   1.0000
   3.750   0.8712   0.01227   0.00738  -0.1241   0.8208   1.0000
   4.000   0.9020   0.01180   0.00702  -0.1239   0.8054   1.0000
   4.250   0.9332   0.01132   0.00662  -0.1236   0.7860   1.0000
   4.500   0.9568   0.01104   0.00641  -0.1219   0.7585   1.0000
   4.750   0.9767   0.01084   0.00629  -0.1195   0.7107   1.0000
   5.000   0.9978   0.01108   0.00594  -0.1170   0.5827   1.0000
   5.250   0.9953   0.01257   0.00650  -0.1109   0.4292   1.0000
   5.500   0.9921   0.01422   0.00736  -0.1053   0.2920   1.0000
   5.750   0.9954   0.01582   0.00829  -0.1014   0.1828   1.0000
   6.000   1.0042   0.01718   0.00925  -0.0982   0.1257   1.0000
   6.250   1.0169   0.01831   0.01022  -0.0957   0.0937   1.0000
   6.500   1.0291   0.01956   0.01134  -0.0931   0.0739   1.0000
   6.750   1.0430   0.02088   0.01264  -0.0908   0.0603   1.0000
   7.000   1.0591   0.02211   0.01388  -0.0889   0.0480   1.0000
   7.250   1.0761   0.02350   0.01531  -0.0871   0.0377   1.0000
   7.500   1.0940   0.02552   0.01735  -0.0857   0.0258   1.0000
   7.750   1.1130   0.02628   0.01837  -0.0840   0.0195   1.0000
   8.000   1.1516   0.03153   0.02388  -0.0860   0.0158   1.0000
   8.250   1.1771   0.03448   0.02724  -0.0852   0.0152   1.0000
   8.500   1.1964   0.03821   0.03145  -0.0836   0.0148   1.0000
   8.750   1.2091   0.04091   0.03455  -0.0810   0.0137   1.0000
   9.000   1.2188   0.04274   0.03666  -0.0785   0.0120   1.0000
   9.250   1.2260   0.04343   0.03743  -0.0761   0.0104   1.0000
   9.500   1.2261   0.04594   0.04015  -0.0727   0.0097   1.0000
   9.750   1.2210   0.04942   0.04401  -0.0685   0.0096   1.0000
  10.000   1.2115   0.05299   0.04794  -0.0641   0.0097   1.0000
  10.250   1.1979   0.05705   0.05238  -0.0596   0.0099   1.0000
  10.500   1.1776   0.06205   0.05778  -0.0554   0.0103   1.0000
  10.750   1.1468   0.06861   0.06478  -0.0518   0.0111   1.0000
  11.000   1.1221   0.07415   0.07058  -0.0501   0.0114   1.0000
  11.250   1.0976   0.08003   0.07668  -0.0498   0.0117   1.0000
  11.500   1.0740   0.08629   0.08312  -0.0510   0.0119   1.0000
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