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MH 20 9.01% (mh20-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 20 9.01% (mh20-il)
Reynolds number: 200,000
Max Cl/Cd: 68.72 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh20-il-200000.txt
Download as CSV file: xf-mh20-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 20  9.01%                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5128   0.09152   0.08855  -0.0123   1.0000   0.0370
  -8.250  -0.5162   0.08666   0.08372  -0.0173   1.0000   0.0370
  -8.000  -0.5229   0.08015   0.07729  -0.0167   1.0000   0.0383
  -7.750  -0.5185   0.07688   0.07404  -0.0168   1.0000   0.0392
  -7.500  -0.5133   0.07325   0.07042  -0.0185   1.0000   0.0401
  -7.250  -0.5081   0.06915   0.06631  -0.0211   1.0000   0.0412
  -7.000  -0.5013   0.06482   0.06195  -0.0240   1.0000   0.0427
  -6.750  -0.4925   0.06017   0.05717  -0.0272   1.0000   0.0445
  -6.500  -0.4813   0.05583   0.05232  -0.0319   1.0000   0.0497
  -6.250  -0.4735   0.04872   0.04505  -0.0343   0.9880   0.0512
  -5.250  -0.3844   0.02770   0.02212  -0.0348   0.9220   0.0234
  -5.000  -0.3677   0.02454   0.01847  -0.0325   0.9095   0.0214
  -4.750  -0.3478   0.02185   0.01524  -0.0303   0.8990   0.0197
  -4.500  -0.3254   0.01970   0.01265  -0.0284   0.8894   0.0187
  -4.250  -0.3015   0.01803   0.01072  -0.0270   0.8801   0.0183
  -4.000  -0.2785   0.01674   0.00925  -0.0255   0.8721   0.0186
  -3.750  -0.2554   0.01567   0.00805  -0.0242   0.8640   0.0198
  -3.500  -0.2319   0.01481   0.00705  -0.0230   0.8562   0.0224
  -3.250  -0.2092   0.01380   0.00592  -0.0214   0.8493   0.0374
  -3.000  -0.1895   0.01236   0.00544  -0.0203   0.8417   0.2087
  -2.750  -0.1684   0.01170   0.00510  -0.0191   0.8354   0.3143
  -2.500  -0.1490   0.01088   0.00490  -0.0177   0.8279   0.4726
  -2.250  -0.1351   0.00986   0.00481  -0.0142   0.8219   0.7097
  -2.000  -0.0454   0.00964   0.00491  -0.0251   0.8181   0.9380
  -1.750   0.0098   0.00988   0.00491  -0.0301   0.8121   0.9765
  -1.500   0.0764   0.00985   0.00459  -0.0380   0.8073   0.9999
  -1.250   0.1002   0.00984   0.00445  -0.0375   0.7991   1.0000
  -1.000   0.1229   0.00984   0.00431  -0.0366   0.7923   1.0000
  -0.750   0.1469   0.00985   0.00421  -0.0359   0.7847   1.0000
  -0.500   0.1706   0.00987   0.00411  -0.0352   0.7778   1.0000
  -0.250   0.1947   0.00989   0.00404  -0.0345   0.7705   1.0000
   0.000   0.2190   0.00994   0.00400  -0.0338   0.7635   1.0000
   0.250   0.2432   0.00998   0.00396  -0.0331   0.7563   1.0000
   0.500   0.2679   0.01003   0.00397  -0.0325   0.7488   1.0000
   0.750   0.2921   0.01008   0.00393  -0.0317   0.7419   1.0000
   1.000   0.3172   0.01014   0.00397  -0.0312   0.7337   1.0000
   1.250   0.3413   0.01019   0.00394  -0.0302   0.7271   1.0000
   1.500   0.3667   0.01025   0.00400  -0.0298   0.7179   1.0000
   1.750   0.3915   0.01031   0.00404  -0.0291   0.7099   1.0000
   2.000   0.4162   0.01035   0.00405  -0.0284   0.7016   1.0000
   2.250   0.4414   0.01041   0.00411  -0.0278   0.6920   1.0000
   2.500   0.4664   0.01046   0.00415  -0.0271   0.6830   1.0000
   2.750   0.4912   0.01050   0.00416  -0.0263   0.6739   1.0000
   3.000   0.5164   0.01054   0.00427  -0.0257   0.6627   1.0000
   3.250   0.5416   0.01058   0.00434  -0.0250   0.6513   1.0000
   3.500   0.5667   0.01061   0.00440  -0.0243   0.6392   1.0000
   3.750   0.5918   0.01063   0.00445  -0.0236   0.6262   1.0000
   4.000   0.6169   0.01065   0.00454  -0.0229   0.6117   1.0000
   4.250   0.6420   0.01066   0.00459  -0.0222   0.5960   1.0000
   4.500   0.6671   0.01069   0.00463  -0.0215   0.5790   1.0000
   4.750   0.6922   0.01072   0.00474  -0.0208   0.5574   1.0000
   5.000   0.7171   0.01080   0.00485  -0.0201   0.5338   1.0000
   5.250   0.7416   0.01092   0.00498  -0.0194   0.5066   1.0000
   5.500   0.7655   0.01114   0.00515  -0.0186   0.4748   1.0000
   5.750   0.7888   0.01148   0.00549  -0.0177   0.4380   1.0000
   6.000   0.8112   0.01194   0.00584  -0.0169   0.3958   1.0000
   6.250   0.8327   0.01251   0.00627  -0.0160   0.3458   1.0000
   6.500   0.8534   0.01319   0.00678  -0.0151   0.2904   1.0000
   6.750   0.8720   0.01413   0.00745  -0.0141   0.2197   1.0000
   7.000   0.8775   0.01678   0.00898  -0.0120   0.0643   1.0000
   7.250   0.8845   0.01927   0.01121  -0.0093   0.0233   1.0000
   7.500   0.8910   0.02153   0.01356  -0.0065   0.0166   1.0000
   7.750   0.9051   0.02291   0.01507  -0.0044   0.0149   1.0000
   8.000   0.9184   0.02456   0.01683  -0.0023   0.0143   1.0000
   8.250   0.9333   0.02644   0.01884  -0.0003   0.0136   1.0000
   8.500   0.9502   0.02854   0.02110   0.0014   0.0134   1.0000
   8.750   0.9676   0.03107   0.02387   0.0030   0.0134   1.0000
   9.000   0.9830   0.03394   0.02704   0.0047   0.0137   1.0000
   9.250   0.9943   0.03708   0.03061   0.0067   0.0141   1.0000
   9.500   1.0004   0.04053   0.03444   0.0088   0.0146   1.0000
   9.750   1.0019   0.04394   0.03816   0.0110   0.0151   1.0000
  10.000   0.9962   0.04788   0.04241   0.0134   0.0156   1.0000
  10.250   1.0050   0.05158   0.04625   0.0148   0.0174   1.0000
  10.500   0.9805   0.05497   0.05027   0.0195   0.0208   1.0000
  10.750   0.9595   0.05944   0.05501   0.0208   0.0223   1.0000
  11.000   0.9432   0.06340   0.05915   0.0208   0.0229   1.0000
  11.250   0.9235   0.06821   0.06413   0.0198   0.0236   1.0000
  11.500   0.9026   0.07360   0.06966   0.0177   0.0240   1.0000
  11.750   0.8841   0.07903   0.07523   0.0147   0.0238   1.0000
  12.000   0.8629   0.08574   0.08207   0.0104   0.0237   1.0000
  12.250   0.8452   0.09274   0.08911   0.0052   0.0230   1.0000
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