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

MH33 (mh33-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: MH33 (mh33-il)
Reynolds number: 500,000
Max Cl/Cd: 72.88 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh33-il-500000.txt
Download as CSV file: xf-mh33-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH33                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5348   0.08196   0.07975  -0.0171   1.0000   0.0114
  -8.250  -0.5383   0.07766   0.07549  -0.0195   1.0000   0.0116
  -8.000  -0.5466   0.07310   0.07097  -0.0227   1.0000   0.0116
  -7.750  -0.5541   0.06780   0.06568  -0.0271   1.0000   0.0116
  -7.500  -0.5564   0.06229   0.06011  -0.0305   1.0000   0.0116
  -7.250  -0.5569   0.05721   0.05495  -0.0323   1.0000   0.0119
  -7.000  -0.5559   0.05239   0.05001  -0.0329   1.0000   0.0121
  -6.750  -0.5531   0.04796   0.04542  -0.0325   1.0000   0.0124
  -6.500  -0.5488   0.04384   0.04112  -0.0313   1.0000   0.0130
  -6.250  -0.5427   0.03995   0.03701  -0.0295   1.0000   0.0137
  -6.000  -0.5336   0.03641   0.03322  -0.0273   1.0000   0.0149
  -5.750  -0.5115   0.03634   0.03287  -0.0250   1.0000   0.0169
  -4.750  -0.4634   0.02282   0.01807  -0.0163   1.0000   0.0166
  -4.500  -0.4445   0.02091   0.01588  -0.0146   1.0000   0.0167
  -4.250  -0.4257   0.01652   0.01102  -0.0127   1.0000   0.0144
  -4.000  -0.4018   0.01414   0.00832  -0.0113   0.9999   0.0122
  -3.750  -0.3687   0.01240   0.00635  -0.0118   0.9980   0.0109
  -3.500  -0.3339   0.01143   0.00525  -0.0131   0.9956   0.0105
  -3.250  -0.2983   0.01063   0.00434  -0.0146   0.9933   0.0106
  -3.000  -0.2639   0.01002   0.00360  -0.0159   0.9899   0.0113
  -2.750  -0.2274   0.00956   0.00303  -0.0175   0.9865   0.0157
  -2.500  -0.1935   0.00827   0.00260  -0.0195   0.9837   0.2192
  -2.250  -0.1560   0.00761   0.00249  -0.0220   0.9814   0.3652
  -2.000  -0.1290   0.00673   0.00240  -0.0223   0.9755   0.5714
  -1.750  -0.0949   0.00622   0.00231  -0.0236   0.9717   0.6936
  -1.500  -0.0597   0.00578   0.00230  -0.0248   0.9690   0.8117
  -1.250  -0.0297   0.00562   0.00233  -0.0246   0.9637   0.8966
  -1.000   0.0082   0.00561   0.00233  -0.0262   0.9598   0.9370
  -0.750   0.0500   0.00559   0.00228  -0.0287   0.9567   0.9584
  -0.500   0.0910   0.00558   0.00222  -0.0312   0.9519   0.9709
  -0.250   0.1358   0.00553   0.00214  -0.0346   0.9458   0.9785
   0.000   0.1812   0.00547   0.00205  -0.0381   0.9389   0.9844
   0.250   0.2257   0.00538   0.00194  -0.0415   0.9283   0.9903
   0.500   0.2678   0.00530   0.00183  -0.0444   0.9124   0.9962
   0.750   0.3048   0.00521   0.00170  -0.0462   0.8903   1.0000
   1.000   0.3273   0.00518   0.00158  -0.0448   0.8596   1.0000
   1.250   0.3491   0.00522   0.00149  -0.0433   0.8215   1.0000
   1.500   0.3706   0.00533   0.00144  -0.0417   0.7763   1.0000
   1.750   0.3918   0.00551   0.00142  -0.0402   0.7279   1.0000
   2.000   0.4129   0.00573   0.00144  -0.0387   0.6769   1.0000
   2.250   0.4342   0.00598   0.00148  -0.0373   0.6256   1.0000
   2.500   0.4555   0.00625   0.00156  -0.0359   0.5739   1.0000
   2.750   0.4769   0.00655   0.00169  -0.0346   0.5211   1.0000
   3.000   0.4984   0.00687   0.00181  -0.0334   0.4680   1.0000
   3.250   0.5200   0.00723   0.00196  -0.0322   0.4134   1.0000
   3.500   0.5418   0.00758   0.00212  -0.0311   0.3629   1.0000
   3.750   0.5640   0.00793   0.00231  -0.0301   0.3179   1.0000
   4.000   0.5861   0.00832   0.00252  -0.0291   0.2705   1.0000
   4.250   0.6084   0.00871   0.00278  -0.0281   0.2287   1.0000
   4.500   0.6305   0.00914   0.00304  -0.0272   0.1849   1.0000
   4.750   0.6526   0.00961   0.00333  -0.0262   0.1423   1.0000
   5.000   0.6753   0.01003   0.00364  -0.0254   0.1121   1.0000
   5.250   0.6984   0.01047   0.00398  -0.0245   0.0871   1.0000
   5.500   0.7215   0.01092   0.00437  -0.0237   0.0652   1.0000
   5.750   0.7442   0.01148   0.00482  -0.0229   0.0432   1.0000
   6.000   0.7670   0.01206   0.00538  -0.0221   0.0276   1.0000
   6.250   0.7900   0.01263   0.00597  -0.0212   0.0201   1.0000
   6.500   0.8134   0.01314   0.00650  -0.0205   0.0156   1.0000
   6.750   0.8364   0.01375   0.00720  -0.0197   0.0119   1.0000
   7.000   0.8558   0.01507   0.00864  -0.0181   0.0055   1.0000
   7.250   0.8767   0.01609   0.00980  -0.0168   0.0046   1.0000
   7.500   0.8968   0.01724   0.01110  -0.0155   0.0041   1.0000
   7.750   0.9154   0.01870   0.01275  -0.0140   0.0038   1.0000
   8.000   0.9326   0.02051   0.01479  -0.0123   0.0037   1.0000
   8.250   0.9492   0.02252   0.01708  -0.0105   0.0037   1.0000
   8.500   0.9638   0.02498   0.01988  -0.0086   0.0037   1.0000
   8.750   0.9720   0.02880   0.02422  -0.0058   0.0039   1.0000
   9.000   0.9665   0.03492   0.03105  -0.0019   0.0044   1.0000
   9.250   0.9513   0.04128   0.03799   0.0020   0.0049   1.0000
   9.500   0.9383   0.04583   0.04287   0.0048   0.0051   1.0000
   9.750   0.9144   0.04997   0.04726   0.0084   0.0053   1.0000
  10.000   0.8976   0.05342   0.05087   0.0097   0.0053   1.0000
  10.250   0.8743   0.05860   0.05622   0.0082   0.0054   1.0000
  10.500   0.8573   0.06470   0.06245   0.0034   0.0053   1.0000
  10.750   0.8408   0.07412   0.07200  -0.0051   0.0052   1.0000
<< Back to MH33 (mh33-il)

Polar data table (+)

Polar graphs


<< Back to MH33 (mh33-il)