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MRC-20 (mrc-20-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MRC-20 (mrc-20-il)
Reynolds number: 50,000
Max Cl/Cd: 23.9 at α=11.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mrc-20-il-50000.txt
Download as CSV file: xf-mrc-20-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MRC-20                                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3681   0.12078   0.11429  -0.0234   1.0000   0.3099
  -9.000  -0.3746   0.11923   0.11282  -0.0209   1.0000   0.3292
  -8.750  -0.3905   0.11831   0.11201  -0.0181   1.0000   0.3472
  -8.500  -0.3675   0.11525   0.10896  -0.0153   1.0000   0.3765
  -8.250  -0.3546   0.11277   0.10651  -0.0125   1.0000   0.4044
  -8.000  -0.3745   0.11256   0.10642  -0.0086   1.0000   0.4270
  -7.750  -0.3587   0.11001   0.10390  -0.0058   1.0000   0.4588
  -7.500  -0.3398   0.10702   0.10094  -0.0036   1.0000   0.4893
  -7.250  -0.3301   0.10443   0.09839  -0.0013   1.0000   0.5159
  -7.000  -0.3440   0.10411   0.09815   0.0032   1.0000   0.5426
  -6.750  -0.3162   0.09984   0.09389   0.0035   1.0000   0.5681
  -6.500  -0.3151   0.09785   0.09196   0.0061   1.0000   0.5896
  -5.000  -0.4761   0.08466   0.07949   0.0223   1.0000   0.5057
  -4.750  -0.5439   0.07998   0.07500   0.0229   1.0000   0.4643
  -4.500  -0.5918   0.05633   0.04997  -0.0128   1.0000   0.2767
  -4.250  -0.5549   0.05209   0.04413  -0.0173   1.0000   0.1999
  -4.000  -0.5266   0.05045   0.04157  -0.0169   1.0000   0.1673
  -3.750  -0.5023   0.04834   0.03895  -0.0160   1.0000   0.1497
  -3.500  -0.4798   0.04582   0.03614  -0.0154   1.0000   0.1388
  -3.250  -0.4566   0.04458   0.03446  -0.0144   1.0000   0.1304
  -3.000  -0.4339   0.04292   0.03253  -0.0136   1.0000   0.1247
  -2.750  -0.4105   0.04217   0.03121  -0.0124   1.0000   0.1193
  -2.500  -0.3887   0.04128   0.03007  -0.0115   1.0000   0.1186
  -2.250  -0.3669   0.04003   0.02874  -0.0109   1.0000   0.1200
  -2.000  -0.3450   0.03913   0.02780  -0.0101   1.0000   0.1216
  -1.750  -0.3229   0.03848   0.02712  -0.0092   1.0000   0.1228
  -1.500  -0.3011   0.03801   0.02666  -0.0082   1.0000   0.1250
  -1.250  -0.2799   0.03776   0.02635  -0.0073   1.0000   0.1305
  -1.000  -0.2588   0.03744   0.02601  -0.0067   1.0000   0.1380
  -0.750  -0.2376   0.03725   0.02573  -0.0063   1.0000   0.1480
  -0.500  -0.2149   0.03699   0.02544  -0.0065   1.0000   0.1671
  -0.250  -0.1849   0.03476   0.02589  -0.0070   1.0000   0.6489
   0.000  -0.1651   0.03495   0.02690  -0.0009   1.0000   1.0000
   0.500  -0.0589   0.03848   0.02927  -0.0144   0.9640   1.0000
   0.750  -0.0123   0.03998   0.03040  -0.0197   0.9471   1.0000
   1.000   0.0288   0.04126   0.03139  -0.0237   0.9313   1.0000
   1.250   0.0611   0.04222   0.03212  -0.0261   0.9167   1.0000
   1.500   0.0921   0.04315   0.03285  -0.0282   0.9022   1.0000
   1.750   0.1201   0.04405   0.03359  -0.0297   0.8884   1.0000
   2.000   0.1486   0.04498   0.03437  -0.0312   0.8744   1.0000
   2.250   0.1762   0.04594   0.03520  -0.0326   0.8610   1.0000
   2.500   0.2049   0.04690   0.03605  -0.0340   0.8475   1.0000
   2.750   0.2348   0.04790   0.03694  -0.0355   0.8345   1.0000
   3.000   0.2686   0.04890   0.03784  -0.0376   0.8218   1.0000
   3.250   0.3011   0.04985   0.03872  -0.0393   0.8091   1.0000
   3.500   0.3197   0.05079   0.03961  -0.0391   0.7956   1.0000
   3.750   0.3409   0.05179   0.04058  -0.0393   0.7824   1.0000
   4.000   0.3645   0.05283   0.04158  -0.0398   0.7696   1.0000
   4.250   0.3933   0.05380   0.04252  -0.0408   0.7571   1.0000
   4.500   0.4325   0.05457   0.04328  -0.0430   0.7457   1.0000
   4.750   0.4455   0.05575   0.04446  -0.0421   0.7320   1.0000
   5.000   0.4605   0.05702   0.04574  -0.0415   0.7190   1.0000
   5.250   0.4812   0.05816   0.04690  -0.0415   0.7062   1.0000
   5.500   0.5104   0.05909   0.04785  -0.0423   0.6944   1.0000
   5.750   0.5426   0.05980   0.04862  -0.0432   0.6825   1.0000
   6.000   0.5504   0.06145   0.05029  -0.0420   0.6690   1.0000
   6.250   0.5636   0.06297   0.05185  -0.0413   0.6561   1.0000
   6.500   0.5846   0.06420   0.05313  -0.0412   0.6438   1.0000
   7.000   0.6330   0.06623   0.05531  -0.0413   0.6192   1.0000
   7.250   0.6378   0.06837   0.05750  -0.0401   0.6061   1.0000
   7.500   0.6515   0.07003   0.05922  -0.0395   0.5933   1.0000
   7.750   0.6787   0.07090   0.06018  -0.0396   0.5816   1.0000
   8.000   0.7096   0.07134   0.06075  -0.0396   0.5694   1.0000
   8.250   0.7047   0.07437   0.06384  -0.0382   0.5559   1.0000
   8.500   0.7096   0.07678   0.06631  -0.0373   0.5429   1.0000
   8.750   0.7250   0.07845   0.06808  -0.0368   0.5306   1.0000
   9.000   0.7625   0.07825   0.06804  -0.0366   0.5186   1.0000
   9.250   0.7690   0.08056   0.07044  -0.0358   0.5055   1.0000
   9.500   0.7580   0.08459   0.07452  -0.0349   0.4927   1.0000
   9.750   0.7634   0.08718   0.07719  -0.0342   0.4801   1.0000
  10.000   0.7836   0.08848   0.07861  -0.0336   0.4679   1.0000
  10.250   0.8245   0.08742   0.07775  -0.0326   0.4548   1.0000
  10.500   0.7992   0.09332   0.08367  -0.0323   0.4426   1.0000
  10.750   0.7907   0.09756   0.08797  -0.0320   0.4309   1.0000
  11.500   1.1760   0.05836   0.05009  -0.0242   0.3757   1.0000
  11.750   1.2630   0.05284   0.04458  -0.0260   0.3426   1.0000
  12.000   1.1932   0.06008   0.05199  -0.0193   0.3410   1.0000
  12.250   1.2614   0.05562   0.04750  -0.0193   0.3085   1.0000
  12.500   0.8666   0.11265   0.10388  -0.0274   0.3431   1.0000
  12.750   0.8180   0.12318   0.11432  -0.0307   0.3389   1.0000
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