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E591 (e591-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E591 (e591-il)
Reynolds number: 50,000
Max Cl/Cd: 10.01 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e591-il-50000.txt
Download as CSV file: xf-e591-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E591                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.2845   0.13271   0.12742  -0.0098   1.0000   0.2123
  -7.000  -0.3051   0.13332   0.12814  -0.0086   1.0000   0.2172
  -6.750  -0.3423   0.13584   0.13080  -0.0074   1.0000   0.2190
  -6.500  -0.3241   0.13065   0.12565  -0.0059   1.0000   0.2221
  -6.250  -0.3066   0.12763   0.12265  -0.0063   0.9971   0.2296
  -6.000  -0.3249   0.12919   0.12428  -0.0117   0.9864   0.2373
  -5.750  -0.2663   0.12157   0.11660  -0.0156   0.9771   0.2475
  -5.500  -0.2903   0.12307   0.11818  -0.0181   0.9658   0.2560
  -5.250  -0.2424   0.11645   0.11152  -0.0214   0.9563   0.2643
  -5.000  -0.2541   0.11639   0.11151  -0.0224   0.9450   0.2741
  -4.750  -0.2424   0.11315   0.10830  -0.0250   0.9356   0.2793
  -4.500  -0.2208   0.10980   0.10494  -0.0259   0.9248   0.2872
  -4.250  -0.2525   0.11115   0.10640  -0.0299   0.9123   0.2964
  -4.000  -0.2143   0.10542   0.10064  -0.0291   0.9042   0.3013
  -3.750  -0.2055   0.10323   0.09846  -0.0295   0.8935   0.3087
  -3.500  -0.2145   0.10179   0.09707  -0.0329   0.8826   0.3184
  -3.250  -0.1252   0.06923   0.06353  -0.0956   0.8795   0.1461
  -3.000   0.0056   0.05529   0.04805  -0.1277   0.8736   0.1425
  -2.750   0.0569   0.05232   0.04427  -0.1351   0.8628   0.1509
  -2.500   0.1236   0.05016   0.04170  -0.1423   0.8548   0.1677
  -2.250   0.1446   0.04985   0.04146  -0.1423   0.8424   0.1804
  -2.000   0.1917   0.04912   0.04077  -0.1456   0.8333   0.2108
  -1.750   0.2237   0.04905   0.04079  -0.1468   0.8220   0.2582
  -1.500   0.2480   0.04995   0.04214  -0.1455   0.8114   0.3273
  -1.250   0.2736   0.05099   0.04337  -0.1437   0.8012   0.4020
  -1.000   0.2839   0.05224   0.04472  -0.1407   0.7902   0.4435
  -0.750   0.3096   0.05280   0.04535  -0.1384   0.7810   0.4878
  -0.500   0.3148   0.05393   0.04653  -0.1353   0.7700   0.5123
  -0.250   0.3490   0.05405   0.04656  -0.1350   0.7615   0.5510
   0.000   0.3565   0.05519   0.04769  -0.1331   0.7503   0.5716
   0.250   0.3948   0.05510   0.04747  -0.1338   0.7420   0.6028
   0.500   0.4008   0.05645   0.04881  -0.1323   0.7311   0.6192
   0.750   0.4440   0.05635   0.04854  -0.1343   0.7227   0.6489
   1.000   0.4497   0.05796   0.05012  -0.1331   0.7119   0.6657
   1.250   0.4843   0.05801   0.05007  -0.1336   0.7037   0.6916
   1.500   0.4909   0.05972   0.05176  -0.1326   0.6936   0.7100
   1.750   0.5187   0.06010   0.05208  -0.1325   0.6851   0.7371
   2.000   0.5305   0.06161   0.05357  -0.1319   0.6761   0.7615
   2.250   0.5441   0.06264   0.05461  -0.1308   0.6672   0.7885
   2.500   0.5691   0.06278   0.05475  -0.1297   0.6599   0.8291
   2.750   0.5571   0.06528   0.05736  -0.1274   0.6508   0.8575
   3.000   0.6000   0.06434   0.05642  -0.1287   0.6433   1.0000
   3.250   0.6011   0.06854   0.06055  -0.1318   0.6328   1.0000
   3.500   0.6869   0.06864   0.06017  -0.1404   0.6236   1.0000
   3.750   0.6633   0.07363   0.06518  -0.1399   0.6147   1.0000
   4.000   0.7073   0.07467   0.06595  -0.1426   0.6061   1.0000
   4.250   0.7066   0.07838   0.06960  -0.1428   0.5989   1.0000
   4.500   0.7085   0.08181   0.07298  -0.1431   0.5920   1.0000
   4.750   0.7605   0.08203   0.07296  -0.1446   0.5833   1.0000
   5.000   0.7350   0.08747   0.07847  -0.1442   0.5790   1.0000
   5.250   0.7323   0.09145   0.08244  -0.1445   0.5751   1.0000
   5.500   0.7862   0.09147   0.08226  -0.1454   0.5641   1.0000
   5.750   0.7708   0.09661   0.08745  -0.1456   0.5625   1.0000
   6.000   0.7688   0.10106   0.09190  -0.1463   0.5625   1.0000
   6.250   0.7748   0.10526   0.09611  -0.1474   0.5635   1.0000
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