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EPPLER 587 AIRFOIL (e587-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 587 AIRFOIL (e587-il)
Reynolds number: 50,000
Max Cl/Cd: 7.29 at α=11.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e587-il-50000.txt
Download as CSV file: xf-e587-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 587 AIRFOIL                              
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.4271   0.12315   0.11795  -0.0102   1.0000   0.3095
  -7.250  -0.4451   0.12170   0.11660  -0.0082   1.0000   0.3221
  -7.000  -0.4658   0.12104   0.11602  -0.0057   1.0000   0.3359
  -6.750  -0.4536   0.11804   0.11304  -0.0036   1.0000   0.3511
  -6.500  -0.4659   0.11656   0.11163  -0.0006   1.0000   0.3687
  -6.250  -0.4703   0.11469   0.10982   0.0021   1.0000   0.3854
  -6.000  -0.4809   0.11319   0.10839   0.0052   1.0000   0.4025
  -5.250  -0.6191   0.07388   0.06818  -0.0375   1.0000   0.1570
  -5.000  -0.5952   0.06558   0.05899  -0.0424   1.0000   0.1211
  -4.750  -0.5697   0.05993   0.05239  -0.0455   1.0000   0.1076
  -4.500  -0.5499   0.05634   0.04858  -0.0461   1.0000   0.1045
  -4.250  -0.5258   0.05303   0.04478  -0.0472   1.0000   0.1021
  -4.000  -0.5010   0.05044   0.04172  -0.0481   1.0000   0.1030
  -3.750  -0.4757   0.04828   0.03907  -0.0486   1.0000   0.1052
  -3.500  -0.4504   0.04647   0.03676  -0.0486   1.0000   0.1071
  -3.250  -0.4272   0.04474   0.03481  -0.0483   1.0000   0.1111
  -3.000  -0.4056   0.04364   0.03364  -0.0476   1.0000   0.1201
  -2.750  -0.3851   0.04255   0.03254  -0.0463   1.0000   0.1307
  -2.500  -0.3657   0.04177   0.03182  -0.0446   1.0000   0.1482
  -2.250  -0.3460   0.04101   0.03121  -0.0430   1.0000   0.1756
  -2.000  -0.3081   0.03786   0.03058  -0.0462   1.0000   0.4139
  -1.750  -0.3335   0.04002   0.03366  -0.0298   1.0000   0.7069
  -1.500  -0.3492   0.04116   0.03474  -0.0178   1.0000   0.7654
  -1.250  -0.3615   0.04169   0.03517  -0.0075   1.0000   0.8111
  -1.000  -0.3721   0.04173   0.03511   0.0019   1.0000   0.8487
  -0.750  -0.1823   0.04874   0.04098  -0.0194   1.0000   1.0000
  -0.500  -0.1807   0.04839   0.04049  -0.0175   1.0000   1.0000
  -0.250  -0.1792   0.04805   0.04001  -0.0155   1.0000   1.0000
   0.000  -0.1777   0.04769   0.03953  -0.0135   1.0000   1.0000
   0.250  -0.1762   0.04731   0.03903  -0.0115   1.0000   1.0000
   0.500  -0.1747   0.04694   0.03855  -0.0095   1.0000   1.0000
   0.750  -0.1730   0.04654   0.03805  -0.0075   1.0000   1.0000
   1.000  -0.1707   0.04617   0.03758  -0.0057   1.0000   1.0000
   1.250  -0.1675   0.04584   0.03715  -0.0040   1.0000   1.0000
   1.500  -0.1610   0.04572   0.03692  -0.0029   1.0000   1.0000
   1.750  -0.1393   0.04668   0.03773  -0.0048   0.9959   1.0000
   2.000  -0.1120   0.04809   0.03897  -0.0077   0.9894   1.0000
   2.250  -0.0759   0.05058   0.04126  -0.0123   0.9828   1.0000
   2.500  -0.0491   0.05168   0.04224  -0.0151   0.9735   1.0000
   2.750  -0.0220   0.05311   0.04355  -0.0179   0.9651   1.0000
   3.000   0.0159   0.05580   0.04609  -0.0227   0.9577   1.0000
   3.250   0.0436   0.05714   0.04735  -0.0255   0.9466   1.0000
   3.500   0.0684   0.05848   0.04861  -0.0279   0.9362   1.0000
   3.750   0.1018   0.06092   0.05096  -0.0318   0.9283   1.0000
   4.000   0.1354   0.06308   0.05305  -0.0357   0.9163   1.0000
   4.250   0.1559   0.06421   0.05414  -0.0373   0.9046   1.0000
   4.500   0.1830   0.06626   0.05616  -0.0401   0.8950   1.0000
   4.750   0.2242   0.06951   0.05935  -0.0451   0.8841   1.0000
   5.000   0.2412   0.07043   0.06027  -0.0461   0.8711   1.0000
   5.250   0.2624   0.07215   0.06199  -0.0479   0.8594   1.0000
   5.500   0.2948   0.07501   0.06483  -0.0516   0.8500   1.0000
   5.750   0.3272   0.07752   0.06734  -0.0549   0.8367   1.0000
   6.000   0.3420   0.07878   0.06863  -0.0557   0.8235   1.0000
   6.250   0.3613   0.08073   0.07061  -0.0573   0.8114   1.0000
   6.500   0.3904   0.08359   0.07348  -0.0603   0.8012   1.0000
   6.750   0.4280   0.08683   0.07676  -0.0641   0.7875   1.0000
   7.000   0.4384   0.08800   0.07797  -0.0644   0.7737   1.0000
   7.250   0.4518   0.08990   0.07991  -0.0652   0.7612   1.0000
   7.500   0.4718   0.09237   0.08243  -0.0669   0.7495   1.0000
   7.750   0.5041   0.09561   0.08574  -0.0700   0.7379   1.0000
   8.000   0.5334   0.09838   0.08857  -0.0724   0.7228   1.0000
   8.250   0.5405   0.09990   0.09016  -0.0725   0.7092   1.0000
   8.500   0.5501   0.10206   0.09239  -0.0731   0.6969   1.0000
   8.750   0.5670   0.10476   0.09517  -0.0744   0.6855   1.0000
   9.000   0.5909   0.10764   0.09813  -0.0763   0.6727   1.0000
   9.250   0.6140   0.11033   0.10090  -0.0779   0.6579   1.0000
   9.500   0.7247   0.10314   0.09378  -0.0772   0.5648   1.0000
   9.750   0.7301   0.10558   0.09632  -0.0773   0.5518   1.0000
  10.000   0.7380   0.10807   0.09892  -0.0777   0.5385   1.0000
  10.250   0.7519   0.11032   0.10128  -0.0782   0.5248   1.0000
  10.500   0.7690   0.11239   0.10345  -0.0787   0.5106   1.0000
  10.750   0.7875   0.11446   0.10565  -0.0792   0.4970   1.0000
  11.000   0.8086   0.11626   0.10758  -0.0797   0.4828   1.0000
  11.250   0.8347   0.11775   0.10921  -0.0802   0.4689   1.0000
  11.500   0.8646   0.11864   0.11025  -0.0803   0.4547   1.0000
  11.750   0.8440   0.12396   0.11563  -0.0812   0.4445   1.0000
  12.000   0.8525   0.12690   0.11869  -0.0817   0.4323   1.0000
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