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

E61 (5.64%) (e61-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: E61 (5.64%) (e61-il)
Reynolds number: 200,000
Max Cl/Cd: 110.84 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e61-il-200000-n5.txt
Download as CSV file: xf-e61-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E61  (5.64%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.1494   0.09164   0.08812  -0.0651   0.9570   0.0135
  -6.250  -0.1270   0.08656   0.08308  -0.0702   0.9455   0.0153
  -5.750  -0.0874   0.08075   0.07728  -0.0819   0.9318   0.0161
  -5.500  -0.0680   0.07725   0.07378  -0.0861   0.9262   0.0162
  -5.250  -0.0397   0.07299   0.06949  -0.0925   0.9234   0.0162
  -5.000  -0.0212   0.06955   0.06605  -0.0961   0.9167   0.0162
  -4.750  -0.0051   0.06414   0.06066  -0.0995   0.9127   0.0170
  -4.500   0.0198   0.06056   0.05705  -0.1031   0.9105   0.0178
  -4.250   0.0424   0.05744   0.05392  -0.1067   0.9047   0.0187
  -4.000   0.0772   0.05360   0.05002  -0.1136   0.9007   0.0198
  -3.750   0.1245   0.04904   0.04534  -0.1237   0.8987   0.0220
  -2.750   0.3801   0.02692   0.02208  -0.1731   0.8969   0.0225
  -2.500   0.4418   0.02247   0.01702  -0.1820   0.8970   0.0216
  -2.250   0.4897   0.02035   0.01437  -0.1866   0.8964   0.0249
  -2.000   0.5293   0.01850   0.01209  -0.1897   0.8932   0.0242
  -1.750   0.5669   0.01713   0.01034  -0.1919   0.8901   0.0234
  -1.500   0.6046   0.01603   0.00898  -0.1940   0.8878   0.0229
  -1.250   0.6423   0.01512   0.00790  -0.1960   0.8858   0.0225
  -1.000   0.6807   0.01434   0.00703  -0.1982   0.8841   0.0224
  -0.750   0.7148   0.01386   0.00645  -0.1996   0.8802   0.0227
  -0.500   0.7472   0.01352   0.00602  -0.2005   0.8751   0.0234
  -0.250   0.7843   0.01312   0.00552  -0.2024   0.8720   0.0251
   0.000   0.8244   0.01270   0.00509  -0.2049   0.8696   0.0355
   0.250   0.8580   0.01207   0.00525  -0.2067   0.8627   0.3091
   0.500   0.8985   0.01150   0.00518  -0.2096   0.8585   0.5384
   1.000   0.9560   0.01062   0.00483  -0.2093   0.8445   1.0000
   1.250   0.9878   0.01063   0.00478  -0.2101   0.8358   1.0000
   1.500   1.0244   0.01056   0.00463  -0.2118   0.8280   1.0000
   1.750   1.0562   0.01057   0.00459  -0.2126   0.8177   1.0000
   2.000   1.0869   0.01062   0.00460  -0.2131   0.8065   1.0000
   2.250   1.1175   0.01068   0.00466  -0.2135   0.7946   1.0000
   2.500   1.1474   0.01076   0.00471  -0.2139   0.7817   1.0000
   2.750   1.1764   0.01087   0.00479  -0.2140   0.7677   1.0000
   3.000   1.2045   0.01100   0.00490  -0.2139   0.7527   1.0000
   3.250   1.2319   0.01116   0.00503  -0.2137   0.7361   1.0000
   3.500   1.2575   0.01135   0.00526  -0.2131   0.7173   1.0000
   3.750   1.2824   0.01157   0.00546  -0.2124   0.6961   1.0000
   4.000   1.3063   0.01183   0.00568  -0.2114   0.6721   1.0000
   4.250   1.3288   0.01214   0.00593  -0.2101   0.6426   1.0000
   4.500   1.3494   0.01253   0.00622  -0.2085   0.6061   1.0000
   4.750   1.3677   0.01305   0.00657  -0.2064   0.5621   1.0000
   5.000   1.3833   0.01371   0.00708  -0.2038   0.5125   1.0000
   5.250   1.3929   0.01471   0.00771  -0.2002   0.4429   1.0000
   5.500   1.3967   0.01614   0.00857  -0.1956   0.3512   1.0000
   5.750   1.3979   0.01788   0.00961  -0.1910   0.2463   1.0000
   6.000   1.4034   0.01954   0.01067  -0.1873   0.1587   1.0000
   6.500   1.4274   0.02202   0.01264  -0.1821   0.0665   1.0000
   6.750   1.4346   0.02376   0.01404  -0.1788   0.0188   1.0000
   7.000   1.4481   0.02493   0.01524  -0.1762   0.0107   1.0000
   7.250   1.4620   0.02608   0.01649  -0.1738   0.0083   1.0000
   7.500   1.4761   0.02717   0.01782  -0.1715   0.0069   1.0000
   7.750   1.4893   0.02838   0.01920  -0.1691   0.0061   1.0000
   8.000   1.5011   0.02974   0.02075  -0.1665   0.0057   1.0000
   8.250   1.5106   0.03131   0.02252  -0.1636   0.0053   1.0000
   8.500   1.5183   0.03308   0.02448  -0.1606   0.0051   1.0000
   8.750   1.5242   0.03509   0.02668  -0.1574   0.0049   1.0000
   9.000   1.5302   0.03718   0.02895  -0.1544   0.0048   1.0000
   9.250   1.5353   0.03959   0.03155  -0.1513   0.0047   1.0000
   9.500   1.5417   0.04209   0.03425  -0.1486   0.0046   1.0000
   9.750   1.5495   0.04476   0.03713  -0.1461   0.0046   1.0000
  10.000   1.5582   0.04758   0.04019  -0.1438   0.0045   1.0000
  10.250   1.5664   0.05069   0.04357  -0.1416   0.0045   1.0000
  10.500   1.5729   0.05373   0.04696  -0.1393   0.0045   1.0000
  10.750   1.5758   0.05727   0.05082  -0.1368   0.0045   1.0000
  11.000   1.5752   0.06092   0.05479  -0.1343   0.0045   1.0000
  11.250   1.5722   0.06466   0.05884  -0.1319   0.0045   1.0000
  11.500   1.5663   0.06866   0.06315  -0.1297   0.0046   1.0000
  11.750   1.5581   0.07294   0.06773  -0.1278   0.0046   1.0000
  12.000   1.5464   0.07782   0.07292  -0.1262   0.0046   1.0000
  12.250   1.5347   0.08270   0.07809  -0.1253   0.0046   1.0000
  12.500   1.5207   0.08809   0.08376  -0.1249   0.0046   1.0000
  12.750   1.5067   0.09371   0.08965  -0.1253   0.0047   1.0000
  13.000   1.4906   0.10003   0.09624  -0.1266   0.0047   1.0000
  13.250   1.4750   0.10667   0.10312  -0.1287   0.0047   1.0000
<< Back to E61 (5.64%) (e61-il)

Polar data table (+)

Polar graphs


<< Back to E61 (5.64%) (e61-il)