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EH 1.5/9.0 (eh1590-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EH 1.5/9.0 (eh1590-il)
Reynolds number: 50,000
Max Cl/Cd: 30.47 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1590-il-50000.txt
Download as CSV file: xf-eh1590-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 1.5/9.0                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5424   0.10970   0.10321   0.0154   1.0000   0.2350
  -8.750  -0.5611   0.10818   0.10184   0.0127   1.0000   0.2448
  -8.500  -0.5523   0.10409   0.09778   0.0131   1.0000   0.2583
  -8.250  -0.5481   0.10038   0.09412   0.0133   1.0000   0.2721
  -8.000  -0.5493   0.09715   0.09097   0.0131   1.0000   0.2864
  -7.750  -0.5234   0.09233   0.08607   0.0157   1.0000   0.3100
  -7.500  -0.5390   0.09025   0.08413   0.0153   1.0000   0.3282
  -7.250  -0.5091   0.08572   0.07956   0.0188   1.0000   0.3622
  -7.000  -0.5000   0.08249   0.07637   0.0208   1.0000   0.3924
  -6.750  -0.4934   0.07949   0.07343   0.0228   1.0000   0.4228
  -5.750  -0.5296   0.04913   0.04160  -0.0156   1.0000   0.1315
  -5.500  -0.5134   0.04500   0.03717  -0.0149   1.0000   0.1273
  -5.250  -0.4959   0.04097   0.03225  -0.0138   1.0000   0.1193
  -5.000  -0.4757   0.03735   0.02824  -0.0126   1.0000   0.1175
  -4.750  -0.4541   0.03415   0.02453  -0.0113   1.0000   0.1174
  -4.500  -0.4321   0.03144   0.02155  -0.0101   1.0000   0.1245
  -4.250  -0.4088   0.02918   0.01881  -0.0087   1.0000   0.1346
  -4.000  -0.3832   0.02679   0.01621  -0.0075   1.0000   0.1466
  -3.750  -0.3594   0.02482   0.01430  -0.0065   1.0000   0.1729
  -3.500  -0.3322   0.02270   0.01235  -0.0055   1.0000   0.2098
  -3.250  -0.1207   0.01812   0.00989  -0.0261   1.0000   1.0000
  -3.000  -0.1031   0.01771   0.00919  -0.0254   1.0000   1.0000
  -2.750  -0.0854   0.01737   0.00857  -0.0246   1.0000   1.0000
  -2.500  -0.0679   0.01709   0.00809  -0.0238   1.0000   1.0000
  -2.250  -0.0510   0.01687   0.00771  -0.0227   1.0000   1.0000
  -2.000  -0.0351   0.01670   0.00741  -0.0215   1.0000   1.0000
  -1.750  -0.0213   0.01659   0.00722  -0.0200   1.0000   1.0000
  -1.500  -0.0119   0.01658   0.00716  -0.0178   1.0000   1.0000
  -1.250  -0.0106   0.01671   0.00724  -0.0145   1.0000   1.0000
  -1.000  -0.0194   0.01704   0.00754  -0.0099   1.0000   1.0000
  -0.750  -0.0305   0.01748   0.00792  -0.0054   1.0000   1.0000
  -0.500  -0.0368   0.01796   0.00830  -0.0018   1.0000   1.0000
  -0.250  -0.0384   0.01847   0.00870   0.0010   1.0000   1.0000
   0.000   0.0303   0.01924   0.00933  -0.0088   0.9783   1.0000
   0.250   0.1084   0.01985   0.00981  -0.0195   0.9548   1.0000
   0.500   0.1812   0.02020   0.01013  -0.0288   0.9306   1.0000
   0.750   0.2506   0.02036   0.01029  -0.0369   0.9070   1.0000
   1.000   0.3071   0.02045   0.01039  -0.0421   0.8834   1.0000
   1.250   0.3468   0.02066   0.01062  -0.0439   0.8594   1.0000
   1.500   0.3745   0.02102   0.01098  -0.0436   0.8355   1.0000
   1.750   0.4018   0.02137   0.01132  -0.0430   0.8140   1.0000
   2.000   0.4235   0.02184   0.01178  -0.0415   0.7923   1.0000
   2.250   0.4455   0.02228   0.01223  -0.0398   0.7727   1.0000
   2.500   0.4646   0.02286   0.01287  -0.0381   0.7524   1.0000
   2.750   0.4855   0.02333   0.01335  -0.0362   0.7342   1.0000
   3.000   0.5045   0.02396   0.01401  -0.0344   0.7153   1.0000
   3.250   0.5238   0.02460   0.01471  -0.0326   0.6968   1.0000
   3.500   0.5441   0.02517   0.01531  -0.0307   0.6796   1.0000
   3.750   0.5649   0.02570   0.01595  -0.0288   0.6631   1.0000
   4.000   0.5830   0.02659   0.01694  -0.0273   0.6443   1.0000
   4.250   0.6027   0.02732   0.01777  -0.0256   0.6268   1.0000
   4.500   0.6234   0.02795   0.01848  -0.0238   0.6100   1.0000
   4.750   0.6450   0.02852   0.01913  -0.0219   0.5937   1.0000
   5.000   0.6629   0.02957   0.02042  -0.0206   0.5748   1.0000
   5.250   0.6830   0.03037   0.02135  -0.0189   0.5568   1.0000
   5.500   0.7049   0.03092   0.02203  -0.0170   0.5395   1.0000
   5.750   0.7244   0.03182   0.02310  -0.0154   0.5206   1.0000
   6.000   0.7443   0.03261   0.02408  -0.0136   0.5008   1.0000
   6.250   0.7684   0.03280   0.02437  -0.0113   0.4817   1.0000
   6.500   0.7865   0.03370   0.02558  -0.0095   0.4587   1.0000
   6.750   0.8077   0.03400   0.02607  -0.0071   0.4344   1.0000
   7.000   0.8312   0.03347   0.02565  -0.0041   0.4047   1.0000
   7.250   0.8565   0.03054   0.02250   0.0005   0.3562   1.0000
   7.500   0.8683   0.02850   0.02023   0.0051   0.2845   1.0000
   7.750   0.8659   0.02981   0.02060   0.0097   0.1816   1.0000
   8.000   0.8724   0.03253   0.02289   0.0123   0.1329   1.0000
   8.250   0.8903   0.03537   0.02546   0.0141   0.1124   1.0000
   8.500   0.9083   0.03802   0.02827   0.0155   0.0986   1.0000
   8.750   0.9283   0.04089   0.03119   0.0166   0.0893   1.0000
   9.000   0.9451   0.04459   0.03534   0.0179   0.0860   1.0000
   9.250   0.9561   0.04841   0.03969   0.0193   0.0843   1.0000
   9.500   0.9612   0.05256   0.04432   0.0207   0.0835   1.0000
   9.750   0.9599   0.05687   0.04908   0.0219   0.0835   1.0000
  10.000   0.9520   0.06130   0.05389   0.0229   0.0837   1.0000
  10.250   0.9382   0.06580   0.05868   0.0234   0.0842   1.0000
  10.500   0.9197   0.07010   0.06316   0.0238   0.0850   1.0000
  10.750   0.8994   0.07495   0.06813   0.0228   0.0860   1.0000
  11.000   0.8814   0.08048   0.07373   0.0205   0.0870   1.0000
  11.250   0.8690   0.08637   0.07965   0.0179   0.0880   1.0000
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