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GOE 440 AIRFOIL (goe440-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 440 AIRFOIL (goe440-il)
Reynolds number: 500,000
Max Cl/Cd: 92.2 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe440-il-500000.txt
Download as CSV file: xf-goe440-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 440 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250   0.2806   0.09388   0.09080  -0.1155   0.7190   0.0040
  -8.000   0.2906   0.09141   0.08827  -0.1167   0.7100   0.0041
  -7.750   0.3008   0.08893   0.08575  -0.1180   0.7012   0.0042
  -7.500   0.3111   0.08650   0.08328  -0.1193   0.6930   0.0044
  -7.250   0.3217   0.08406   0.08081  -0.1207   0.6845   0.0045
  -7.000   0.3324   0.08166   0.07838  -0.1221   0.6768   0.0047
  -6.750   0.3435   0.07929   0.07599  -0.1236   0.6689   0.0048
  -6.500   0.3549   0.07699   0.07366  -0.1253   0.6615   0.0050
  -6.250   0.3663   0.07478   0.07143  -0.1270   0.6534   0.0051
  -6.000   0.3769   0.07267   0.06930  -0.1289   0.6455   0.0052
  -5.750   0.3867   0.07073   0.06734  -0.1307   0.6375   0.0053
  -5.500   0.4021   0.06831   0.06491  -0.1339   0.6290   0.0053
  -5.250   0.4188   0.06578   0.06234  -0.1371   0.6207   0.0053
  -5.000   0.4375   0.06308   0.05961  -0.1408   0.6116   0.0054
  -4.750   0.4578   0.06028   0.05678  -0.1449   0.6029   0.0054
  -4.500   0.4734   0.05656   0.05302  -0.1484   0.5940   0.0056
  -4.250   0.4882   0.05392   0.05035  -0.1500   0.5838   0.0060
  -4.000   0.5112   0.05110   0.04747  -0.1544   0.5732   0.0064
  -3.750   0.5390   0.04805   0.04433  -0.1602   0.5618   0.0069
  -3.500   0.5725   0.04467   0.04086  -0.1673   0.5495   0.0074
  -3.250   0.6137   0.04084   0.03691  -0.1760   0.5371   0.0081
  -2.250   0.8194   0.02443   0.01928  -0.2121   0.4889   0.0122
  -2.000   0.8616   0.02225   0.01677  -0.2164   0.4786   0.0150
  -1.750   0.8981   0.02195   0.01603  -0.2172   0.4703   0.0196
  -1.500   0.9437   0.01865   0.01234  -0.2228   0.4623   0.0237
  -1.250   0.9812   0.01754   0.01078  -0.2248   0.4555   0.0315
  -1.000   1.0180   0.01599   0.00895  -0.2259   0.4489   0.0169
  -0.750   1.0529   0.01492   0.00760  -0.2268   0.4430   0.0089
  -0.500   1.0883   0.01416   0.00668  -0.2284   0.4373   0.0088
  -0.250   1.1208   0.01385   0.00616  -0.2297   0.4321   0.0119
   0.000   1.1534   0.01361   0.00592  -0.2309   0.4273   0.1014
   0.250   1.1857   0.01329   0.00618  -0.2328   0.4223   0.4054
   0.500   1.2113   0.01347   0.00625  -0.2328   0.4176   0.4343
   0.750   1.2363   0.01362   0.00636  -0.2327   0.4054   0.4632
   1.000   1.2610   0.01381   0.00657  -0.2327   0.3969   0.5149
   1.250   1.2862   0.01395   0.00678  -0.2327   0.3888   0.5824
   1.500   1.3056   0.01434   0.00700  -0.2317   0.3572   0.6310
   1.750   1.3260   0.01469   0.00733  -0.2310   0.3345   0.6960
   2.000   1.3320   0.01550   0.00804  -0.2276   0.2806   0.8444
   2.250   1.3223   0.01693   0.00909  -0.2213   0.2180   1.0000
   2.500   1.2953   0.02022   0.01173  -0.2131   0.1046   1.0000
   2.750   1.3024   0.02151   0.01291  -0.2104   0.0263   1.0000
   3.000   1.3208   0.02206   0.01347  -0.2095   0.0266   1.0000
   3.250   1.3388   0.02265   0.01407  -0.2085   0.0286   1.0000
   4.000   1.3860   0.02495   0.01658  -0.2047   0.0222   1.0000
   4.250   1.3977   0.02605   0.01779  -0.2030   0.0180   1.0000
   4.500   1.4110   0.02704   0.01884  -0.2016   0.0160   1.0000
   4.750   1.4244   0.02804   0.01989  -0.2002   0.0134   1.0000
   5.000   1.4336   0.02941   0.02137  -0.1984   0.0117   1.0000
   5.250   1.4374   0.03124   0.02333  -0.1961   0.0105   1.0000
   5.500   1.4328   0.03385   0.02610  -0.1932   0.0097   1.0000
   5.750   1.4176   0.03756   0.03001  -0.1897   0.0091   1.0000
   6.000   1.4173   0.04008   0.03263  -0.1878   0.0090   1.0000
   6.250   1.4509   0.03939   0.03185  -0.1883   0.0079   1.0000
   6.500   1.4586   0.04121   0.03374  -0.1870   0.0071   1.0000
   6.750   1.4592   0.04382   0.03645  -0.1854   0.0065   1.0000
   7.000   1.4561   0.04692   0.03966  -0.1838   0.0061   1.0000
   7.250   1.4493   0.05061   0.04348  -0.1822   0.0058   1.0000
   7.500   1.4384   0.05498   0.04799  -0.1808   0.0056   1.0000
   7.750   1.4242   0.05993   0.05309  -0.1795   0.0055   1.0000
   8.000   1.4068   0.06548   0.05879  -0.1785   0.0054   1.0000
   8.250   1.3882   0.07143   0.06488  -0.1777   0.0053   1.0000
   8.500   1.3693   0.07756   0.07115  -0.1771   0.0053   1.0000
   8.750   1.3498   0.08392   0.07764  -0.1767   0.0052   1.0000
   9.000   1.3306   0.09035   0.08418  -0.1765   0.0052   1.0000
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