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

GOE 223 (MVA H.34) AIRFOIL (goe223-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 223 (MVA H.34) AIRFOIL (goe223-il)
Reynolds number: 50,000
Max Cl/Cd: 5.02 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe223-il-50000.txt
Download as CSV file: xf-goe223-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 223 (MVA H.34) 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3020   0.12730   0.12209  -0.0237   0.9795   0.3155
  -7.500  -0.3992   0.11106   0.10596  -0.0368   0.9734   0.1861
  -7.250  -0.4787   0.09493   0.08996  -0.0501   0.9670   0.1571
  -7.000  -0.5190   0.07290   0.06734  -0.0760   0.9606   0.1405
  -6.750  -0.4907   0.06483   0.05869  -0.0870   0.9513   0.1357
  -6.500  -0.4473   0.05805   0.05095  -0.0983   0.9418   0.1323
  -6.250  -0.4119   0.05512   0.04754  -0.1030   0.9308   0.1339
  -6.000  -0.3686   0.05281   0.04472  -0.1079   0.9211   0.1363
  -5.750  -0.3313   0.05115   0.04267  -0.1110   0.9114   0.1378
  -5.500  -0.2999   0.05012   0.04133  -0.1126   0.9020   0.1398
  -5.250  -0.2595   0.04921   0.04022  -0.1150   0.8937   0.1441
  -5.000  -0.2395   0.04876   0.03990  -0.1144   0.8842   0.1496
  -4.750  -0.1972   0.04833   0.03929  -0.1168   0.8768   0.1580
  -4.500  -0.1823   0.04814   0.03930  -0.1159   0.8686   0.1645
  -4.250  -0.1426   0.04761   0.03893  -0.1190   0.8618   0.1809
  -4.000  -0.1124   0.04694   0.03889  -0.1219   0.8546   0.2234
  -3.750  -0.0933   0.04961   0.04244  -0.1190   0.8471   0.3838
  -3.500  -0.0898   0.05374   0.04658  -0.1110   0.8412   0.4396
  -3.250  -0.0931   0.05620   0.04902  -0.1049   0.8368   0.4670
  -3.000  -0.0938   0.05855   0.05141  -0.0981   0.8330   0.4884
  -2.750  -0.0844   0.06081   0.05361  -0.0928   0.8290   0.5188
  -2.500  -0.0836   0.06286   0.05572  -0.0852   0.8258   0.5427
  -2.250  -0.1763   0.06134   0.05442  -0.0860   0.9547   0.4932
  -2.000  -0.1694   0.06322   0.05624  -0.0819   0.9550   0.5188
  -1.750  -0.1719   0.06473   0.05782  -0.0751   0.9545   0.5402
  -1.500  -0.1717   0.06573   0.05883  -0.0696   0.9530   0.5619
  -1.250  -0.1627   0.06659   0.05961  -0.0664   0.9480   0.5838
  -1.000  -0.1440   0.06795   0.06086  -0.0649   0.9427   0.6064
  -0.750  -0.1277   0.06943   0.06226  -0.0626   0.9395   0.6273
  -0.500  -0.1207   0.06934   0.06211  -0.0600   0.9330   0.6435
  -0.250  -0.1005   0.07027   0.06294  -0.0595   0.9276   0.6624
   0.000  -0.0653   0.07213   0.06461  -0.0631   0.9243   0.6754
   0.250  -0.0554   0.07165   0.06405  -0.0626   0.9179   0.6796
   0.500  -0.0181   0.07301   0.06519  -0.0675   0.9110   0.6845
   0.750   0.0182   0.07476   0.06675  -0.0726   0.9072   0.6888
   1.000   0.0324   0.07473   0.06664  -0.0727   0.8991   0.6918
   1.250   0.0676   0.07646   0.06823  -0.0766   0.8939   0.6969
   1.500   0.0885   0.07741   0.06907  -0.0785   0.8885   0.7015
   1.750   0.1184   0.07865   0.07018  -0.0818   0.8799   0.7062
   2.000   0.1576   0.08132   0.07274  -0.0862   0.8760   0.7118
   2.250   0.1628   0.08101   0.07240  -0.0852   0.8672   0.7158
   2.500   0.2028   0.08348   0.07475  -0.0901   0.8610   0.7225
   2.750   0.2121   0.08398   0.07523  -0.0895   0.8534   0.7268
   3.000   0.2457   0.08605   0.07721  -0.0930   0.8457   0.7335
   3.250   0.2627   0.08737   0.07848  -0.0941   0.8383   0.7391
   3.500   0.2957   0.08947   0.08052  -0.0970   0.8284   0.7469
   3.750   0.3109   0.09050   0.08151  -0.0976   0.8160   0.7532
   4.000   0.3344   0.09230   0.08329  -0.0990   0.8053   0.7600
   4.250   0.3819   0.09570   0.08660  -0.1041   0.7932   0.7710
   4.500   0.3860   0.09585   0.08679  -0.1025   0.7794   0.7774
   4.750   0.3997   0.09740   0.08833  -0.1030   0.7677   0.7857
   5.000   0.4336   0.10031   0.09124  -0.1058   0.7584   0.7969
   5.250   0.4555   0.10201   0.09297  -0.1070   0.7442   0.8078
   5.500   0.4636   0.10331   0.09431  -0.1065   0.7303   0.8182
   5.750   0.4759   0.10509   0.09616  -0.1066   0.7169   0.8305
   6.000   0.4924   0.10720   0.09835  -0.1072   0.7041   0.8461
   6.250   0.5252   0.11010   0.10136  -0.1093   0.6930   0.8718
   6.500   0.5313   0.11117   0.10260  -0.1082   0.6791   0.9083
   6.750   0.5317   0.11276   0.10425  -0.1080   0.6667   1.0000
   7.000   0.5563   0.11643   0.10785  -0.1115   0.6583   1.0000
   7.250   0.5875   0.11972   0.11106  -0.1152   0.6464   1.0000
   7.500   0.5883   0.12187   0.11318  -0.1156   0.6360   1.0000
   7.750   0.6361   0.12667   0.11786  -0.1207   0.6289   1.0000
   8.000   0.6222   0.12769   0.11887  -0.1195   0.6193   1.0000
   8.250   0.6570   0.13169   0.12277  -0.1227   0.6136   1.0000
<< Back to GOE 223 (MVA H.34) AIRFOIL (goe223-il)

Polar data table (+)

Polar graphs


<< Back to GOE 223 (MVA H.34) AIRFOIL (goe223-il)