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

GOE 310 (MVA H.42) AIRFOIL (goe310-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 310 (MVA H.42) AIRFOIL (goe310-il)
Reynolds number: 50,000
Max Cl/Cd: 38.45 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe310-il-50000-n5.txt
Download as CSV file: xf-goe310-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 310 (MVA H.42) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3357   0.10462   0.09803  -0.0322   1.0000   0.1188
  -8.000  -0.3501   0.10370   0.09727  -0.0317   1.0000   0.1195
  -7.750  -0.3641   0.10248   0.09620  -0.0311   1.0000   0.1198
  -7.500  -0.3735   0.10084   0.09466  -0.0313   1.0000   0.1201
  -6.500  -0.3817   0.08466   0.07865  -0.0308   1.0000   0.0784
  -6.250  -0.3778   0.08172   0.07574  -0.0273   1.0000   0.0735
  -5.750  -0.3813   0.07363   0.06758  -0.0306   1.0000   0.0643
  -5.500  -0.3789   0.07091   0.06487  -0.0292   1.0000   0.0635
  -5.250  -0.3540   0.06602   0.05979  -0.0342   0.9932   0.0634
  -5.000  -0.3217   0.06023   0.05371  -0.0409   0.9842   0.0640
  -4.750  -0.2905   0.05504   0.04824  -0.0458   0.9755   0.0638
  -4.500  -0.2611   0.05015   0.04303  -0.0496   0.9661   0.0632
  -4.250  -0.2316   0.04460   0.03695  -0.0531   0.9568   0.0639
  -4.000  -0.1995   0.04060   0.03248  -0.0560   0.9488   0.0663
  -3.750  -0.1709   0.03731   0.02872  -0.0574   0.9390   0.0672
  -3.500  -0.1396   0.03423   0.02503  -0.0589   0.9305   0.0682
  -3.250  -0.1058   0.03186   0.02213  -0.0604   0.9223   0.0701
  -3.000  -0.0742   0.03015   0.01994  -0.0613   0.9127   0.0745
  -2.750  -0.0345   0.02840   0.01763  -0.0634   0.9061   0.0781
  -2.500  -0.0027   0.02740   0.01644  -0.0642   0.8960   0.0811
  -2.250   0.0440   0.02637   0.01510  -0.0675   0.8905   0.0866
  -2.000   0.0801   0.02567   0.01420  -0.0689   0.8806   0.0939
  -1.750   0.1257   0.02503   0.01357  -0.0721   0.8750   0.1147
  -1.500   0.1554   0.02491   0.01353  -0.0723   0.8637   0.1547
  -1.250   0.1853   0.02501   0.01345  -0.0726   0.8531   0.2015
  -1.000   0.2222   0.02488   0.01316  -0.0741   0.8459   0.2367
  -0.750   0.2495   0.02480   0.01297  -0.0740   0.8345   0.2523
  -0.500   0.2781   0.02472   0.01288  -0.0742   0.8245   0.2734
  -0.250   0.3131   0.02449   0.01259  -0.0753   0.8172   0.2934
   0.000   0.3376   0.02451   0.01266  -0.0748   0.8062   0.3205
   0.250   0.3667   0.02437   0.01256  -0.0750   0.7973   0.3494
   0.500   0.3972   0.02415   0.01232  -0.0754   0.7891   0.3679
   0.750   0.4226   0.02404   0.01224  -0.0750   0.7794   0.3901
   1.000   0.4539   0.02356   0.01192  -0.0754   0.7725   0.4285
   1.250   0.5465   0.02224   0.01170  -0.0886   0.7668   1.0000
   1.500   0.5756   0.02237   0.01164  -0.0886   0.7584   1.0000
   1.750   0.5984   0.02269   0.01184  -0.0876   0.7484   1.0000
   2.000   0.6301   0.02273   0.01175  -0.0879   0.7415   1.0000
   2.250   0.6500   0.02315   0.01212  -0.0866   0.7309   1.0000
   2.500   0.6755   0.02341   0.01233  -0.0860   0.7223   1.0000
   2.750   0.7013   0.02364   0.01252  -0.0854   0.7136   1.0000
   3.000   0.7229   0.02404   0.01291  -0.0843   0.7039   1.0000
   3.250   0.7530   0.02412   0.01299  -0.0842   0.6963   1.0000
   3.500   0.7718   0.02460   0.01351  -0.0826   0.6853   1.0000
   3.750   0.7951   0.02491   0.01384  -0.0816   0.6753   1.0000
   4.000   0.8251   0.02489   0.01383  -0.0813   0.6656   1.0000
   4.250   0.8477   0.02499   0.01397  -0.0798   0.6509   1.0000
   4.500   0.8713   0.02497   0.01396  -0.0782   0.6348   1.0000
   4.750   0.8904   0.02509   0.01411  -0.0760   0.6170   1.0000
   5.000   0.9105   0.02521   0.01431  -0.0740   0.6000   1.0000
   5.250   0.9309   0.02539   0.01455  -0.0722   0.5843   1.0000
   5.500   0.9502   0.02558   0.01482  -0.0702   0.5678   1.0000
   5.750   0.9685   0.02573   0.01503  -0.0680   0.5492   1.0000
   6.000   0.9871   0.02585   0.01523  -0.0658   0.5301   1.0000
   6.250   1.0038   0.02611   0.01558  -0.0635   0.5117   1.0000
   6.500   1.0188   0.02651   0.01612  -0.0611   0.4938   1.0000
   6.750   1.0345   0.02693   0.01670  -0.0589   0.4768   1.0000
   7.000   1.0499   0.02737   0.01734  -0.0566   0.4593   1.0000
   7.250   1.0652   0.02782   0.01794  -0.0544   0.4414   1.0000
   7.500   1.0795   0.02827   0.01852  -0.0519   0.4217   1.0000
   7.750   1.0900   0.02880   0.01915  -0.0490   0.3967   1.0000
   8.000   1.1002   0.02938   0.01977  -0.0460   0.3703   1.0000
   8.250   1.1097   0.03004   0.02046  -0.0431   0.3446   1.0000
   8.500   1.1178   0.03084   0.02133  -0.0399   0.3215   1.0000
   8.750   1.1226   0.03181   0.02231  -0.0366   0.2962   1.0000
   9.000   1.1236   0.03305   0.02348  -0.0330   0.2666   1.0000
   9.250   1.1210   0.03466   0.02496  -0.0295   0.2338   1.0000
   9.500   1.1162   0.03664   0.02679  -0.0263   0.2003   1.0000
   9.750   1.1093   0.03901   0.02899  -0.0235   0.1615   1.0000
  10.000   1.1019   0.04170   0.03146  -0.0212   0.1272   1.0000
  10.250   1.0945   0.04463   0.03416  -0.0193   0.1018   1.0000
  10.500   1.0882   0.04767   0.03703  -0.0178   0.0875   1.0000
  10.750   1.0828   0.05079   0.04009  -0.0166   0.0792   1.0000
  11.000   1.0772   0.05405   0.04331  -0.0155   0.0732   1.0000
  11.250   1.0706   0.05750   0.04672  -0.0147   0.0683   1.0000
  11.500   1.0678   0.06064   0.04997  -0.0136   0.0638   1.0000
  11.750   1.0685   0.06344   0.05281  -0.0124   0.0599   1.0000
  12.000   1.0761   0.06561   0.05497  -0.0103   0.0554   1.0000
  12.250   1.0872   0.06772   0.05730  -0.0087   0.0502   1.0000
  12.750   1.1304   0.07144   0.06149  -0.0048   0.0432   1.0000
  13.000   1.1369   0.07445   0.06474  -0.0041   0.0409   1.0000
  13.250   1.1554   0.07722   0.06749  -0.0031   0.0379   1.0000
  13.500   1.1467   0.08123   0.07191  -0.0031   0.0375   1.0000
  13.750   1.1360   0.08572   0.07674  -0.0036   0.0370   1.0000
  14.000   1.1251   0.09046   0.08178  -0.0044   0.0367   1.0000
  14.250   1.1115   0.09566   0.08725  -0.0059   0.0366   1.0000
  14.500   1.0964   0.10128   0.09313  -0.0079   0.0365   1.0000
  14.750   1.0810   0.10722   0.09929  -0.0104   0.0365   1.0000
  15.000   1.0642   0.11377   0.10604  -0.0135   0.0366   1.0000
  15.250   1.0476   0.12073   0.11313  -0.0172   0.0368   1.0000
<< Back to GOE 310 (MVA H.42) AIRFOIL (goe310-il)

Polar data table (+)

Polar graphs


<< Back to GOE 310 (MVA H.42) AIRFOIL (goe310-il)