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

GOE 184 (MVA H.29) AIRFOIL (goe184-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 184 (MVA H.29) AIRFOIL (goe184-il)
Reynolds number: 500,000
Max Cl/Cd: 102.74 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe184-il-500000-n5.txt
Download as CSV file: xf-goe184-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 184 (MVA H.29) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3445   0.09979   0.09756  -0.0214   1.0000   0.0102
  -9.000  -0.3432   0.09529   0.09308  -0.0236   1.0000   0.0105
  -8.750  -0.3460   0.08991   0.08772  -0.0262   1.0000   0.0109
  -7.250  -0.2735   0.02014   0.01605  -0.1021   0.8984   0.0182
  -7.000  -0.2482   0.01916   0.01489  -0.1022   0.8796   0.0187
  -6.750  -0.2222   0.01865   0.01427  -0.1021   0.8625   0.0192
  -6.500  -0.1960   0.01789   0.01330  -0.1021   0.8466   0.0199
  -6.250  -0.1694   0.01702   0.01219  -0.1021   0.8316   0.0207
  -6.000  -0.1425   0.01626   0.01114  -0.1021   0.8174   0.0217
  -5.750  -0.1152   0.01566   0.01026  -0.1021   0.8030   0.0225
  -5.500  -0.0878   0.01520   0.00955  -0.1019   0.7880   0.0230
  -5.250  -0.0612   0.01400   0.00812  -0.1021   0.7736   0.0239
  -5.000  -0.0340   0.01346   0.00744  -0.1020   0.7602   0.0247
  -4.500   0.0215   0.01252   0.00618  -0.1020   0.7362   0.0260
  -4.250   0.0494   0.01212   0.00565  -0.1019   0.7243   0.0267
  -4.000   0.0773   0.01179   0.00517  -0.1018   0.7116   0.0275
  -3.750   0.1052   0.01149   0.00473  -0.1017   0.6984   0.0280
  -3.500   0.1332   0.01120   0.00432  -0.1016   0.6847   0.0283
  -3.250   0.1613   0.01094   0.00397  -0.1016   0.6714   0.0286
  -3.000   0.1895   0.01072   0.00365  -0.1015   0.6575   0.0288
  -2.750   0.2175   0.01041   0.00325  -0.1014   0.6409   0.0294
  -2.500   0.2454   0.01017   0.00290  -0.1014   0.6207   0.0304
  -2.250   0.2732   0.01003   0.00264  -0.1013   0.5976   0.0312
  -2.000   0.3010   0.00996   0.00243  -0.1011   0.5778   0.0319
  -1.750   0.3289   0.00990   0.00226  -0.1011   0.5621   0.0326
  -1.500   0.3569   0.00985   0.00212  -0.1010   0.5493   0.0333
  -1.250   0.3850   0.00981   0.00200  -0.1009   0.5389   0.0342
  -1.000   0.4130   0.00980   0.00191  -0.1008   0.5288   0.0351
  -0.750   0.4409   0.00980   0.00185  -0.1008   0.5182   0.0363
  -0.500   0.4690   0.00977   0.00180  -0.1007   0.5085   0.0400
  -0.250   0.4968   0.00978   0.00187  -0.1007   0.4999   0.0623
   0.000   0.5250   0.00983   0.00189  -0.1006   0.4924   0.0682
   0.250   0.5530   0.00987   0.00192  -0.1006   0.4851   0.0721
   0.500   0.5811   0.00991   0.00194  -0.1005   0.4778   0.0753
   0.750   0.6090   0.00997   0.00196  -0.1005   0.4693   0.0770
   1.000   0.6370   0.01001   0.00199  -0.1004   0.4599   0.0783
   1.250   0.6648   0.01005   0.00200  -0.1004   0.4495   0.0800
   1.500   0.6926   0.01009   0.00202  -0.1004   0.4391   0.0821
   1.750   0.7205   0.01012   0.00205  -0.1003   0.4294   0.0853
   2.000   0.7481   0.01019   0.00210  -0.1003   0.4175   0.0893
   2.500   0.8024   0.01046   0.00223  -0.1000   0.3805   0.0944
   2.750   0.8294   0.01061   0.00234  -0.0999   0.3650   0.0978
   3.000   0.8566   0.01073   0.00245  -0.0998   0.3544   0.1040
   3.250   0.8838   0.01080   0.00261  -0.0998   0.3459   0.1513
   3.500   0.9114   0.01073   0.00281  -0.0999   0.3391   0.2748
   3.750   0.9379   0.01004   0.00308  -0.1002   0.3326   0.7275
   4.000   0.9597   0.00969   0.00319  -0.0986   0.3272   1.0000
   4.250   0.9867   0.00988   0.00335  -0.0984   0.3202   1.0000
   4.500   1.0136   0.01008   0.00352  -0.0983   0.3142   1.0000
   4.750   1.0407   0.01025   0.00369  -0.0981   0.3082   1.0000
   5.000   1.0674   0.01045   0.00388  -0.0980   0.3022   1.0000
   5.250   1.0942   0.01065   0.00408  -0.0978   0.2950   1.0000
   5.500   1.1203   0.01091   0.00429  -0.0976   0.2842   1.0000
   5.750   1.1465   0.01116   0.00452  -0.0974   0.2705   1.0000
   6.000   1.1722   0.01145   0.00476  -0.0971   0.2536   1.0000
   6.250   1.1971   0.01184   0.00505  -0.0968   0.2270   1.0000
   6.500   1.2174   0.01279   0.00565  -0.0959   0.1714   1.0000
   6.750   1.2408   0.01334   0.00613  -0.0954   0.1557   1.0000
   7.000   1.2647   0.01379   0.00655  -0.0949   0.1444   1.0000
   7.250   1.2889   0.01419   0.00694  -0.0944   0.1324   1.0000
   7.500   1.3113   0.01479   0.00742  -0.0938   0.1079   1.0000
   7.750   1.3315   0.01560   0.00806  -0.0929   0.0835   1.0000
   8.000   1.3539   0.01614   0.00860  -0.0922   0.0724   1.0000
   8.250   1.3700   0.01732   0.00954  -0.0907   0.0354   1.0000
   8.500   1.3899   0.01804   0.01026  -0.0897   0.0281   1.0000
   8.750   1.4107   0.01862   0.01091  -0.0888   0.0246   1.0000
   9.000   1.4298   0.01935   0.01168  -0.0876   0.0209   1.0000
   9.250   1.4501   0.01991   0.01230  -0.0867   0.0187   1.0000
   9.500   1.4684   0.02060   0.01303  -0.0855   0.0166   1.0000
   9.750   1.4852   0.02137   0.01385  -0.0840   0.0149   1.0000
  10.000   1.5024   0.02207   0.01463  -0.0827   0.0137   1.0000
  10.250   1.5174   0.02284   0.01546  -0.0811   0.0126   1.0000
  10.500   1.5278   0.02375   0.01641  -0.0787   0.0116   1.0000
  10.750   1.5371   0.02475   0.01749  -0.0764   0.0108   1.0000
  11.000   1.5471   0.02576   0.01859  -0.0743   0.0103   1.0000
  11.250   1.5557   0.02690   0.01982  -0.0723   0.0098   1.0000
  11.500   1.5634   0.02819   0.02121  -0.0704   0.0094   1.0000
  11.750   1.5698   0.02966   0.02277  -0.0687   0.0090   1.0000
  12.000   1.5744   0.03139   0.02459  -0.0671   0.0086   1.0000
  12.250   1.5760   0.03353   0.02682  -0.0657   0.0082   1.0000
  12.500   1.5769   0.03590   0.02930  -0.0646   0.0079   1.0000
  12.750   1.5799   0.03818   0.03170  -0.0638   0.0077   1.0000
  13.000   1.5819   0.04067   0.03431  -0.0632   0.0074   1.0000
  13.250   1.5822   0.04345   0.03722  -0.0628   0.0072   1.0000
  13.500   1.5808   0.04651   0.04040  -0.0626   0.0070   1.0000
  13.750   1.5785   0.04979   0.04379  -0.0625   0.0068   1.0000
  14.000   1.5744   0.05337   0.04749  -0.0626   0.0066   1.0000
  14.250   1.5693   0.05722   0.05146  -0.0630   0.0065   1.0000
  14.500   1.5632   0.06129   0.05565  -0.0635   0.0064   1.0000
  14.750   1.5558   0.06565   0.06012  -0.0642   0.0063   1.0000
  15.000   1.5475   0.07023   0.06481  -0.0651   0.0062   1.0000
  15.250   1.5380   0.07506   0.06976  -0.0661   0.0061   1.0000
  15.500   1.5275   0.08011   0.07493  -0.0673   0.0060   1.0000
  15.750   1.5159   0.08543   0.08036  -0.0686   0.0059   1.0000
  16.000   1.5036   0.09097   0.08601  -0.0701   0.0058   1.0000
<< Back to GOE 184 (MVA H.29) AIRFOIL (goe184-il)

Polar data table (+)

Polar graphs


<< Back to GOE 184 (MVA H.29) AIRFOIL (goe184-il)