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

GOE 289 (MVA 289) AIRFOIL (goe289-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 289 (MVA 289) AIRFOIL (goe289-il)
Reynolds number: 50,000
Max Cl/Cd: 20.65 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe289-il-50000-n5.txt
Download as CSV file: xf-goe289-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 289 (MVA 289) 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
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.2282   0.15292   0.14619  -0.0418   1.0000   0.1219
 -12.000  -0.2394   0.15240   0.14577  -0.0405   1.0000   0.1231
 -11.750  -0.2450   0.15101   0.14445  -0.0420   0.9976   0.1241
 -11.500  -0.2404   0.14821   0.14164  -0.0479   0.9907   0.1247
 -11.250  -0.2183   0.14282   0.13625  -0.0515   0.9848   0.1257
 -11.000  -0.1860   0.13732   0.13070  -0.0535   0.9790   0.1285
 -10.750  -0.1680   0.13334   0.12668  -0.0574   0.9724   0.1297
 -10.500  -0.1557   0.12955   0.12287  -0.0611   0.9646   0.1293
 -10.000  -0.1538   0.11679   0.10999  -0.0728   0.9491   0.0852
  -9.750  -0.1315   0.11271   0.10586  -0.0762   0.9438   0.0837
  -9.500  -0.1234   0.10906   0.10221  -0.0789   0.9342   0.0834
  -9.250  -0.1083   0.10473   0.09785  -0.0834   0.9285   0.0828
  -9.000  -0.1026   0.10117   0.09429  -0.0859   0.9183   0.0827
  -8.750  -0.0894   0.09669   0.08979  -0.0906   0.9125   0.0823
  -8.500  -0.0860   0.09316   0.08626  -0.0928   0.9019   0.0815
  -8.250  -0.0760   0.08840   0.08147  -0.0976   0.8957   0.0805
  -8.000  -0.0806   0.08470   0.07779  -0.0994   0.8837   0.0797
  -7.750  -0.0829   0.07973   0.07280  -0.1037   0.8750   0.0788
  -7.500  -0.1003   0.07471   0.06779  -0.1064   0.8620   0.0779
  -7.000  -0.1695   0.05969   0.05220  -0.1127   0.8338   0.0755
  -6.750  -0.1609   0.05549   0.04771  -0.1146   0.8269   0.0759
  -6.500  -0.1606   0.05339   0.04545  -0.1129   0.8155   0.0764
  -6.000  -0.1335   0.04937   0.04105  -0.1124   0.7990   0.0796
  -5.750  -0.1129   0.04630   0.03754  -0.1135   0.7932   0.0819
  -5.500  -0.1062   0.04428   0.03512  -0.1117   0.7831   0.0834
  -5.250  -0.0826   0.04143   0.03165  -0.1122   0.7772   0.0856
  -5.000  -0.0646   0.04035   0.03045  -0.1111   0.7691   0.0874
  -4.750  -0.0415   0.03918   0.02912  -0.1107   0.7618   0.0900
  -4.500  -0.0069   0.03744   0.02700  -0.1119   0.7577   0.0951
  -4.250   0.0026   0.03699   0.02641  -0.1093   0.7469   0.0989
  -4.000   0.0353   0.03593   0.02521  -0.1101   0.7419   0.1065
  -3.750   0.0517   0.03535   0.02449  -0.1083   0.7330   0.1132
  -3.500   0.0792   0.03445   0.02338  -0.1081   0.7265   0.1252
  -3.250   0.1159   0.03342   0.02229  -0.1094   0.7225   0.1436
  -3.000   0.1243   0.03354   0.02229  -0.1065   0.7114   0.1574
  -2.750   0.1577   0.03289   0.02171  -0.1075   0.7064   0.1832
  -2.500   0.1726   0.03303   0.02186  -0.1057   0.6973   0.2066
  -2.250   0.1989   0.03274   0.02156  -0.1054   0.6906   0.2308
  -2.000   0.2388   0.03201   0.02076  -0.1070   0.6866   0.2530
  -1.750   0.2451   0.03249   0.02123  -0.1038   0.6754   0.2664
  -1.500   0.2810   0.03197   0.02065  -0.1047   0.6705   0.2912
  -1.250   0.2949   0.03226   0.02094  -0.1027   0.6613   0.3134
  -1.000   0.3225   0.03204   0.02072  -0.1024   0.6548   0.3464
  -0.750   0.3620   0.03137   0.02001  -0.1038   0.6509   0.3938
  -0.500   0.3641   0.03208   0.02079  -0.1002   0.6397   0.4226
  -0.250   0.3988   0.03145   0.02027  -0.1008   0.6350   0.4751
   0.000   0.4068   0.03190   0.02090  -0.0978   0.6257   0.5165
   0.250   0.4325   0.03153   0.02080  -0.0970   0.6197   0.5902
   0.750   0.5168   0.03111   0.02091  -0.1019   0.6056   1.0000
   1.000   0.5513   0.03101   0.02047  -0.1028   0.6007   1.0000
   1.250   0.5556   0.03197   0.02129  -0.0996   0.5917   1.0000
   1.500   0.5808   0.03219   0.02127  -0.0991   0.5856   1.0000
   1.750   0.6213   0.03190   0.02070  -0.1007   0.5820   1.0000
   2.000   0.6079   0.03353   0.02230  -0.0952   0.5711   1.0000
   2.250   0.6433   0.03340   0.02194  -0.0960   0.5668   1.0000
   2.750   0.6633   0.03514   0.02347  -0.0914   0.5519   1.0000
   3.000   0.7030   0.03482   0.02295  -0.0927   0.5484   1.0000
   3.250   0.6835   0.03715   0.02529  -0.0872   0.5376   1.0000
   3.500   0.7152   0.03709   0.02508  -0.0875   0.5333   1.0000
   3.750   0.7570   0.03665   0.02447  -0.0889   0.5305   1.0000
   4.000   0.7254   0.03989   0.02777  -0.0827   0.5183   1.0000
   4.250   0.7597   0.03969   0.02744  -0.0832   0.5151   1.0000
   4.750   0.7601   0.04344   0.03115  -0.0783   0.4996   1.0000
   5.000   0.7966   0.04301   0.03060  -0.0787   0.4971   1.0000
   5.500   0.7859   0.04810   0.03570  -0.0740   0.4810   1.0000
   5.750   0.8207   0.04771   0.03521  -0.0741   0.4791   1.0000
   6.250   0.7748   0.05685   0.04446  -0.0693   0.4584   1.0000
   6.500   0.7618   0.06108   0.04873  -0.0681   0.4491   1.0000
   6.750   0.7851   0.06155   0.04915  -0.0676   0.4466   1.0000
   7.000   0.8127   0.06158   0.04914  -0.0672   0.4449   1.0000
   8.000   0.7673   0.07919   0.06691  -0.0646   0.4151   1.0000
   8.250   0.7882   0.08010   0.06781  -0.0642   0.4128   1.0000
   8.750   0.7692   0.08874   0.07653  -0.0638   0.4002   1.0000
   9.000   0.7849   0.09034   0.07813  -0.0635   0.3976   1.0000
   9.250   0.8041   0.09157   0.07938  -0.0632   0.3957   1.0000
   9.500   0.7754   0.09821   0.08610  -0.0635   0.3880   1.0000
   9.750   0.7822   0.10076   0.08869  -0.0634   0.3840   1.0000
  10.000   0.7968   0.10252   0.09047  -0.0632   0.3814   1.0000
  10.250   0.8147   0.10400   0.09197  -0.0631   0.3795   1.0000
  10.500   0.7975   0.10945   0.09750  -0.0636   0.3743   1.0000
  10.750   0.7961   0.11296   0.10107  -0.0639   0.3696   1.0000
<< Back to GOE 289 (MVA 289) AIRFOIL (goe289-il)

Polar data table (+)

Polar graphs


<< Back to GOE 289 (MVA 289) AIRFOIL (goe289-il)