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GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 587 AIRFOIL (goe587-il)
Reynolds number: 1,000,000
Max Cl/Cd: 77.94 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe587-il-1000000.txt
Download as CSV file: xf-goe587-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 587 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4621   0.08879   0.08721  -0.0132   1.0000   0.0064
  -7.750  -0.4649   0.08567   0.08412  -0.0146   1.0000   0.0064
  -7.500  -0.4632   0.08217   0.08062  -0.0164   1.0000   0.0064
  -7.250  -0.4582   0.07859   0.07704  -0.0184   1.0000   0.0064
  -7.000  -0.4608   0.07280   0.07125  -0.0208   1.0000   0.0066
  -6.750  -0.4567   0.06868   0.06712  -0.0222   1.0000   0.0068
  -6.500  -0.4484   0.06548   0.06389  -0.0230   1.0000   0.0070
  -6.250  -0.4389   0.06232   0.06070  -0.0236   1.0000   0.0071
  -6.000  -0.4283   0.05925   0.05759  -0.0241   1.0000   0.0073
  -5.750  -0.4167   0.05595   0.05424  -0.0244   1.0000   0.0075
  -5.500  -0.4044   0.05289   0.05112  -0.0244   1.0000   0.0077
  -4.000  -0.2234   0.02390   0.02087  -0.0377   0.9726   0.0095
  -3.750  -0.1953   0.01762   0.01404  -0.0371   0.9623   0.0096
  -3.500  -0.1609   0.01325   0.00907  -0.0381   0.9527   0.0098
  -3.250  -0.1214   0.01138   0.00687  -0.0405   0.9394   0.0104
  -3.000  -0.0840   0.00951   0.00464  -0.0424   0.9189   0.0107
  -2.750  -0.0559   0.00838   0.00320  -0.0422   0.8940   0.0117
  -2.500  -0.0313   0.00777   0.00243  -0.0414   0.8706   0.0144
  -2.250  -0.0064   0.00758   0.00212  -0.0406   0.8474   0.0165
  -2.000   0.0193   0.00762   0.00209  -0.0401   0.8231   0.0185
  -1.750   0.0428   0.00721   0.00155  -0.0391   0.7990   0.0261
  -1.500   0.0683   0.00726   0.00149  -0.0385   0.7757   0.0296
  -1.250   0.0928   0.00710   0.00124  -0.0378   0.7538   0.0386
  -1.000   0.1186   0.00712   0.00118  -0.0374   0.7339   0.0416
  -0.750   0.1439   0.00700   0.00097  -0.0368   0.7157   0.0472
  -0.500   0.1695   0.00695   0.00086  -0.0363   0.6990   0.0550
  -0.250   0.1946   0.00679   0.00079  -0.0358   0.6806   0.1032
   0.000   0.2186   0.00663   0.00075  -0.0351   0.6513   0.1855
   0.250   0.2358   0.00568   0.00072  -0.0335   0.6246   0.5620
   0.500   0.2453   0.00479   0.00079  -0.0293   0.6030   0.9019
   0.750   0.2686   0.00487   0.00084  -0.0281   0.5829   0.9295
   1.000   0.2925   0.00499   0.00090  -0.0270   0.5628   0.9474
   1.250   0.3181   0.00511   0.00096  -0.0264   0.5443   0.9605
   1.500   0.3487   0.00528   0.00105  -0.0268   0.5250   0.9724
   1.750   0.4000   0.00560   0.00126  -0.0319   0.5021   0.9830
   2.000   0.4363   0.00576   0.00131  -0.0340   0.4763   0.9843
   2.250   0.4684   0.00601   0.00136  -0.0352   0.4219   0.9856
   2.500   0.4953   0.00677   0.00154  -0.0356   0.2831   0.9874
   2.750   0.5168   0.00816   0.00202  -0.0352   0.0506   0.9896
   3.000   0.5501   0.00837   0.00219  -0.0365   0.0401   0.9920
   3.250   0.5806   0.00854   0.00241  -0.0373   0.0379   0.9940
   3.500   0.6101   0.00876   0.00265  -0.0378   0.0334   0.9949
   3.750   0.6388   0.00920   0.00318  -0.0382   0.0275   0.9959
   4.000   0.6687   0.00932   0.00330  -0.0389   0.0265   0.9968
   4.250   0.6982   0.00950   0.00349  -0.0395   0.0237   0.9979
   4.500   0.7254   0.01014   0.00421  -0.0397   0.0170   0.9991
   4.750   0.7557   0.01017   0.00422  -0.0405   0.0155   0.9999
   5.000   0.7771   0.01044   0.00452  -0.0393   0.0134   1.0000
   5.250   0.7962   0.01075   0.00484  -0.0377   0.0118   1.0000
   5.500   0.8032   0.01215   0.00643  -0.0335   0.0098   1.0000
   5.750   0.8237   0.01223   0.00652  -0.0322   0.0091   1.0000
   6.000   0.8387   0.01287   0.00723  -0.0297   0.0086   1.0000
   6.250   0.8550   0.01353   0.00797  -0.0275   0.0079   1.0000
   6.500   0.8726   0.01422   0.00873  -0.0256   0.0074   1.0000
   6.750   0.8915   0.01488   0.00945  -0.0240   0.0069   1.0000
   7.000   0.9107   0.01557   0.01020  -0.0226   0.0063   1.0000
   7.250   0.9183   0.02014   0.01516  -0.0188   0.0055   1.0000
  15.250   0.8228   0.18271   0.18101  -0.0489   0.0054   1.0000
  15.500   0.8279   0.18732   0.18561  -0.0510   0.0054   1.0000
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