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

GOE 396 AIRFOIL (goe396-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 396 AIRFOIL (goe396-il)
Reynolds number: 500,000
Max Cl/Cd: 121.3 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe396-il-500000.txt
Download as CSV file: xf-goe396-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 396 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2378   0.09443   0.09239  -0.0325   1.0000   0.0101
  -8.500  -0.2378   0.09179   0.08979  -0.0321   1.0000   0.0101
  -8.250  -0.2401   0.08955   0.08760  -0.0310   1.0000   0.0101
  -8.000  -0.2378   0.08662   0.08469  -0.0315   0.9989   0.0101
  -7.750  -0.2286   0.07986   0.07796  -0.0340   0.9966   0.0105
  -7.500  -0.2139   0.07539   0.07349  -0.0364   0.9935   0.0109
  -7.250  -0.1998   0.07135   0.06944  -0.0396   0.9885   0.0111
  -7.000  -0.1839   0.06719   0.06526  -0.0435   0.9838   0.0114
  -6.750  -0.1694   0.06322   0.06129  -0.0473   0.9760   0.0118
  -6.500  -0.1502   0.05878   0.05684  -0.0526   0.9703   0.0123
  -6.250  -0.1308   0.05442   0.05247  -0.0579   0.9616   0.0131
  -6.000  -0.1079   0.04992   0.04794  -0.0639   0.9538   0.0138
  -5.750  -0.0846   0.04562   0.04359  -0.0699   0.9429   0.0151
  -5.500  -0.1173   0.06094   0.05880  -0.0722   0.9597   0.0141
  -5.250  -0.0689   0.05671   0.05448  -0.0831   0.9522   0.0162
  -3.000   0.1998   0.01786   0.01365  -0.1119   0.8621   0.0173
  -2.500   0.2493   0.01199   0.00659  -0.1105   0.8468   0.0179
  -2.250   0.2749   0.01130   0.00576  -0.1100   0.8380   0.0204
  -2.000   0.3012   0.01086   0.00515  -0.1095   0.8291   0.0237
  -1.750   0.3277   0.01034   0.00441  -0.1088   0.8200   0.0262
  -1.500   0.3527   0.00922   0.00309  -0.1080   0.8108   0.0313
  -1.250   0.3792   0.00905   0.00284  -0.1076   0.8029   0.0379
  -1.000   0.4048   0.00855   0.00225  -0.1070   0.7947   0.0439
  -0.750   0.4309   0.00838   0.00204  -0.1065   0.7856   0.0508
  -0.500   0.4569   0.00817   0.00176  -0.1060   0.7767   0.0562
  -0.250   0.4826   0.00797   0.00158  -0.1054   0.7663   0.0746
   0.000   0.5084   0.00785   0.00156  -0.1049   0.7547   0.1294
   0.250   0.5346   0.00782   0.00151  -0.1044   0.7434   0.1466
   0.500   0.5607   0.00778   0.00147  -0.1040   0.7328   0.1619
   0.750   0.5866   0.00774   0.00144  -0.1036   0.7216   0.1804
   1.000   0.6124   0.00769   0.00143  -0.1031   0.7096   0.2084
   1.250   0.6378   0.00760   0.00146  -0.1026   0.6972   0.2653
   1.500   0.6569   0.00677   0.00155  -0.1011   0.6846   0.6796
   1.750   0.7179   0.00628   0.00154  -0.1087   0.6687   1.0000
   2.000   0.7429   0.00640   0.00158  -0.1080   0.6532   1.0000
   2.250   0.7681   0.00653   0.00163  -0.1074   0.6390   1.0000
   2.500   0.7931   0.00666   0.00170  -0.1068   0.6246   1.0000
   2.750   0.8181   0.00680   0.00179  -0.1062   0.6103   1.0000
   3.000   0.8430   0.00695   0.00189  -0.1055   0.5949   1.0000
   3.250   0.8657   0.00720   0.00202  -0.1044   0.5620   1.0000
   3.500   0.8889   0.00745   0.00215  -0.1035   0.5302   1.0000
   3.750   0.9104   0.00781   0.00231  -0.1023   0.4833   1.0000
   4.000   0.9302   0.00836   0.00255  -0.1008   0.4201   1.0000
   4.250   0.9478   0.00916   0.00291  -0.0991   0.3322   1.0000
   4.500   0.9554   0.01105   0.00380  -0.0960   0.1389   1.0000
   4.750   0.9748   0.01181   0.00433  -0.0946   0.0892   1.0000
   5.000   0.9919   0.01284   0.00501  -0.0927   0.0196   1.0000
   5.250   1.0144   0.01328   0.00558  -0.0916   0.0174   1.0000
   5.500   1.0359   0.01384   0.00626  -0.0903   0.0159   1.0000
   5.750   1.0558   0.01455   0.00708  -0.0888   0.0144   1.0000
   6.000   1.0729   0.01551   0.00814  -0.0868   0.0127   1.0000
   6.250   1.0864   0.01682   0.00957  -0.0842   0.0119   1.0000
   6.500   1.0945   0.01885   0.01176  -0.0807   0.0113   1.0000
   6.750   1.1095   0.02063   0.01361  -0.0784   0.0112   1.0000
   7.000   1.1300   0.02173   0.01475  -0.0770   0.0116   1.0000
   7.250   1.1520   0.02229   0.01538  -0.0758   0.0123   1.0000
<< Back to GOE 396 AIRFOIL (goe396-il)

Polar data table (+)

Polar graphs


<< Back to GOE 396 AIRFOIL (goe396-il)