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

GOE 359 AIRFOIL (goe359-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 359 AIRFOIL (goe359-il)
Reynolds number: 500,000
Max Cl/Cd: 110.81 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe359-il-500000.txt
Download as CSV file: xf-goe359-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 359 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.2845   0.10996   0.10773  -0.0248   1.0000   0.0152
  -9.000  -0.2815   0.10769   0.10549  -0.0267   1.0000   0.0154
  -8.750  -0.2802   0.10540   0.10324  -0.0281   1.0000   0.0155
  -8.500  -0.2000   0.08920   0.08729  -0.0308   1.0000   0.0159
  -8.250  -0.1950   0.08609   0.08420  -0.0305   0.9995   0.0161
  -8.000  -0.1799   0.08192   0.08003  -0.0336   0.9965   0.0164
  -7.750  -0.1653   0.07780   0.07591  -0.0371   0.9917   0.0168
  -7.500  -0.2775   0.09154   0.08958  -0.0257   1.0000   0.0159
  -7.250  -0.2618   0.08801   0.08606  -0.0283   0.9971   0.0162
  -7.000  -0.2369   0.08392   0.08196  -0.0341   0.9907   0.0166
  -6.750  -0.2119   0.07988   0.07792  -0.0403   0.9799   0.0172
  -6.500  -0.1863   0.07574   0.07377  -0.0468   0.9624   0.0180
  -6.250  -0.1399   0.07029   0.06825  -0.0614   0.9426   0.0201
  -6.000  -0.0901   0.06437   0.06221  -0.0757   0.9231   0.0204
  -5.750  -0.0535   0.05949   0.05716  -0.0839   0.9013   0.0204
  -5.500  -0.0419   0.05478   0.05236  -0.0852   0.8807   0.0211
  -5.250  -0.0224   0.05230   0.04976  -0.0866   0.8629   0.0216
  -5.000   0.0014   0.04952   0.04687  -0.0894   0.8472   0.0224
  -4.750   0.0289   0.04643   0.04365  -0.0930   0.8330   0.0238
  -4.500   0.0769   0.04253   0.03947  -0.1004   0.8204   0.0268
  -4.250   0.1092   0.03868   0.03540  -0.1037   0.8084   0.0269
  -4.000   0.1325   0.03268   0.02918  -0.1071   0.7972   0.0280
  -3.500   0.1816   0.02930   0.02554  -0.1085   0.7721   0.0307
  -3.250   0.2182   0.02743   0.02330  -0.1095   0.7598   0.0358
  -3.000   0.2467   0.02208   0.01740  -0.1113   0.7488   0.0372
  -2.750   0.2715   0.02066   0.01593  -0.1116   0.7364   0.0385
  -2.500   0.2976   0.01957   0.01468  -0.1117   0.7240   0.0403
  -2.250   0.3253   0.01840   0.01324  -0.1117   0.7114   0.0442
  -2.000   0.3538   0.01657   0.01097  -0.1117   0.6984   0.0495
  -1.750   0.3800   0.01573   0.01004  -0.1117   0.6842   0.0520
  -1.500   0.4074   0.01512   0.00920  -0.1114   0.6702   0.0567
  -1.250   0.4343   0.01417   0.00805  -0.1114   0.6575   0.0641
  -1.000   0.4617   0.01391   0.00764  -0.1112   0.6453   0.0705
  -0.500   0.5191   0.01193   0.00500  -0.1099   0.6217   0.0481
   0.000   0.5734   0.01088   0.00377  -0.1094   0.5980   0.0460
   0.250   0.6006   0.01058   0.00339  -0.1091   0.5863   0.0460
   0.500   0.6276   0.01041   0.00316  -0.1088   0.5752   0.0466
   0.750   0.6546   0.01032   0.00301  -0.1086   0.5646   0.0472
   1.000   0.6816   0.01011   0.00276  -0.1083   0.5533   0.0472
   1.250   0.7087   0.00994   0.00256  -0.1081   0.5420   0.0477
   1.500   0.7357   0.00978   0.00234  -0.1079   0.5299   0.0497
   1.750   0.7626   0.00975   0.00225  -0.1076   0.5173   0.0526
   2.000   0.7894   0.00978   0.00221  -0.1074   0.5050   0.0559
   2.500   0.8432   0.00952   0.00239  -0.1072   0.4836   0.2878
   2.750   0.8675   0.00827   0.00255  -0.1067   0.4745   1.0000
   3.000   0.8940   0.00843   0.00262  -0.1064   0.4652   1.0000
   3.250   0.9209   0.00856   0.00272  -0.1062   0.4567   1.0000
   3.500   0.9474   0.00873   0.00283  -0.1059   0.4481   1.0000
   3.750   0.9738   0.00889   0.00294  -0.1057   0.4372   1.0000
   4.000   1.0002   0.00905   0.00306  -0.1054   0.4257   1.0000
   4.250   1.0261   0.00926   0.00320  -0.1051   0.4098   1.0000
   4.500   1.0514   0.00950   0.00334  -0.1047   0.3883   1.0000
   4.750   1.0768   0.00975   0.00350  -0.1043   0.3662   1.0000
   5.000   1.1015   0.01005   0.00369  -0.1039   0.3382   1.0000
   5.250   1.1234   0.01066   0.00401  -0.1030   0.2849   1.0000
   5.500   1.1439   0.01143   0.00448  -0.1020   0.2380   1.0000
   5.750   1.1655   0.01209   0.00492  -0.1012   0.2056   1.0000
   6.000   1.1867   0.01277   0.00535  -0.1003   0.1630   1.0000
   6.250   1.1981   0.01449   0.00642  -0.0981   0.0633   1.0000
   6.500   1.2190   0.01517   0.00702  -0.0971   0.0452   1.0000
   6.750   1.2419   0.01562   0.00744  -0.0964   0.0341   1.0000
   7.000   1.2617   0.01638   0.00813  -0.0952   0.0177   1.0000
   7.250   1.2843   0.01683   0.00866  -0.0944   0.0162   1.0000
   7.500   1.3062   0.01732   0.00927  -0.0935   0.0149   1.0000
   7.750   1.3270   0.01791   0.00995  -0.0924   0.0143   1.0000
   8.000   1.3468   0.01856   0.01071  -0.0912   0.0137   1.0000
   8.250   1.3652   0.01931   0.01156  -0.0898   0.0132   1.0000
   8.500   1.3816   0.02018   0.01253  -0.0881   0.0128   1.0000
   8.750   1.3957   0.02119   0.01366  -0.0861   0.0124   1.0000
   9.000   1.4071   0.02231   0.01487  -0.0838   0.0122   1.0000
   9.250   1.4131   0.02355   0.01620  -0.0806   0.0119   1.0000
   9.500   1.4139   0.02508   0.01782  -0.0770   0.0116   1.0000
   9.750   1.4109   0.02696   0.01979  -0.0733   0.0113   1.0000
  10.000   1.4034   0.02939   0.02232  -0.0698   0.0110   1.0000
  10.250   1.4004   0.03175   0.02476  -0.0673   0.0109   1.0000
  10.500   1.4001   0.03406   0.02714  -0.0653   0.0109   1.0000
  10.750   1.4021   0.03631   0.02945  -0.0635   0.0109   1.0000
  11.000   1.4066   0.03838   0.03156  -0.0618   0.0109   1.0000
  11.250   1.4134   0.04031   0.03355  -0.0603   0.0109   1.0000
<< Back to GOE 359 AIRFOIL (goe359-il)

Polar data table (+)

Polar graphs


<< Back to GOE 359 AIRFOIL (goe359-il)