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

GOE 744 AIRFOIL (goe744-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 744 AIRFOIL (goe744-il)
Reynolds number: 500,000
Max Cl/Cd: 97.29 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe744-il-500000-n5.txt
Download as CSV file: xf-goe744-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 744 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2102   0.09041   0.08636  -0.0009   0.4846   0.0269
  -7.750  -0.2079   0.08730   0.08325  -0.0025   0.4831   0.0269
  -7.500  -0.2075   0.08425   0.08021  -0.0041   0.4813   0.0269
  -7.250  -0.2092   0.08138   0.07734  -0.0052   0.4798   0.0269
  -7.000  -0.2062   0.07856   0.07451  -0.0064   0.4780   0.0269
  -6.750  -0.2021   0.07490   0.07083  -0.0088   0.4767   0.0269
  -6.500  -0.1931   0.07186   0.06777  -0.0104   0.4754   0.0269
  -6.250  -0.1807   0.06972   0.06565  -0.0109   0.4737   0.0272
  -6.000  -0.1662   0.06767   0.06358  -0.0114   0.4718   0.0266
  -5.750  -0.1526   0.06525   0.06115  -0.0124   0.4697   0.0262
  -5.500  -0.1395   0.06213   0.05799  -0.0140   0.4678   0.0257
  -5.250  -0.1254   0.05863   0.05442  -0.0159   0.4660   0.0255
  -5.000  -0.1096   0.05507   0.05077  -0.0176   0.4643   0.0253
  -4.750  -0.0923   0.05149   0.04708  -0.0192   0.4625   0.0251
  -4.500  -0.0739   0.04783   0.04330  -0.0205   0.4607   0.0251
  -4.250  -0.0546   0.04407   0.03938  -0.0215   0.4587   0.0251
  -4.000  -0.0343   0.04032   0.03549  -0.0221   0.4572   0.0254
  -3.500  -0.0197   0.01933   0.01252  -0.0185   0.4553   0.0265
  -3.250   0.0054   0.01785   0.01068  -0.0178   0.4532   0.0268
  -3.000   0.0320   0.01694   0.00959  -0.0173   0.4509   0.0271
  -2.750   0.0595   0.01648   0.00906  -0.0171   0.4483   0.0274
  -2.500   0.0874   0.01614   0.00866  -0.0169   0.4457   0.0277
  -2.250   0.1154   0.01581   0.00824  -0.0166   0.4432   0.0280
  -2.000   0.1434   0.01544   0.00777  -0.0164   0.4409   0.0284
  -1.750   0.1717   0.01498   0.00725  -0.0162   0.4389   0.0288
  -1.500   0.2001   0.01455   0.00674  -0.0160   0.4362   0.0292
  -1.250   0.2285   0.01417   0.00628  -0.0158   0.4334   0.0296
  -1.000   0.2570   0.01382   0.00587  -0.0157   0.4306   0.0301
  -0.750   0.2855   0.01354   0.00551  -0.0155   0.4280   0.0305
  -0.500   0.3137   0.01327   0.00519  -0.0154   0.4253   0.0309
  -0.250   0.3420   0.01307   0.00499  -0.0152   0.4227   0.0313
   0.000   0.3705   0.01289   0.00484  -0.0151   0.4199   0.0319
   0.250   0.3989   0.01274   0.00469  -0.0150   0.4170   0.0325
   0.500   0.4272   0.01261   0.00455  -0.0149   0.4137   0.0334
   0.750   0.4554   0.01251   0.00440  -0.0148   0.4105   0.0344
   1.000   0.4833   0.01239   0.00427  -0.0146   0.4074   0.0352
   1.250   0.5115   0.01229   0.00420  -0.0145   0.4046   0.0361
   1.500   0.5396   0.01220   0.00414  -0.0144   0.4011   0.0371
   1.750   0.5675   0.01214   0.00407  -0.0143   0.3974   0.0383
   2.000   0.5953   0.01209   0.00401  -0.0142   0.3938   0.0396
   2.250   0.6229   0.01209   0.00400  -0.0141   0.3904   0.0414
   2.500   0.6509   0.01207   0.00400  -0.0140   0.3872   0.0437
   2.750   0.6785   0.01204   0.00401  -0.0139   0.3837   0.0466
   3.000   0.7061   0.01206   0.00402  -0.0138   0.3800   0.0496
   3.250   0.7333   0.01209   0.00404  -0.0136   0.3761   0.0540
   3.500   0.7605   0.01210   0.00408  -0.0135   0.3727   0.0623
   3.750   0.7863   0.01198   0.00418  -0.0132   0.3691   0.1325
   4.000   0.8128   0.01199   0.00433  -0.0130   0.3654   0.1964
   4.250   0.8396   0.01208   0.00445  -0.0129   0.3617   0.2203
   4.500   0.8658   0.01215   0.00458  -0.0127   0.3583   0.2527
   4.750   0.8847   0.01164   0.00473  -0.0113   0.3551   0.5439
   5.500   1.0738   0.01173   0.00569  -0.0333   0.3395   0.9923
   5.750   1.1089   0.01191   0.00585  -0.0350   0.3354   0.9940
   6.000   1.1443   0.01212   0.00603  -0.0369   0.3312   0.9958
   6.250   1.1803   0.01237   0.00624  -0.0390   0.3271   0.9979
   6.500   1.2156   0.01255   0.00642  -0.0408   0.3234   0.9995
   6.750   1.2444   0.01279   0.00665  -0.0414   0.3193   1.0000
   7.000   1.2686   0.01309   0.00693  -0.0412   0.3151   1.0000
   7.250   1.2922   0.01343   0.00724  -0.0409   0.3112   1.0000
   7.500   1.3163   0.01368   0.00752  -0.0406   0.3078   1.0000
   7.750   1.3397   0.01399   0.00783  -0.0403   0.3041   1.0000
   8.000   1.3622   0.01435   0.00819  -0.0399   0.3005   1.0000
   8.250   1.3836   0.01479   0.00861  -0.0394   0.2969   1.0000
   8.500   1.4053   0.01515   0.00899  -0.0389   0.2940   1.0000
   8.750   1.4265   0.01553   0.00940  -0.0384   0.2908   1.0000
   9.000   1.4461   0.01598   0.00988  -0.0377   0.2874   1.0000
   9.250   1.4636   0.01655   0.01045  -0.0369   0.2840   1.0000
   9.500   1.4780   0.01728   0.01117  -0.0359   0.2806   1.0000
   9.750   1.4934   0.01793   0.01188  -0.0350   0.2778   1.0000
  10.000   1.5032   0.01896   0.01296  -0.0341   0.2745   1.0000
  10.250   1.4970   0.02028   0.01432  -0.0309   0.2718   1.0000
  10.500   1.4949   0.02166   0.01572  -0.0285   0.2690   1.0000
  10.750   1.4943   0.02319   0.01726  -0.0266   0.2662   1.0000
  11.000   1.4984   0.02455   0.01864  -0.0252   0.2638   1.0000
  11.250   1.5051   0.02580   0.01994  -0.0241   0.2615   1.0000
  11.500   1.5103   0.02721   0.02141  -0.0230   0.2589   1.0000
  11.750   1.5136   0.02883   0.02305  -0.0220   0.2561   1.0000
  12.000   1.5151   0.03062   0.02486  -0.0210   0.2532   1.0000
  12.250   1.5151   0.03258   0.02682  -0.0201   0.2502   1.0000
  12.500   1.5207   0.03414   0.02844  -0.0194   0.2481   1.0000
  12.750   1.5263   0.03573   0.03009  -0.0188   0.2460   1.0000
  13.000   1.5309   0.03743   0.03184  -0.0182   0.2436   1.0000
  13.250   1.5341   0.03931   0.03375  -0.0177   0.2411   1.0000
  13.500   1.5372   0.04124   0.03571  -0.0173   0.2386   1.0000
  13.750   1.5382   0.04341   0.03788  -0.0169   0.2357   1.0000
  14.000   1.5424   0.04536   0.03989  -0.0167   0.2330   1.0000
  14.250   1.5463   0.04739   0.04198  -0.0165   0.2300   1.0000
  14.500   1.5479   0.04968   0.04431  -0.0164   0.2266   1.0000
  14.750   1.5485   0.05207   0.04670  -0.0162   0.2231   1.0000
  15.000   1.5484   0.05453   0.04917  -0.0161   0.2200   1.0000
  15.250   1.5526   0.05666   0.05138  -0.0162   0.2171   1.0000
  15.500   1.5558   0.05889   0.05367  -0.0162   0.2147   1.0000
  15.750   1.5572   0.06131   0.05614  -0.0163   0.2116   1.0000
  16.000   1.5579   0.06382   0.05866  -0.0163   0.2084   1.0000
  16.250   1.5573   0.06650   0.06135  -0.0165   0.2055   1.0000
  16.500   1.5606   0.06882   0.06376  -0.0167   0.2030   1.0000
  16.750   1.5610   0.07149   0.06649  -0.0169   0.1996   1.0000
  17.000   1.5610   0.07420   0.06925  -0.0172   0.1969   1.0000
  17.250   1.5620   0.07680   0.07187  -0.0175   0.1942   1.0000
  17.500   1.5607   0.07968   0.07478  -0.0179   0.1913   1.0000
  17.750   1.5610   0.08246   0.07765  -0.0183   0.1884   1.0000
  18.000   1.5621   0.08512   0.08038  -0.0187   0.1856   1.0000
  18.250   1.5564   0.08869   0.08399  -0.0193   0.1817   1.0000
  18.500   1.5527   0.09199   0.08730  -0.0200   0.1780   1.0000
  18.750   1.5497   0.09527   0.09067  -0.0206   0.1745   1.0000
  19.000   1.5450   0.09879   0.09423  -0.0214   0.1702   1.0000
<< Back to GOE 744 AIRFOIL (goe744-il)

Polar data table (+)

Polar graphs


<< Back to GOE 744 AIRFOIL (goe744-il)