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

GOE 403 AIRFOIL (goe403-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 403 AIRFOIL (goe403-il)
Reynolds number: 500,000
Max Cl/Cd: 92.88 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe403-il-500000-n5.txt
Download as CSV file: xf-goe403-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 403 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3537   0.09699   0.09479  -0.0244   1.0000   0.0041
  -8.500  -0.3532   0.09338   0.09122  -0.0253   1.0000   0.0042
  -8.000  -0.3502   0.08731   0.08521  -0.0263   1.0000   0.0041
  -7.750  -0.3434   0.08667   0.08459  -0.0261   1.0000   0.0053
  -7.500  -0.3462   0.08414   0.08211  -0.0257   1.0000   0.0056
  -7.000  -0.3226   0.07540   0.07340  -0.0352   0.9862   0.0055
  -6.750  -0.3008   0.06991   0.06791  -0.0431   0.9744   0.0054
  -6.500  -0.2708   0.06275   0.06070  -0.0544   0.9613   0.0054
  -6.000  -0.1996   0.01762   0.01371  -0.0971   0.9121   0.0048
  -5.750  -0.1720   0.01543   0.01103  -0.0977   0.8923   0.0051
  -5.500  -0.1455   0.01393   0.00913  -0.0978   0.8760   0.0055
  -5.250  -0.1194   0.01282   0.00767  -0.0975   0.8619   0.0061
  -5.000  -0.0932   0.01200   0.00655  -0.0972   0.8497   0.0066
  -4.750  -0.0672   0.01111   0.00540  -0.0969   0.8392   0.0083
  -4.500  -0.0408   0.01061   0.00472  -0.0965   0.8295   0.0097
  -4.250  -0.0142   0.01008   0.00399  -0.0962   0.8203   0.0121
  -4.000   0.0127   0.00982   0.00369  -0.0959   0.8118   0.0198
  -3.750   0.0397   0.00974   0.00354  -0.0956   0.8032   0.0251
  -3.500   0.0670   0.00968   0.00339  -0.0954   0.7948   0.0295
  -3.250   0.0942   0.00975   0.00333  -0.0952   0.7864   0.0320
  -3.000   0.1208   0.00942   0.00296  -0.0950   0.7774   0.0359
  -2.750   0.1477   0.00925   0.00270  -0.0947   0.7682   0.0381
  -2.500   0.1745   0.00908   0.00243  -0.0944   0.7583   0.0396
  -2.250   0.2013   0.00892   0.00216  -0.0942   0.7473   0.0407
  -2.000   0.2281   0.00878   0.00194  -0.0939   0.7353   0.0415
  -1.750   0.2549   0.00866   0.00173  -0.0936   0.7216   0.0422
  -1.500   0.2814   0.00854   0.00151  -0.0932   0.7057   0.0439
  -1.250   0.3077   0.00845   0.00135  -0.0928   0.6871   0.0484
  -1.000   0.3340   0.00840   0.00122  -0.0924   0.6650   0.0538
  -0.750   0.3599   0.00837   0.00113  -0.0920   0.6408   0.0701
  -0.500   0.3856   0.00839   0.00110  -0.0916   0.6159   0.0930
  -0.250   0.4116   0.00844   0.00107  -0.0912   0.5927   0.1068
   0.000   0.4376   0.00850   0.00107  -0.0908   0.5728   0.1184
   0.250   0.4639   0.00855   0.00107  -0.0905   0.5559   0.1318
   0.500   0.4902   0.00856   0.00110  -0.0903   0.5420   0.1571
   1.000   0.5426   0.00849   0.00122  -0.0898   0.5184   0.2666
   1.250   0.5674   0.00807   0.00134  -0.0895   0.5090   0.4882
   1.500   0.6026   0.00705   0.00144  -0.0912   0.4996   1.0000
   1.750   0.6289   0.00718   0.00151  -0.0909   0.4904   1.0000
   2.000   0.6554   0.00730   0.00159  -0.0906   0.4810   1.0000
   2.250   0.6814   0.00745   0.00167  -0.0903   0.4669   1.0000
   2.500   0.7064   0.00768   0.00177  -0.0898   0.4386   1.0000
   2.750   0.7319   0.00788   0.00189  -0.0894   0.4182   1.0000
   3.000   0.7566   0.00817   0.00202  -0.0888   0.3825   1.0000
   3.250   0.7798   0.00863   0.00221  -0.0881   0.3275   1.0000
   3.500   0.7948   0.01012   0.00285  -0.0864   0.1600   1.0000
   3.750   0.8120   0.01149   0.00364  -0.0848   0.0163   1.0000
   4.000   0.8372   0.01181   0.00404  -0.0842   0.0116   1.0000
   4.250   0.8623   0.01211   0.00441  -0.0837   0.0103   1.0000
   4.500   0.8867   0.01251   0.00490  -0.0831   0.0090   1.0000
   4.750   0.9104   0.01300   0.00547  -0.0823   0.0076   1.0000
   5.000   0.9320   0.01375   0.00631  -0.0813   0.0065   1.0000
   5.250   0.9544   0.01438   0.00702  -0.0803   0.0061   1.0000
   5.500   0.9753   0.01516   0.00788  -0.0791   0.0056   1.0000
   5.750   0.9950   0.01606   0.00886  -0.0777   0.0053   1.0000
   6.000   1.0140   0.01699   0.00989  -0.0762   0.0050   1.0000
   6.250   1.0341   0.01771   0.01063  -0.0751   0.0044   1.0000
   6.500   1.0513   0.01884   0.01184  -0.0735   0.0039   1.0000
   6.750   1.0688   0.02008   0.01318  -0.0717   0.0037   1.0000
   7.000   1.0865   0.02154   0.01474  -0.0699   0.0034   1.0000
   7.250   1.1057   0.02316   0.01645  -0.0683   0.0033   1.0000
   7.500   1.1265   0.02498   0.01839  -0.0671   0.0031   1.0000
   7.750   1.1483   0.02708   0.02063  -0.0659   0.0031   1.0000
   8.000   1.1698   0.02942   0.02321  -0.0647   0.0031   1.0000
   8.250   1.1894   0.03196   0.02598  -0.0633   0.0032   1.0000
   8.500   1.2063   0.03475   0.02903  -0.0616   0.0034   1.0000
   8.750   1.2199   0.03767   0.03222  -0.0596   0.0035   1.0000
   9.000   1.2303   0.04066   0.03546  -0.0575   0.0037   1.0000
   9.250   1.2365   0.04390   0.03894  -0.0552   0.0038   1.0000
  16.250   0.8242   0.19424   0.19248  -0.0982   0.0058   1.0000
  16.500   0.8268   0.19841   0.19666  -0.1002   0.0049   1.0000
<< Back to GOE 403 AIRFOIL (goe403-il)

Polar data table (+)

Polar graphs


<< Back to GOE 403 AIRFOIL (goe403-il)