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

GOE 174 (ALBATROS 5020) AIRFOIL (goe174-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 174 (ALBATROS 5020) AIRFOIL (goe174-il)
Reynolds number: 200,000
Max Cl/Cd: 76.57 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe174-il-200000-n5.txt
Download as CSV file: xf-goe174-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 174 (ALBATROS 5020) AIRFOIL                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.2811   0.09963   0.09630  -0.0299   1.0000   0.0227
  -8.000  -0.2842   0.09736   0.09411  -0.0295   1.0000   0.0227
  -7.750  -0.2901   0.09534   0.09217  -0.0283   1.0000   0.0227
  -7.500  -0.2798   0.09161   0.08848  -0.0319   0.9961   0.0227
  -7.250  -0.2569   0.08666   0.08352  -0.0390   0.9880   0.0228
  -6.750  -0.2206   0.07827   0.07514  -0.0455   0.9678   0.0235
  -6.500  -0.2005   0.07544   0.07229  -0.0477   0.9562   0.0252
  -6.250  -0.1691   0.07073   0.06754  -0.0571   0.9420   0.0278
  -6.000  -0.1247   0.06434   0.06101  -0.0715   0.9285   0.0287
  -5.750  -0.0868   0.05859   0.05512  -0.0806   0.9174   0.0288
  -5.250  -0.0318   0.05124   0.04764  -0.0882   0.8944   0.0315
  -5.000   0.0157   0.04601   0.04212  -0.0979   0.8803   0.0364
  -4.750   0.0402   0.04177   0.03773  -0.1007   0.8652   0.0318
  -4.500   0.0737   0.03787   0.03357  -0.1048   0.8487   0.0337
  -4.250   0.1069   0.03330   0.02863  -0.1085   0.8312   0.0334
  -4.000   0.1400   0.02793   0.02271  -0.1115   0.8128   0.0319
  -3.750   0.1688   0.02524   0.01958  -0.1125   0.7922   0.0340
  -3.500   0.1976   0.02258   0.01639  -0.1131   0.7726   0.0343
  -3.250   0.2259   0.02058   0.01390  -0.1133   0.7539   0.0345
  -3.000   0.2539   0.01927   0.01216  -0.1132   0.7343   0.0357
  -2.750   0.2816   0.01840   0.01087  -0.1130   0.7141   0.0366
  -2.250   0.3360   0.01669   0.00846  -0.1124   0.6783   0.0369
  -2.000   0.3632   0.01605   0.00756  -0.1120   0.6633   0.0371
  -1.750   0.3901   0.01520   0.00648  -0.1118   0.6501   0.0378
  -1.500   0.4170   0.01460   0.00578  -0.1116   0.6383   0.0394
  -1.250   0.4441   0.01422   0.00529  -0.1113   0.6276   0.0401
  -1.000   0.4711   0.01390   0.00486  -0.1110   0.6173   0.0403
  -0.750   0.4983   0.01362   0.00450  -0.1107   0.6068   0.0407
  -0.500   0.5255   0.01339   0.00423  -0.1105   0.5969   0.0412
   0.000   0.5800   0.01311   0.00382  -0.1101   0.5776   0.0428
   0.250   0.6072   0.01303   0.00367  -0.1099   0.5675   0.0438
   0.750   0.6614   0.01295   0.00347  -0.1094   0.5453   0.0485
   1.000   0.6885   0.01292   0.00341  -0.1092   0.5341   0.0515
   1.250   0.7154   0.01293   0.00338  -0.1089   0.5225   0.0559
   1.500   0.7419   0.01269   0.00348  -0.1088   0.5093   0.1888
   1.750   0.7682   0.01270   0.00352  -0.1084   0.4941   0.2331
   2.000   0.7918   0.01130   0.00362  -0.1076   0.4782   1.0000
   2.250   0.8178   0.01150   0.00368  -0.1072   0.4615   1.0000
   2.500   0.8437   0.01171   0.00375  -0.1068   0.4456   1.0000
   2.750   0.8694   0.01195   0.00386  -0.1063   0.4306   1.0000
   3.000   0.8951   0.01220   0.00400  -0.1059   0.4168   1.0000
   3.250   0.9206   0.01246   0.00415  -0.1055   0.4038   1.0000
   3.500   0.9460   0.01273   0.00433  -0.1050   0.3915   1.0000
   3.750   0.9715   0.01300   0.00454  -0.1047   0.3804   1.0000
   4.000   0.9970   0.01328   0.00475  -0.1043   0.3715   1.0000
   4.250   1.0225   0.01355   0.00499  -0.1039   0.3637   1.0000
   4.500   1.0479   0.01383   0.00525  -0.1035   0.3566   1.0000
   4.750   1.0730   0.01413   0.00551  -0.1030   0.3482   1.0000
   5.000   1.0980   0.01443   0.00579  -0.1026   0.3399   1.0000
   5.250   1.1227   0.01475   0.00610  -0.1021   0.3320   1.0000
   5.500   1.1478   0.01504   0.00640  -0.1017   0.3252   1.0000
   5.750   1.1726   0.01534   0.00673  -0.1013   0.3189   1.0000
   6.000   1.1971   0.01568   0.00707  -0.1008   0.3134   1.0000
   6.250   1.2220   0.01596   0.00743  -0.1003   0.3070   1.0000
   6.500   1.2458   0.01632   0.00779  -0.0998   0.3006   1.0000
   6.750   1.2704   0.01661   0.00817  -0.0993   0.2939   1.0000
   7.000   1.2941   0.01696   0.00857  -0.0987   0.2871   1.0000
   7.250   1.3179   0.01730   0.00899  -0.0981   0.2818   1.0000
   7.500   1.3414   0.01763   0.00940  -0.0975   0.2731   1.0000
   7.750   1.3637   0.01800   0.00980  -0.0967   0.2595   1.0000
   8.000   1.3854   0.01841   0.01025  -0.0959   0.2438   1.0000
   8.250   1.4064   0.01887   0.01073  -0.0950   0.2274   1.0000
   8.500   1.4257   0.01945   0.01127  -0.0938   0.2042   1.0000
   8.750   1.4394   0.02047   0.01206  -0.0920   0.1619   1.0000
   9.000   1.4491   0.02185   0.01319  -0.0897   0.1276   1.0000
   9.250   1.4604   0.02304   0.01432  -0.0875   0.1045   1.0000
   9.750   1.4657   0.02624   0.01727  -0.0809   0.0472   1.0000
  10.000   1.4720   0.02747   0.01859  -0.0781   0.0354   1.0000
  10.500   1.4781   0.03057   0.02171  -0.0726   0.0200   1.0000
  10.750   1.4805   0.03231   0.02355  -0.0702   0.0182   1.0000
  11.000   1.4812   0.03433   0.02571  -0.0681   0.0168   1.0000
  11.250   1.4834   0.03635   0.02793  -0.0665   0.0160   1.0000
  11.500   1.4834   0.03871   0.03047  -0.0651   0.0151   1.0000
  11.750   1.4816   0.04141   0.03333  -0.0640   0.0144   1.0000
  12.000   1.4778   0.04450   0.03659  -0.0632   0.0137   1.0000
  12.250   1.4717   0.04807   0.04032  -0.0629   0.0132   1.0000
  12.500   1.4631   0.05217   0.04458  -0.0629   0.0129   1.0000
  12.750   1.4527   0.05672   0.04929  -0.0634   0.0126   1.0000
  13.000   1.4402   0.06164   0.05437  -0.0641   0.0124   1.0000
  13.250   1.4266   0.06681   0.05970  -0.0649   0.0122   1.0000
  13.500   1.4144   0.07188   0.06490  -0.0657   0.0120   1.0000
  13.750   1.4051   0.07663   0.06979  -0.0666   0.0119   1.0000
  14.000   1.3961   0.08146   0.07476  -0.0676   0.0117   1.0000
  14.250   1.3878   0.08625   0.07968  -0.0686   0.0116   1.0000
  14.500   1.3801   0.09095   0.08450  -0.0697   0.0114   1.0000
<< Back to GOE 174 (ALBATROS 5020) AIRFOIL (goe174-il)

Polar data table (+)

Polar graphs


<< Back to GOE 174 (ALBATROS 5020) AIRFOIL (goe174-il)