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

GOE 173 (ALBATROS 6020) AIRFOIL (goe173-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 173 (ALBATROS 6020) AIRFOIL (goe173-il)
Reynolds number: 100,000
Max Cl/Cd: 56.89 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe173-il-100000.txt
Download as CSV file: xf-goe173-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 173 (ALBATROS 6020) AIRFOIL                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.2781   0.10277   0.09826  -0.0270   1.0000   0.0464
  -7.500  -0.2841   0.10232   0.09794  -0.0284   1.0000   0.0468
  -7.250  -0.2843   0.10155   0.09726  -0.0320   1.0000   0.0470
  -7.000  -0.2817   0.10037   0.09616  -0.0354   1.0000   0.0472
  -6.750  -0.2777   0.09196   0.08782  -0.0249   1.0000   0.0495
  -6.500  -0.2786   0.08971   0.08565  -0.0232   1.0000   0.0509
  -6.250  -0.2844   0.08814   0.08419  -0.0216   1.0000   0.0519
  -6.000  -0.2958   0.08717   0.08334  -0.0193   1.0000   0.0526
  -5.750  -0.3075   0.08630   0.08256  -0.0174   1.0000   0.0533
  -5.500  -0.3113   0.08482   0.08114  -0.0173   0.9995   0.0545
  -5.250  -0.2294   0.08121   0.07721  -0.0444   0.9878   0.0599
  -5.000  -0.2128   0.07447   0.07057  -0.0446   0.9829   0.0615
  -4.750  -0.1889   0.07024   0.06632  -0.0464   0.9751   0.0650
  -4.500  -0.1069   0.06718   0.06280  -0.0680   0.9645   0.0740
  -4.250  -0.0956   0.06105   0.05688  -0.0661   0.9592   0.0768
  -4.000  -0.0389   0.05767   0.05317  -0.0777   0.9495   0.0890
  -3.750  -0.0136   0.05327   0.04889  -0.0788   0.9443   0.0957
  -3.500   0.0294   0.04958   0.04501  -0.0855   0.9346   0.1068
  -3.250   0.0806   0.04566   0.04091  -0.0929   0.9298   0.1222
  -3.000   0.1311   0.04325   0.03811  -0.0998   0.9192   0.1476
  -2.750   0.1674   0.03925   0.03420  -0.1026   0.9117   0.1680
  -1.750   0.3103   0.02861   0.02330  -0.1101   0.8719   0.3819
  -1.500   0.4019   0.02534   0.01797  -0.1196   0.8615   0.1316
  -1.250   0.4372   0.02308   0.01533  -0.1200   0.8501   0.1141
  -1.000   0.4711   0.02151   0.01337  -0.1202   0.8387   0.1076
  -0.750   0.5041   0.02054   0.01194  -0.1198   0.8270   0.1019
  -0.500   0.5343   0.01952   0.01071  -0.1193   0.8149   0.1008
  -0.250   0.5620   0.01887   0.00988  -0.1185   0.8014   0.1017
   0.000   0.5886   0.01822   0.00922  -0.1178   0.7877   0.1073
   0.250   0.6148   0.01777   0.00872  -0.1168   0.7739   0.1110
   0.500   0.6405   0.01743   0.00829  -0.1156   0.7601   0.1161
   0.750   0.6668   0.01713   0.00794  -0.1147   0.7459   0.1263
   1.000   0.6937   0.01683   0.00766  -0.1139   0.7315   0.1599
   1.250   0.7168   0.01504   0.00738  -0.1121   0.7173   1.0000
   1.500   0.7426   0.01518   0.00725  -0.1111   0.7016   1.0000
   1.750   0.7684   0.01531   0.00715  -0.1101   0.6852   1.0000
   2.000   0.7942   0.01541   0.00705  -0.1092   0.6682   1.0000
   2.250   0.8186   0.01555   0.00706  -0.1081   0.6477   1.0000
   2.500   0.8437   0.01567   0.00702  -0.1071   0.6272   1.0000
   2.750   0.8683   0.01582   0.00703  -0.1061   0.6046   1.0000
   3.000   0.8932   0.01600   0.00701  -0.1051   0.5817   1.0000
   3.250   0.9173   0.01628   0.00713  -0.1041   0.5564   1.0000
   3.500   0.9419   0.01661   0.00730  -0.1032   0.5336   1.0000
   3.750   0.9665   0.01700   0.00753  -0.1024   0.5125   1.0000
   4.000   0.9916   0.01743   0.00779  -0.1018   0.4952   1.0000
   4.250   1.0169   0.01790   0.00814  -0.1013   0.4801   1.0000
   4.500   1.0421   0.01840   0.00858  -0.1009   0.4668   1.0000
   4.750   1.0675   0.01893   0.00912  -0.1005   0.4551   1.0000
   5.000   1.0931   0.01949   0.00965  -0.1002   0.4454   1.0000
   5.250   1.1194   0.02003   0.01011  -0.0999   0.4370   1.0000
   5.500   1.1442   0.02063   0.01080  -0.0995   0.4281   1.0000
   5.750   1.1711   0.02124   0.01136  -0.0994   0.4217   1.0000
   6.000   1.1954   0.02192   0.01223  -0.0990   0.4148   1.0000
   6.250   1.2218   0.02258   0.01291  -0.0989   0.4095   1.0000
   6.500   1.2469   0.02337   0.01383  -0.0986   0.4044   1.0000
   6.750   1.2715   0.02414   0.01482  -0.0983   0.3991   1.0000
   7.000   1.2980   0.02486   0.01557  -0.0982   0.3944   1.0000
   7.250   1.3199   0.02555   0.01647  -0.0974   0.3860   1.0000
   7.500   1.3434   0.02595   0.01689  -0.0967   0.3753   1.0000
   7.750   1.3651   0.02617   0.01715  -0.0956   0.3613   1.0000
   8.000   1.3862   0.02642   0.01743  -0.0945   0.3474   1.0000
   8.250   1.4090   0.02681   0.01786  -0.0937   0.3364   1.0000
   8.500   1.4277   0.02711   0.01833  -0.0923   0.3233   1.0000
   8.750   1.4435   0.02717   0.01847  -0.0903   0.3050   1.0000
   9.000   1.4604   0.02741   0.01888  -0.0885   0.2911   1.0000
   9.250   1.4704   0.02745   0.01923  -0.0856   0.2708   1.0000
   9.500   1.4759   0.02747   0.01942  -0.0822   0.2401   1.0000
   9.750   1.4799   0.02812   0.02007  -0.0788   0.1747   1.0000
  10.000   1.4608   0.03138   0.02241  -0.0737   0.0847   1.0000
  10.250   1.4489   0.03434   0.02516  -0.0694   0.0600   1.0000
  10.500   1.4421   0.03692   0.02782  -0.0661   0.0524   1.0000
  10.750   1.4348   0.03967   0.03071  -0.0635   0.0485   1.0000
  11.000   1.4238   0.04301   0.03419  -0.0615   0.0463   1.0000
  11.250   1.4153   0.04643   0.03785  -0.0603   0.0448   1.0000
  11.500   1.4064   0.05020   0.04184  -0.0597   0.0436   1.0000
  11.750   1.3960   0.05437   0.04622  -0.0595   0.0427   1.0000
  12.000   1.3851   0.05883   0.05086  -0.0598   0.0419   1.0000
  12.250   1.3741   0.06343   0.05563  -0.0602   0.0411   1.0000
  12.500   1.3639   0.06800   0.06034  -0.0606   0.0404   1.0000
  12.750   1.3552   0.07237   0.06483  -0.0610   0.0394   1.0000
  13.000   1.3482   0.07636   0.06889  -0.0611   0.0383   1.0000
  13.250   1.3468   0.07898   0.07146  -0.0596   0.0368   1.0000
  13.500   1.3558   0.08025   0.07275  -0.0572   0.0354   1.0000
  13.750   1.3714   0.08092   0.07350  -0.0542   0.0346   1.0000
  14.000   1.3884   0.08191   0.07461  -0.0512   0.0341   1.0000
  14.250   1.4021   0.08390   0.07681  -0.0489   0.0339   1.0000
  14.500   1.4096   0.08696   0.08010  -0.0476   0.0339   1.0000
  14.750   1.4096   0.09100   0.08439  -0.0472   0.0341   1.0000
  15.000   1.4039   0.09570   0.08934  -0.0477   0.0343   1.0000
  15.250   1.3946   0.10091   0.09480  -0.0488   0.0346   1.0000
  15.500   1.3823   0.10659   0.10073  -0.0507   0.0348   1.0000
  15.750   1.3683   0.11269   0.10706  -0.0532   0.0350   1.0000
  16.000   1.3533   0.11917   0.11376  -0.0562   0.0353   1.0000
  16.250   1.3377   0.12608   0.12087  -0.0599   0.0355   1.0000
  16.500   1.3217   0.13339   0.12837  -0.0641   0.0358   1.0000
  16.750   1.3067   0.14097   0.13610  -0.0685   0.0361   1.0000
  17.000   1.2915   0.14881   0.14414  -0.0738   0.0364   1.0000
  17.250   1.2680   0.15938   0.15495  -0.0826   0.0370   1.0000
<< Back to GOE 173 (ALBATROS 6020) AIRFOIL (goe173-il)

Polar data table (+)

Polar graphs


<< Back to GOE 173 (ALBATROS 6020) AIRFOIL (goe173-il)