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

GOE 770 AIRFOIL (goe770-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 770 AIRFOIL (goe770-il)
Reynolds number: 50,000
Max Cl/Cd: 2.92 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe770-il-50000.txt
Download as CSV file: xf-goe770-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 770 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.2480   0.12938   0.12346  -0.0237   0.9970   0.3397
 -10.750  -0.2622   0.12711   0.12121  -0.0289   0.9868   0.3480
 -10.500  -0.4929   0.08643   0.08043  -0.0585   0.9820   0.1839
 -10.250  -0.5089   0.07841   0.07228  -0.0649   0.9705   0.1814
 -10.000  -0.5498   0.07151   0.06519  -0.0680   0.9569   0.1785
  -9.750  -0.5885   0.06540   0.05872  -0.0688   0.9435   0.1762
  -9.500  -0.6020   0.06066   0.05357  -0.0695   0.9328   0.1763
  -9.250  -0.6126   0.05769   0.05026  -0.0677   0.9212   0.1773
  -9.000  -0.6037   0.05400   0.04607  -0.0688   0.9134   0.1798
  -8.750  -0.6119   0.05218   0.04390  -0.0654   0.9028   0.1812
  -8.500  -0.5907   0.04942   0.04074  -0.0665   0.8954   0.1852
  -8.250  -0.5704   0.04834   0.03970  -0.0661   0.8869   0.1906
  -8.000  -0.5572   0.04688   0.03794  -0.0650   0.8786   0.1960
  -7.750  -0.5168   0.04504   0.03598  -0.0677   0.8729   0.2051
  -7.500  -0.5215   0.04484   0.03569  -0.0635   0.8641   0.2097
  -7.250  -0.4972   0.04372   0.03444  -0.0636   0.8569   0.2192
  -7.000  -0.4564   0.04250   0.03311  -0.0659   0.8513   0.2349
  -6.750  -0.4663   0.04280   0.03350  -0.0608   0.8433   0.2403
  -6.500  -0.4460   0.04222   0.03292  -0.0601   0.8367   0.2558
  -6.250  -0.4050   0.04114   0.03187  -0.0622   0.8311   0.2844
  -6.000  -0.4149   0.04135   0.03221  -0.0572   0.8242   0.2964
  -5.750  -0.4071   0.04091   0.03192  -0.0547   0.8180   0.3250
  -5.500  -0.3807   0.04021   0.03201  -0.0538   0.8123   0.3905
  -5.250  -0.3728   0.04124   0.03351  -0.0495   0.8066   0.4555
  -5.000  -0.3819   0.04240   0.03481  -0.0437   0.8005   0.4914
  -4.750  -0.3692   0.04378   0.03627  -0.0401   0.7945   0.5350
  -4.500  -0.3243   0.04532   0.03777  -0.0396   0.7881   0.5831
  -4.250  -0.3413   0.04628   0.03878  -0.0335   0.7827   0.5976
  -4.000  -0.3375   0.04726   0.03977  -0.0296   0.7773   0.6209
  -3.750  -0.3182   0.04815   0.04060  -0.0272   0.7712   0.6495
  -3.500  -0.2922   0.04930   0.04170  -0.0250   0.7648   0.6780
  -3.250  -0.3012   0.05017   0.04258  -0.0201   0.7604   0.6986
  -3.000  -0.2967   0.05126   0.04369  -0.0158   0.7556   0.7220
  -2.750  -0.2720   0.05262   0.04503  -0.0127   0.7490   0.7498
  -2.500  -0.2483   0.05395   0.04628  -0.0101   0.7428   0.7774
  -2.250  -0.2526   0.05483   0.04715  -0.0061   0.7397   0.7980
  -2.000  -0.2470   0.05577   0.04806  -0.0033   0.7361   0.8195
  -1.750  -0.2294   0.05681   0.04902  -0.0023   0.7319   0.8406
  -1.500  -0.1671   0.05822   0.05020  -0.0068   0.7217   0.8612
  -1.250  -0.1620   0.05930   0.05124  -0.0059   0.7203   0.8780
  -1.000  -0.1508   0.06050   0.05238  -0.0060   0.7192   0.8948
  -0.750  -0.1255   0.06221   0.05401  -0.0087   0.7192   0.9103
  -0.500  -0.0199   0.06442   0.05595  -0.0214   0.7025   0.9227
  -0.250   0.0116   0.06650   0.05796  -0.0258   0.7024   0.9361
   0.750   0.0001   0.07654   0.06823  -0.0369   0.8148   0.9924
   1.000   0.0325   0.07840   0.06999  -0.0415   0.8040   1.0000
   1.500   0.0232   0.07813   0.06952  -0.0366   0.7854   1.0000
   1.750   0.0040   0.07638   0.06768  -0.0316   0.7715   1.0000
   2.000   0.0159   0.07800   0.06909  -0.0315   0.7662   1.0000
   2.250   0.0043   0.07653   0.06749  -0.0278   0.7507   1.0000
   2.500   0.0464   0.08044   0.07118  -0.0319   0.7453   1.0000
   2.750   0.0331   0.07912   0.06975  -0.0282   0.7311   1.0000
   3.000   0.0682   0.08220   0.07265  -0.0310   0.7252   1.0000
   3.250   0.0644   0.08244   0.07280  -0.0288   0.7158   1.0000
   3.500   0.0885   0.08440   0.07462  -0.0299   0.7066   1.0000
   3.750   0.1315   0.08886   0.07892  -0.0335   0.7022   1.0000
   4.000   0.1138   0.08742   0.07743  -0.0295   0.6894   1.0000
   4.250   0.1451   0.09031   0.08021  -0.0314   0.6831   1.0000
   4.500   0.1736   0.09403   0.08381  -0.0334   0.6797   1.0000
   4.750   0.1601   0.09284   0.08260  -0.0300   0.6675   1.0000
   5.000   0.1907   0.09586   0.08552  -0.0318   0.6616   1.0000
   5.250   0.2108   0.09893   0.08850  -0.0327   0.6575   1.0000
   5.500   0.2051   0.09849   0.08803  -0.0305   0.6456   1.0000
   5.750   0.2333   0.10149   0.09097  -0.0320   0.6403   1.0000
   6.000   0.2743   0.10682   0.09622  -0.0351   0.6374   1.0000
   6.250   0.2456   0.10432   0.09371  -0.0307   0.6260   1.0000
   6.500   0.2716   0.10712   0.09647  -0.0319   0.6196   1.0000
   6.750   0.3111   0.11229   0.10158  -0.0347   0.6163   1.0000
   7.000   0.2859   0.11048   0.09977  -0.0312   0.6068   1.0000
   7.250   0.3068   0.11290   0.10216  -0.0319   0.5996   1.0000
   7.500   0.3434   0.11765   0.10688  -0.0342   0.5956   1.0000
   7.750   0.3260   0.11697   0.10620  -0.0318   0.5877   1.0000
   8.000   0.3423   0.11906   0.10828  -0.0322   0.5797   1.0000
<< Back to GOE 770 AIRFOIL (goe770-il)

Polar data table (+)

Polar graphs


<< Back to GOE 770 AIRFOIL (goe770-il)