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GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 50,000
Max Cl/Cd: 32.74 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe55-il-50000-n5.txt
Download as CSV file: xf-goe55-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6048   0.11140   0.10432   0.0327   1.0000   0.1158
  -8.000  -0.6098   0.10889   0.10189   0.0276   1.0000   0.1187
  -7.750  -0.6088   0.10563   0.09864   0.0177   1.0000   0.1196
  -7.500  -0.5920   0.10037   0.09341   0.0256   1.0000   0.1230
  -7.250  -0.5822   0.09682   0.08989   0.0242   1.0000   0.1286
  -7.000  -0.5760   0.09307   0.08611   0.0109   1.0000   0.1341
  -6.750  -0.5632   0.08875   0.08185   0.0163   1.0000   0.1372
  -6.500  -0.5498   0.08496   0.07806   0.0146   1.0000   0.1410
  -6.250  -0.5338   0.08062   0.07361   0.0044   1.0000   0.1492
  -5.750  -0.4815   0.06650   0.05898  -0.0082   1.0000   0.0912
  -5.500  -0.4604   0.06232   0.05470  -0.0106   1.0000   0.0899
  -5.250  -0.4368   0.05795   0.05017  -0.0136   1.0000   0.0872
  -5.000  -0.4007   0.05122   0.04272  -0.0209   1.0000   0.0797
  -4.750  -0.3753   0.04770   0.03904  -0.0225   1.0000   0.0789
  -4.500  -0.3473   0.04416   0.03523  -0.0245   1.0000   0.0782
  -4.250  -0.3162   0.04069   0.03128  -0.0268   1.0000   0.0794
  -4.000  -0.2858   0.03769   0.02784  -0.0285   1.0000   0.0805
  -3.750  -0.2561   0.03505   0.02487  -0.0296   1.0000   0.0807
  -3.500  -0.2263   0.03266   0.02216  -0.0304   1.0000   0.0809
  -3.250  -0.1969   0.03059   0.01984  -0.0310   1.0000   0.0815
  -3.000  -0.1676   0.02879   0.01783  -0.0314   1.0000   0.0826
  -2.750  -0.1387   0.02736   0.01627  -0.0316   1.0000   0.0857
  -2.500  -0.1091   0.02598   0.01464  -0.0318   1.0000   0.0894
  -2.250  -0.0794   0.02465   0.01301  -0.0317   1.0000   0.0917
  -2.000  -0.0508   0.02339   0.01165  -0.0315   1.0000   0.0937
  -1.750  -0.0231   0.02229   0.01059  -0.0312   1.0000   0.0968
  -1.500   0.0047   0.02141   0.00967  -0.0308   1.0000   0.1024
  -1.250   0.0327   0.02057   0.00886  -0.0305   1.0000   0.1097
  -1.000   0.0609   0.01980   0.00811  -0.0301   1.0000   0.1176
  -0.750   0.0890   0.01903   0.00750  -0.0300   1.0000   0.1282
  -0.500   0.1176   0.01836   0.00704  -0.0300   1.0000   0.1468
  -0.250   0.1592   0.01766   0.00667  -0.0327   0.8651   0.1889
   0.000   0.1832   0.01512   0.00619  -0.0300   0.6883   1.0000
   0.250   0.2081   0.01577   0.00597  -0.0282   0.6100   1.0000
   0.500   0.2346   0.01629   0.00592  -0.0274   0.5670   1.0000
   0.750   0.2620   0.01676   0.00597  -0.0269   0.5390   1.0000
   1.000   0.2898   0.01721   0.00609  -0.0265   0.5184   1.0000
   1.250   0.3180   0.01763   0.00627  -0.0263   0.5020   1.0000
   1.500   0.3464   0.01806   0.00652  -0.0261   0.4879   1.0000
   1.750   0.3748   0.01849   0.00680  -0.0260   0.4750   1.0000
   2.000   0.4032   0.01895   0.00713  -0.0258   0.4629   1.0000
   2.250   0.4313   0.01942   0.00750  -0.0256   0.4498   1.0000
   2.500   0.4594   0.01986   0.00787  -0.0254   0.4346   1.0000
   2.750   0.4872   0.02031   0.00826  -0.0252   0.4189   1.0000
   3.000   0.5148   0.02077   0.00870  -0.0249   0.4037   1.0000
   3.250   0.5425   0.02125   0.00918  -0.0247   0.3893   1.0000
   3.500   0.5700   0.02175   0.00971  -0.0245   0.3752   1.0000
   3.750   0.5975   0.02228   0.01030  -0.0243   0.3615   1.0000
   4.000   0.6250   0.02283   0.01091  -0.0240   0.3478   1.0000
   4.250   0.6523   0.02342   0.01154  -0.0238   0.3343   1.0000
   4.500   0.6794   0.02400   0.01221  -0.0235   0.3201   1.0000
   4.750   0.7064   0.02453   0.01286  -0.0232   0.3028   1.0000
   5.000   0.7332   0.02491   0.01339  -0.0229   0.2818   1.0000
   5.250   0.7593   0.02500   0.01352  -0.0225   0.2594   1.0000
   5.500   0.7857   0.02501   0.01370  -0.0222   0.2313   1.0000
   5.750   0.8118   0.02508   0.01375  -0.0220   0.2008   1.0000
   6.000   0.8375   0.02558   0.01414  -0.0218   0.1721   1.0000
   6.250   0.8625   0.02663   0.01513  -0.0217   0.1454   1.0000
   6.500   0.8868   0.02813   0.01661  -0.0215   0.1242   1.0000
   6.750   0.9105   0.02981   0.01831  -0.0213   0.1100   1.0000
   7.000   0.9336   0.03145   0.01999  -0.0209   0.1010   1.0000
   7.250   0.9570   0.03322   0.02202  -0.0205   0.0940   1.0000
   7.500   0.9793   0.03481   0.02367  -0.0202   0.0883   1.0000
   7.750   1.0013   0.03674   0.02583  -0.0198   0.0839   1.0000
   8.000   1.0225   0.03894   0.02835  -0.0194   0.0807   1.0000
   8.250   1.0428   0.04121   0.03091  -0.0191   0.0783   1.0000
   8.500   1.0621   0.04353   0.03342  -0.0187   0.0764   1.0000
   8.750   1.0804   0.04596   0.03598  -0.0183   0.0749   1.0000
   9.000   1.0934   0.04975   0.04031  -0.0184   0.0733   1.0000
   9.250   1.1008   0.05423   0.04536  -0.0189   0.0715   1.0000
   9.500   1.1023   0.05921   0.05082  -0.0199   0.0700   1.0000
   9.750   1.0952   0.06511   0.05715  -0.0217   0.0691   1.0000
  10.000   1.0699   0.07357   0.06599  -0.0264   0.0693   1.0000
  10.250   1.0154   0.08955   0.08216  -0.0403   0.0713   1.0000
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