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GOE 411 AIRFOIL (goe411-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 411 AIRFOIL (goe411-il)
Reynolds number: 200,000
Max Cl/Cd: 43.75 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe411-il-200000.txt
Download as CSV file: xf-goe411-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 411 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.8809   0.03223   0.02499   0.0012   1.0000   0.0357
  -7.750  -0.8562   0.02932   0.02165   0.0030   1.0000   0.0300
  -7.500  -0.8398   0.02859   0.02069   0.0057   1.0000   0.0284
  -7.250  -0.8231   0.02786   0.01985   0.0079   1.0000   0.0278
  -7.000  -0.8024   0.02619   0.01810   0.0093   1.0000   0.0278
  -6.750  -0.7828   0.02384   0.01575   0.0106   1.0000   0.0291
  -6.500  -0.7701   0.02274   0.01463   0.0134   1.0000   0.0292
  -6.250  -0.7602   0.02173   0.01358   0.0167   1.0000   0.0292
  -6.000  -0.7520   0.02082   0.01260   0.0202   1.0000   0.0294
  -5.750  -0.7443   0.01998   0.01173   0.0238   1.0000   0.0297
  -5.500  -0.7363   0.01919   0.01086   0.0273   1.0000   0.0304
  -5.250  -0.7256   0.01856   0.01013   0.0304   1.0000   0.0312
  -5.000  -0.7135   0.01798   0.00942   0.0332   1.0000   0.0327
  -4.750  -0.6999   0.01748   0.00880   0.0357   1.0000   0.0356
  -4.500  -0.6849   0.01709   0.00827   0.0380   1.0000   0.0384
  -4.000  -0.6340   0.01493   0.00717   0.0373   0.9913   0.2333
  -3.750  -0.6086   0.01390   0.00707   0.0368   0.9847   0.3993
  -3.500  -0.5793   0.01359   0.00708   0.0362   0.9778   0.4816
  -3.250  -0.5450   0.01344   0.00711   0.0346   0.9715   0.5410
  -3.000  -0.5152   0.01326   0.00705   0.0341   0.9642   0.5858
  -2.750  -0.4804   0.01306   0.00702   0.0327   0.9579   0.6308
  -2.500  -0.4482   0.01285   0.00695   0.0318   0.9506   0.6628
  -2.250  -0.4098   0.01268   0.00682   0.0297   0.9445   0.6890
  -2.000  -0.3747   0.01251   0.00673   0.0283   0.9376   0.7162
  -1.750  -0.3361   0.01233   0.00662   0.0262   0.9311   0.7442
  -1.500  -0.2958   0.01216   0.00650   0.0238   0.9254   0.7717
  -1.250  -0.2574   0.01200   0.00641   0.0219   0.9180   0.7983
  -1.000  -0.2071   0.01186   0.00635   0.0177   0.9148   0.8270
  -0.750  -0.1635   0.01188   0.00644   0.0151   0.9075   0.8533
  -0.500  -0.1106   0.01187   0.00645   0.0104   0.9033   0.8724
  -0.250  -0.0492   0.01188   0.00645   0.0039   0.9006   0.8827
   0.000   0.0000   0.01192   0.00650   0.0000   0.8912   0.8912
   0.250   0.0490   0.01188   0.00645  -0.0039   0.8825   0.9006
   0.500   0.1105   0.01187   0.00645  -0.0104   0.8725   0.9033
   0.750   0.1635   0.01188   0.00644  -0.0151   0.8531   0.9074
   1.000   0.2072   0.01186   0.00635  -0.0178   0.8275   0.9148
   1.250   0.2575   0.01200   0.00641  -0.0219   0.7981   0.9180
   1.500   0.2959   0.01216   0.00650  -0.0238   0.7716   0.9254
   1.750   0.3364   0.01233   0.00662  -0.0262   0.7440   0.9311
   2.000   0.3747   0.01251   0.00673  -0.0282   0.7157   0.9376
   2.250   0.4099   0.01268   0.00682  -0.0297   0.6888   0.9445
   2.500   0.4484   0.01286   0.00695  -0.0318   0.6625   0.9507
   2.750   0.4802   0.01306   0.00702  -0.0326   0.6307   0.9579
   3.000   0.5149   0.01326   0.00704  -0.0340   0.5845   0.9641
   3.250   0.5454   0.01344   0.00711  -0.0347   0.5413   0.9715
   3.500   0.5795   0.01358   0.00708  -0.0362   0.4832   0.9777
   3.750   0.6085   0.01391   0.00707  -0.0368   0.3972   0.9847
   4.000   0.6340   0.01495   0.00718  -0.0373   0.2311   0.9914
   4.250   0.6577   0.01661   0.00778  -0.0378   0.0472   0.9973
   4.750   0.7000   0.01748   0.00880  -0.0357   0.0358   1.0000
   5.000   0.7134   0.01800   0.00944  -0.0331   0.0324   1.0000
   5.250   0.7260   0.01854   0.01010  -0.0304   0.0314   1.0000
   5.500   0.7354   0.01927   0.01094  -0.0271   0.0301   1.0000
   5.750   0.7437   0.02004   0.01178  -0.0237   0.0295   1.0000
   6.000   0.7506   0.02095   0.01273  -0.0199   0.0291   1.0000
   6.250   0.7600   0.02176   0.01361  -0.0166   0.0292   1.0000
   6.500   0.7700   0.02275   0.01463  -0.0134   0.0291   1.0000
   6.750   0.7827   0.02404   0.01594  -0.0106   0.0287   1.0000
   7.000   0.8004   0.02565   0.01756  -0.0089   0.0283   1.0000
   7.250   0.8226   0.02771   0.01971  -0.0078   0.0279   1.0000
   7.500   0.8391   0.02840   0.02050  -0.0055   0.0286   1.0000
   7.750   0.8566   0.02934   0.02168  -0.0030   0.0301   1.0000
   8.000   0.8805   0.03209   0.02484  -0.0012   0.0354   1.0000
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