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AG47ct -02f rot. (ag47ct02r-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG47ct -02f rot. (ag47ct02r-il)
Reynolds number: 100,000
Max Cl/Cd: 38.88 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag47ct02r-il-100000-n5.txt
Download as CSV file: xf-ag47ct02r-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG47ct  -02f rot.                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6492   0.13607   0.13105   0.0378   1.0000   0.0373
 -10.000  -0.6477   0.13293   0.12795   0.0346   1.0000   0.0377
  -9.750  -0.6452   0.12952   0.12458   0.0317   1.0000   0.0379
  -9.500  -0.6422   0.12593   0.12103   0.0290   1.0000   0.0381
  -9.250  -0.6388   0.12220   0.11733   0.0263   1.0000   0.0382
  -9.000  -0.6354   0.11836   0.11353   0.0236   1.0000   0.0382
  -8.750  -0.6317   0.11440   0.10961   0.0207   1.0000   0.0383
  -8.500  -0.6268   0.10831   0.10356   0.0221   1.0000   0.0390
  -8.250  -0.6197   0.10402   0.09924   0.0230   1.0000   0.0398
  -8.000  -0.6138   0.10018   0.09542   0.0222   1.0000   0.0404
  -7.750  -0.6087   0.09636   0.09164   0.0206   1.0000   0.0410
  -7.500  -0.6036   0.09245   0.08776   0.0181   1.0000   0.0415
  -7.250  -0.5956   0.08813   0.08347   0.0143   1.0000   0.0419
  -7.000  -0.5857   0.08352   0.07886   0.0097   1.0000   0.0421
  -6.750  -0.5735   0.07857   0.07390   0.0046   1.0000   0.0421
  -6.500  -0.5504   0.06962   0.06475  -0.0057   1.0000   0.0259
  -6.250  -0.5385   0.06516   0.06026  -0.0077   1.0000   0.0245
  -6.000  -0.5211   0.05994   0.05494  -0.0114   1.0000   0.0233
  -5.750  -0.5007   0.05431   0.04913  -0.0154   1.0000   0.0220
  -5.500  -0.4779   0.04842   0.04297  -0.0190   1.0000   0.0209
  -5.250  -0.4532   0.04236   0.03651  -0.0218   1.0000   0.0199
  -5.000  -0.4285   0.03784   0.03160  -0.0234   1.0000   0.0202
  -4.750  -0.4034   0.03445   0.02784  -0.0243   1.0000   0.0219
  -4.500  -0.3768   0.03085   0.02374  -0.0248   1.0000   0.0233
  -4.250  -0.3496   0.02738   0.01971  -0.0250   1.0000   0.0236
  -4.000  -0.3219   0.02441   0.01614  -0.0248   1.0000   0.0239
  -3.750  -0.2941   0.02197   0.01321  -0.0245   1.0000   0.0246
  -3.500  -0.2665   0.02000   0.01086  -0.0239   1.0000   0.0257
  -3.250  -0.2398   0.01825   0.00891  -0.0235   1.0000   0.0283
  -3.000  -0.2133   0.01717   0.00779  -0.0232   1.0000   0.0336
  -2.750  -0.1866   0.01597   0.00643  -0.0225   1.0000   0.0377
  -2.500  -0.1603   0.01499   0.00543  -0.0222   1.0000   0.0485
  -2.250  -0.1337   0.01408   0.00450  -0.0217   1.0000   0.0647
  -2.000  -0.1074   0.01317   0.00383  -0.0215   1.0000   0.1097
  -1.750  -0.0818   0.01207   0.00340  -0.0214   1.0000   0.2569
  -1.500  -0.0617   0.01040   0.00322  -0.0199   1.0000   0.6333
  -1.250  -0.0269   0.00951   0.00291  -0.0197   1.0000   1.0000
  -1.000  -0.0006   0.00950   0.00269  -0.0194   1.0000   1.0000
  -0.750   0.0257   0.00949   0.00253  -0.0190   1.0000   1.0000
  -0.500   0.0518   0.00950   0.00241  -0.0187   1.0000   1.0000
  -0.250   0.0778   0.00952   0.00232  -0.0184   1.0000   1.0000
   0.000   0.1037   0.00955   0.00229  -0.0180   1.0000   1.0000
   0.250   0.1294   0.00959   0.00230  -0.0177   1.0000   1.0000
   0.500   0.1597   0.00966   0.00235  -0.0185   0.9911   1.0000
   0.750   0.2046   0.00976   0.00241  -0.0220   0.9427   1.0000
   1.000   0.2463   0.00987   0.00245  -0.0244   0.8800   1.0000
   1.250   0.2793   0.01005   0.00246  -0.0247   0.8173   1.0000
   1.500   0.3056   0.01031   0.00253  -0.0236   0.7662   1.0000
   1.750   0.3309   0.01060   0.00263  -0.0225   0.7252   1.0000
   2.000   0.3564   0.01089   0.00278  -0.0216   0.6879   1.0000
   2.250   0.3819   0.01113   0.00294  -0.0208   0.6450   1.0000
   2.500   0.4072   0.01139   0.00307  -0.0199   0.5965   1.0000
   2.750   0.4324   0.01169   0.00323  -0.0190   0.5399   1.0000
   3.000   0.4575   0.01208   0.00339  -0.0182   0.4765   1.0000
   3.250   0.4825   0.01255   0.00360  -0.0176   0.4087   1.0000
   3.500   0.5076   0.01311   0.00389  -0.0171   0.3414   1.0000
   3.750   0.5328   0.01372   0.00425  -0.0167   0.2807   1.0000
   4.000   0.5581   0.01436   0.00474  -0.0164   0.2298   1.0000
   4.250   0.5836   0.01501   0.00525  -0.0161   0.1893   1.0000
   4.500   0.6091   0.01568   0.00581  -0.0158   0.1581   1.0000
   4.750   0.6346   0.01637   0.00646  -0.0155   0.1343   1.0000
   5.000   0.6598   0.01713   0.00718  -0.0152   0.1143   1.0000
   5.250   0.6850   0.01793   0.00802  -0.0149   0.1000   1.0000
   5.500   0.7101   0.01875   0.00888  -0.0145   0.0863   1.0000
   5.750   0.7347   0.01967   0.00984  -0.0141   0.0752   1.0000
   6.000   0.7597   0.02066   0.01101  -0.0136   0.0663   1.0000
   6.250   0.7842   0.02170   0.01221  -0.0132   0.0575   1.0000
   6.500   0.8082   0.02291   0.01351  -0.0128   0.0515   1.0000
   6.750   0.8320   0.02454   0.01537  -0.0121   0.0466   1.0000
   7.000   0.8559   0.02567   0.01667  -0.0118   0.0407   1.0000
   7.250   0.8780   0.02774   0.01891  -0.0113   0.0375   1.0000
   7.500   0.9005   0.03021   0.02187  -0.0105   0.0350   1.0000
   7.750   0.9217   0.03258   0.02468  -0.0100   0.0320   1.0000
   8.000   0.9422   0.03408   0.02640  -0.0098   0.0289   1.0000
   8.250   0.9581   0.03768   0.03040  -0.0094   0.0274   1.0000
   8.500   0.9708   0.04211   0.03554  -0.0091   0.0265   1.0000
   8.750   0.9769   0.04761   0.04174  -0.0093   0.0258   1.0000
   9.000   0.9754   0.05393   0.04864  -0.0104   0.0255   1.0000
   9.250   0.9663   0.06078   0.05593  -0.0128   0.0254   1.0000
   9.500   0.9504   0.06798   0.06341  -0.0171   0.0256   1.0000
   9.750   0.9309   0.07679   0.07234  -0.0259   0.0260   1.0000
  10.000   0.9133   0.08828   0.08381  -0.0356   0.0265   1.0000
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