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RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RAE 101 AIRFOIL (rae101-il)
Reynolds number: 200,000
Max Cl/Cd: 49.65 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae101-il-200000.txt
Download as CSV file: xf-rae101-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 101 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.8410   0.05365   0.04905  -0.0242   1.0000   0.0364
  -9.250  -0.8464   0.04864   0.04383  -0.0228   1.0000   0.0354
  -9.000  -0.8474   0.04473   0.03957  -0.0209   1.0000   0.0354
  -8.750  -0.8444   0.04081   0.03525  -0.0189   1.0000   0.0351
  -8.500  -0.8381   0.03685   0.03088  -0.0169   1.0000   0.0347
  -8.250  -0.8274   0.03330   0.02691  -0.0149   1.0000   0.0343
  -8.000  -0.8128   0.03034   0.02355  -0.0131   1.0000   0.0345
  -7.750  -0.7950   0.02796   0.02080  -0.0115   1.0000   0.0353
  -7.500  -0.7748   0.02656   0.01908  -0.0102   1.0000   0.0367
  -7.250  -0.7555   0.02376   0.01607  -0.0090   1.0000   0.0388
  -7.000  -0.7341   0.02221   0.01448  -0.0080   1.0000   0.0409
  -6.750  -0.7121   0.02094   0.01309  -0.0069   1.0000   0.0433
  -6.500  -0.6900   0.01989   0.01192  -0.0058   1.0000   0.0464
  -6.250  -0.6701   0.01842   0.01043  -0.0044   1.0000   0.0504
  -6.000  -0.6498   0.01748   0.00950  -0.0031   1.0000   0.0551
  -5.750  -0.6293   0.01663   0.00857  -0.0016   1.0000   0.0604
  -5.500  -0.6109   0.01564   0.00763   0.0000   1.0000   0.0694
  -5.250  -0.5924   0.01470   0.00674   0.0016   1.0000   0.0805
  -5.000  -0.5737   0.01383   0.00595   0.0032   1.0000   0.1014
  -4.750  -0.5572   0.01269   0.00516   0.0050   1.0000   0.1546
  -4.500  -0.5448   0.01124   0.00455   0.0071   1.0000   0.3009
  -4.250  -0.5275   0.01062   0.00434   0.0089   1.0000   0.4089
  -4.000  -0.5077   0.01032   0.00420   0.0105   1.0000   0.4644
  -3.750  -0.4871   0.01011   0.00409   0.0119   1.0000   0.5047
  -3.500  -0.4665   0.00995   0.00400   0.0134   1.0000   0.5377
  -3.250  -0.4460   0.00982   0.00392   0.0149   1.0000   0.5672
  -3.000  -0.4258   0.00972   0.00387   0.0163   1.0000   0.5940
  -2.750  -0.4060   0.00964   0.00385   0.0179   1.0000   0.6200
  -2.500  -0.3867   0.00960   0.00387   0.0194   1.0000   0.6456
  -2.250  -0.3637   0.00957   0.00392   0.0202   0.9985   0.6717
  -2.000  -0.3200   0.00953   0.00396   0.0169   0.9890   0.7030
  -1.750  -0.2773   0.00945   0.00398   0.0139   0.9796   0.7324
  -1.500  -0.2339   0.00936   0.00396   0.0109   0.9708   0.7590
  -1.250  -0.1934   0.00923   0.00391   0.0085   0.9600   0.7836
  -1.000  -0.1514   0.00911   0.00385   0.0059   0.9502   0.8078
  -0.750  -0.1092   0.00898   0.00380   0.0034   0.9402   0.8290
  -0.500  -0.0712   0.00888   0.00374   0.0018   0.9266   0.8516
  -0.250  -0.0347   0.00882   0.00372   0.0007   0.9105   0.8717
   0.000   0.0000   0.00879   0.00371   0.0000   0.8914   0.8914
   0.250   0.0347   0.00882   0.00372  -0.0007   0.8717   0.9105
   0.500   0.0711   0.00888   0.00374  -0.0018   0.8515   0.9265
   0.750   0.1092   0.00898   0.00380  -0.0034   0.8291   0.9402
   1.000   0.1513   0.00911   0.00385  -0.0059   0.8078   0.9502
   1.250   0.1934   0.00923   0.00391  -0.0085   0.7835   0.9600
   1.500   0.2339   0.00935   0.00396  -0.0109   0.7592   0.9708
   1.750   0.2774   0.00945   0.00398  -0.0139   0.7325   0.9796
   2.000   0.3200   0.00952   0.00396  -0.0169   0.7032   0.9891
   2.250   0.3636   0.00956   0.00392  -0.0202   0.6717   0.9986
   2.500   0.3865   0.00960   0.00387  -0.0194   0.6457   1.0000
   2.750   0.4058   0.00964   0.00385  -0.0178   0.6199   1.0000
   3.000   0.4256   0.00971   0.00387  -0.0163   0.5941   1.0000
   3.250   0.4458   0.00982   0.00392  -0.0148   0.5674   1.0000
   3.500   0.4663   0.00995   0.00400  -0.0134   0.5380   1.0000
   3.750   0.4869   0.01011   0.00409  -0.0119   0.5045   1.0000
   4.000   0.5075   0.01031   0.00420  -0.0104   0.4649   1.0000
   4.250   0.5273   0.01062   0.00434  -0.0089   0.4090   1.0000
   4.500   0.5446   0.01123   0.00455  -0.0071   0.3012   1.0000
   4.750   0.5571   0.01269   0.00516  -0.0050   0.1543   1.0000
   5.000   0.5736   0.01383   0.00595  -0.0032   0.1015   1.0000
   5.250   0.5923   0.01471   0.00674  -0.0016   0.0807   1.0000
   5.500   0.6108   0.01564   0.00763   0.0001   0.0693   1.0000
   5.750   0.6292   0.01663   0.00856   0.0016   0.0604   1.0000
   6.000   0.6498   0.01748   0.00950   0.0031   0.0551   1.0000
   6.250   0.6701   0.01843   0.01043   0.0044   0.0504   1.0000
   6.500   0.6900   0.01989   0.01192   0.0058   0.0464   1.0000
   6.750   0.7121   0.02094   0.01309   0.0069   0.0433   1.0000
   7.000   0.7341   0.02223   0.01449   0.0080   0.0410   1.0000
   7.250   0.7556   0.02375   0.01606   0.0090   0.0389   1.0000
   7.500   0.7748   0.02665   0.01917   0.0101   0.0367   1.0000
   7.750   0.7951   0.02796   0.02080   0.0115   0.0353   1.0000
   8.000   0.8129   0.03035   0.02355   0.0131   0.0346   1.0000
   8.250   0.8276   0.03328   0.02688   0.0149   0.0344   1.0000
   8.500   0.8383   0.03685   0.03088   0.0168   0.0347   1.0000
   8.750   0.8447   0.04080   0.03524   0.0188   0.0352   1.0000
   9.000   0.8477   0.04475   0.03958   0.0208   0.0355   1.0000
   9.250   0.8467   0.04895   0.04410   0.0226   0.0358   1.0000
   9.500   0.8416   0.05332   0.04874   0.0242   0.0362   1.0000
  12.250   0.5664   0.12453   0.12116  -0.0012   0.0597   1.0000
  12.500   0.5552   0.12989   0.12649  -0.0053   0.0594   1.0000
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