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RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAE 5215 AIRFOIL (rae5215-il)
Reynolds number: 50,000
Max Cl/Cd: 28.5 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae5215-il-50000.txt
Download as CSV file: xf-rae5215-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 5215 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5366   0.12839   0.12091   0.0058   1.0000   0.2553
  -9.750  -0.5177   0.12342   0.11594   0.0066   1.0000   0.2647
  -9.500  -0.5294   0.12145   0.11407   0.0051   1.0000   0.2757
  -9.250  -0.5213   0.11839   0.11103   0.0055   1.0000   0.2901
  -9.000  -0.5013   0.11367   0.10632   0.0064   1.0000   0.3019
  -8.750  -0.4959   0.11038   0.10308   0.0066   1.0000   0.3155
  -8.500  -0.4902   0.10728   0.10003   0.0070   1.0000   0.3318
  -8.250  -0.4814   0.10414   0.09691   0.0077   1.0000   0.3501
  -8.000  -0.4725   0.10104   0.09386   0.0085   1.0000   0.3691
  -7.750  -0.4645   0.09799   0.09085   0.0092   1.0000   0.3877
  -7.500  -0.4565   0.09486   0.08777   0.0098   1.0000   0.4049
  -7.250  -0.4487   0.09171   0.08467   0.0104   1.0000   0.4213
  -7.000  -0.4405   0.08863   0.08164   0.0111   1.0000   0.4386
  -6.750  -0.4333   0.08554   0.07860   0.0117   1.0000   0.4542
  -6.500  -0.5574   0.05726   0.04972  -0.0404   1.0000   0.1678
  -6.250  -0.5447   0.05123   0.04317  -0.0431   1.0000   0.1504
  -6.000  -0.5274   0.04635   0.03730  -0.0451   1.0000   0.1376
  -5.750  -0.5078   0.04254   0.03336  -0.0448   1.0000   0.1348
  -5.500  -0.4863   0.03914   0.02958  -0.0448   1.0000   0.1317
  -5.250  -0.4636   0.03636   0.02638  -0.0445   1.0000   0.1323
  -5.000  -0.4397   0.03389   0.02351  -0.0440   1.0000   0.1335
  -4.750  -0.4150   0.03165   0.02090  -0.0432   1.0000   0.1345
  -4.500  -0.3897   0.02997   0.01876  -0.0424   1.0000   0.1382
  -4.250  -0.3667   0.02793   0.01686  -0.0413   1.0000   0.1443
  -4.000  -0.3424   0.02648   0.01530  -0.0399   1.0000   0.1505
  -3.750  -0.3195   0.02508   0.01397  -0.0383   1.0000   0.1608
  -3.500  -0.2967   0.02382   0.01279  -0.0367   1.0000   0.1750
  -3.250  -0.2740   0.02248   0.01163  -0.0355   1.0000   0.2005
  -3.000  -0.2478   0.01911   0.01018  -0.0363   1.0000   0.3940
  -2.750  -0.2596   0.02058   0.01266  -0.0220   1.0000   0.7135
  -2.500  -0.2635   0.02106   0.01316  -0.0117   1.0000   0.7781
  -2.250  -0.2658   0.02102   0.01311  -0.0025   1.0000   0.8278
  -2.000  -0.0909   0.02130   0.01258  -0.0170   1.0000   0.9848
  -1.750  -0.0326   0.02069   0.01171  -0.0245   1.0000   1.0000
  -1.500  -0.0321   0.02041   0.01140  -0.0215   1.0000   1.0000
  -1.250  -0.0337   0.02014   0.01112  -0.0182   1.0000   1.0000
  -1.000  -0.0367   0.01987   0.01085  -0.0147   1.0000   1.0000
  -0.750  -0.0405   0.01959   0.01056  -0.0111   1.0000   1.0000
  -0.500  -0.0441   0.01929   0.01025  -0.0076   1.0000   1.0000
  -0.250  -0.0426   0.01903   0.00996  -0.0050   1.0000   1.0000
   0.000  -0.0285   0.01891   0.00979  -0.0045   1.0000   1.0000
   0.250  -0.0063   0.01894   0.00976  -0.0054   1.0000   1.0000
   0.500   0.0200   0.01908   0.00983  -0.0069   1.0000   1.0000
   0.750   0.0483   0.01929   0.00999  -0.0087   1.0000   1.0000
   1.000   0.0773   0.01957   0.01024  -0.0105   1.0000   1.0000
   1.250   0.1064   0.01992   0.01057  -0.0123   1.0000   1.0000
   1.500   0.1352   0.02032   0.01097  -0.0140   1.0000   1.0000
   1.750   0.1634   0.02077   0.01144  -0.0156   1.0000   1.0000
   2.000   0.1909   0.02128   0.01199  -0.0170   1.0000   1.0000
   2.250   0.2176   0.02185   0.01262  -0.0184   1.0000   1.0000
   2.500   0.2435   0.02249   0.01334  -0.0196   1.0000   1.0000
   2.750   0.3125   0.02336   0.01445  -0.0284   0.9770   1.0000
   3.000   0.4343   0.02259   0.01427  -0.0423   0.9053   1.0000
   3.250   0.5195   0.01823   0.00947  -0.0369   0.5455   1.0000
   3.500   0.5352   0.02069   0.01015  -0.0338   0.3846   1.0000
   3.750   0.5643   0.02226   0.01126  -0.0339   0.3380   1.0000
   4.000   0.5966   0.02368   0.01234  -0.0345   0.3083   1.0000
   4.250   0.6285   0.02501   0.01360  -0.0350   0.2856   1.0000
   4.500   0.6596   0.02647   0.01500  -0.0354   0.2675   1.0000
   4.750   0.6894   0.02807   0.01654  -0.0356   0.2519   1.0000
   5.000   0.7184   0.02984   0.01822  -0.0358   0.2381   1.0000
   5.250   0.7458   0.03153   0.02023  -0.0357   0.2261   1.0000
   5.500   0.7725   0.03367   0.02253  -0.0355   0.2160   1.0000
   5.750   0.7985   0.03573   0.02471  -0.0353   0.2058   1.0000
   6.000   0.8217   0.03836   0.02772  -0.0349   0.1981   1.0000
   6.250   0.8444   0.04079   0.03044  -0.0344   0.1898   1.0000
   6.500   0.8658   0.04390   0.03371  -0.0340   0.1838   1.0000
   6.750   0.8817   0.04734   0.03779  -0.0331   0.1797   1.0000
   7.000   0.8963   0.05108   0.04197  -0.0324   0.1766   1.0000
   7.250   0.9084   0.05521   0.04650  -0.0318   0.1749   1.0000
   7.500   0.9128   0.06063   0.05242  -0.0316   0.1780   1.0000
   7.750   0.9166   0.06613   0.05826  -0.0316   0.1813   1.0000
   8.000   0.9245   0.07148   0.06376  -0.0318   0.1842   1.0000
   8.250   0.7202   0.10748   0.10037  -0.0765   0.4261   1.0000
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