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BOEING AIRFOIL J (no closed TE) (bacj-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: BOEING AIRFOIL J (no closed TE) (bacj-il)
Reynolds number: 200,000
Max Cl/Cd: -1 at α=-100°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-bacj-il-200000-n5.txt
Download as CSV file: xf-bacj-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING AIRFOIL J                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.7503   0.07341   0.06901  -0.0578   1.0000   0.0226
 -12.750  -0.7715   0.06753   0.06312  -0.0607   1.0000   0.0232
 -12.500  -0.8031   0.06095   0.05633  -0.0641   1.0000   0.0232
 -12.250  -0.8349   0.05540   0.05056  -0.0655   1.0000   0.0230
 -12.000  -0.8575   0.05167   0.04657  -0.0651   1.0000   0.0226
 -11.750  -0.8770   0.04851   0.04326  -0.0638   1.0000   0.0228
 -11.500  -0.8795   0.04723   0.04205  -0.0622   1.0000   0.0234
 -11.250  -0.9040   0.04415   0.03868  -0.0595   1.0000   0.0231
 -11.000  -0.9089   0.04309   0.03762  -0.0571   1.0000   0.0236
 -10.750  -0.9232   0.04135   0.03574  -0.0537   0.9999   0.0238
 -10.500  -0.9231   0.03901   0.03310  -0.0530   0.9963   0.0236
 -10.250  -0.9171   0.03798   0.03198  -0.0515   0.9931   0.0243
 -10.000  -0.9149   0.03600   0.02972  -0.0494   0.9894   0.0241
  -9.750  -0.9028   0.03510   0.02871  -0.0483   0.9865   0.0248
  -9.500  -0.8912   0.03311   0.02646  -0.0474   0.9842   0.0247
  -9.250  -0.8764   0.03216   0.02534  -0.0464   0.9819   0.0261
  -9.000  -0.8651   0.03096   0.02397  -0.0445   0.9784   0.0264
  -8.750  -0.8507   0.02940   0.02220  -0.0431   0.9757   0.0264
  -8.500  -0.8336   0.02863   0.02122  -0.0419   0.9732   0.0283
  -8.250  -0.8129   0.02748   0.01991  -0.0415   0.9713   0.0285
  -8.000  -0.7895   0.02659   0.01885  -0.0413   0.9696   0.0292
  -7.750  -0.7733   0.02567   0.01782  -0.0397   0.9669   0.0290
  -7.500  -0.7619   0.02434   0.01647  -0.0374   0.9638   0.0303
  -7.250  -0.7441   0.02357   0.01564  -0.0360   0.9610   0.0312
  -7.000  -0.7231   0.02282   0.01484  -0.0353   0.9584   0.0324
  -6.750  -0.7010   0.02216   0.01410  -0.0347   0.9565   0.0341
  -6.500  -0.6772   0.02160   0.01346  -0.0344   0.9550   0.0348
  -6.250  -0.6624   0.02102   0.01281  -0.0322   0.9519   0.0349
  -6.000  -0.6468   0.02063   0.01234  -0.0300   0.9481   0.0369
  -5.750  -0.6307   0.01984   0.01157  -0.0282   0.9450   0.0406
  -5.500  -0.6087   0.01941   0.01107  -0.0274   0.9427   0.0438
  -5.250  -0.5836   0.01899   0.01057  -0.0270   0.9408   0.0457
  -5.000  -0.5582   0.01854   0.01010  -0.0268   0.9393   0.0494
  -4.750  -0.5500   0.01813   0.00971  -0.0231   0.9341   0.0546
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