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RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAE 102 AIRFOIL (rae102-il)
Reynolds number: 100,000
Max Cl/Cd: 39.63 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae102-il-100000.txt
Download as CSV file: xf-rae102-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 102 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5374   0.10150   0.09681  -0.0056   1.0000   0.1500
 -10.000  -0.5223   0.09753   0.09283  -0.0042   1.0000   0.1560
  -9.750  -0.5388   0.09282   0.08818  -0.0070   1.0000   0.1635
  -9.500  -0.5394   0.08781   0.08320  -0.0075   1.0000   0.1680
  -9.250  -0.5345   0.08377   0.07916  -0.0074   1.0000   0.1731
  -9.000  -0.5601   0.07779   0.07328  -0.0107   1.0000   0.1815
  -8.750  -0.5531   0.07358   0.06907  -0.0102   1.0000   0.1850
  -8.250  -0.7931   0.05135   0.04476  -0.0198   1.0000   0.0703
  -8.000  -0.7774   0.04679   0.04011  -0.0189   1.0000   0.0660
  -7.750  -0.7695   0.04241   0.03532  -0.0170   1.0000   0.0632
  -7.500  -0.7586   0.03904   0.03152  -0.0150   1.0000   0.0638
  -7.250  -0.7447   0.03606   0.02808  -0.0130   1.0000   0.0648
  -7.000  -0.7279   0.03309   0.02466  -0.0112   1.0000   0.0652
  -6.750  -0.7084   0.03038   0.02152  -0.0095   1.0000   0.0660
  -6.500  -0.6869   0.02825   0.01896  -0.0079   1.0000   0.0680
  -6.250  -0.6651   0.02594   0.01657  -0.0070   1.0000   0.0732
  -6.000  -0.6417   0.02428   0.01478  -0.0058   1.0000   0.0786
  -5.750  -0.6181   0.02246   0.01288  -0.0047   1.0000   0.0847
  -5.500  -0.5967   0.02131   0.01168  -0.0034   1.0000   0.0955
  -5.250  -0.5773   0.01976   0.01030  -0.0017   1.0000   0.1080
  -5.000  -0.5612   0.01839   0.00917   0.0003   1.0000   0.1289
  -4.750  -0.5480   0.01689   0.00798   0.0029   1.0000   0.1689
  -4.500  -0.5477   0.01425   0.00688   0.0070   1.0000   0.3696
  -4.250  -0.5402   0.01362   0.00702   0.0117   1.0000   0.5634
  -4.000  -0.5248   0.01359   0.00711   0.0150   1.0000   0.6272
  -3.750  -0.5083   0.01363   0.00719   0.0181   1.0000   0.6715
  -3.500  -0.4916   0.01367   0.00720   0.0211   1.0000   0.7069
  -3.250  -0.4752   0.01370   0.00724   0.0241   1.0000   0.7381
  -3.000  -0.4592   0.01375   0.00729   0.0272   1.0000   0.7663
  -2.750  -0.4432   0.01381   0.00738   0.0304   1.0000   0.7921
  -2.500  -0.4277   0.01388   0.00745   0.0335   1.0000   0.8171
  -2.250  -0.4123   0.01396   0.00750   0.0365   1.0000   0.8425
  -2.000  -0.3954   0.01406   0.00758   0.0390   1.0000   0.8678
  -1.750  -0.3691   0.01427   0.00775   0.0398   1.0000   0.8901
  -1.500  -0.3393   0.01443   0.00786   0.0392   1.0000   0.9124
  -1.250  -0.2938   0.01471   0.00802   0.0355   1.0000   0.9290
  -1.000  -0.2434   0.01496   0.00818   0.0304   1.0000   0.9438
  -0.750  -0.1890   0.01519   0.00832   0.0243   1.0000   0.9571
  -0.500  -0.1346   0.01538   0.00845   0.0178   1.0000   0.9703
  -0.250  -0.0679   0.01552   0.00855   0.0090   0.9949   0.9788
   0.000   0.0000   0.01556   0.00856   0.0000   0.9868   0.9868
   0.250   0.0681   0.01552   0.00854  -0.0091   0.9788   0.9949
   0.500   0.1344   0.01538   0.00845  -0.0178   0.9703   1.0000
   0.750   0.1890   0.01519   0.00832  -0.0243   0.9571   1.0000
   1.000   0.2433   0.01496   0.00818  -0.0304   0.9439   1.0000
   1.250   0.2936   0.01470   0.00802  -0.0354   0.9291   1.0000
   1.500   0.3391   0.01443   0.00785  -0.0392   0.9124   1.0000
   1.750   0.3689   0.01426   0.00774  -0.0397   0.8901   1.0000
   2.000   0.3951   0.01406   0.00758  -0.0390   0.8678   1.0000
   2.250   0.4119   0.01395   0.00750  -0.0364   0.8426   1.0000
   2.500   0.4273   0.01388   0.00745  -0.0334   0.8173   1.0000
   2.750   0.4428   0.01381   0.00737  -0.0303   0.7922   1.0000
   3.000   0.4588   0.01375   0.00729  -0.0271   0.7664   1.0000
   3.250   0.4748   0.01371   0.00724  -0.0240   0.7382   1.0000
   3.500   0.4912   0.01367   0.00720  -0.0210   0.7073   1.0000
   3.750   0.5077   0.01362   0.00718  -0.0180   0.6712   1.0000
   4.000   0.5243   0.01359   0.00711  -0.0149   0.6273   1.0000
   4.250   0.5398   0.01362   0.00702  -0.0116   0.5639   1.0000
   4.500   0.5476   0.01423   0.00688  -0.0070   0.3728   1.0000
   4.750   0.5477   0.01689   0.00797  -0.0028   0.1691   1.0000
   5.000   0.5610   0.01839   0.00917  -0.0003   0.1290   1.0000
   5.250   0.5770   0.01975   0.01029   0.0018   0.1082   1.0000
   5.500   0.5965   0.02130   0.01167   0.0034   0.0956   1.0000
   5.750   0.6179   0.02246   0.01287   0.0047   0.0846   1.0000
   6.000   0.6415   0.02428   0.01478   0.0059   0.0786   1.0000
   6.250   0.6649   0.02593   0.01655   0.0070   0.0732   1.0000
   6.500   0.6868   0.02823   0.01894   0.0079   0.0680   1.0000
   6.750   0.7083   0.03039   0.02153   0.0095   0.0659   1.0000
   7.000   0.7280   0.03308   0.02465   0.0112   0.0652   1.0000
   7.250   0.7448   0.03604   0.02807   0.0130   0.0647   1.0000
   7.500   0.7588   0.03894   0.03140   0.0149   0.0633   1.0000
   7.750   0.7698   0.04241   0.03532   0.0169   0.0632   1.0000
   8.000   0.7778   0.04678   0.04009   0.0188   0.0660   1.0000
   8.250   0.7934   0.05140   0.04480   0.0197   0.0703   1.0000
   9.000   0.6918   0.08167   0.07698   0.0161   0.1639   1.0000
   9.250   0.6480   0.08877   0.08396   0.0078   0.1619   1.0000
   9.500   0.7033   0.08979   0.08505   0.0171   0.1508   1.0000
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