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RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAE 102 AIRFOIL (rae102-il)
Reynolds number: 50,000
Max Cl/Cd: 27.39 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rae102-il-50000-n5.txt
Download as CSV file: xf-rae102-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.7472   0.07692   0.06954  -0.0272   1.0000   0.0445
 -10.000  -0.7632   0.07278   0.06530  -0.0276   1.0000   0.0443
  -9.750  -0.7788   0.06905   0.06142  -0.0267   1.0000   0.0443
  -9.500  -0.7904   0.06511   0.05725  -0.0258   1.0000   0.0443
  -9.250  -0.7972   0.06125   0.05311  -0.0246   1.0000   0.0443
  -9.000  -0.8007   0.05742   0.04893  -0.0232   1.0000   0.0445
  -8.750  -0.7993   0.05360   0.04480  -0.0217   1.0000   0.0448
  -8.500  -0.7928   0.05000   0.04093  -0.0205   1.0000   0.0455
  -8.250  -0.7829   0.04693   0.03761  -0.0192   1.0000   0.0467
  -8.000  -0.7702   0.04447   0.03488  -0.0181   1.0000   0.0491
  -7.750  -0.7565   0.04188   0.03192  -0.0167   1.0000   0.0520
  -7.500  -0.7400   0.03908   0.02861  -0.0152   1.0000   0.0546
  -7.250  -0.7194   0.03648   0.02547  -0.0140   1.0000   0.0565
  -7.000  -0.6980   0.03422   0.02326  -0.0133   1.0000   0.0602
  -6.750  -0.6760   0.03251   0.02132  -0.0123   1.0000   0.0665
  -6.500  -0.6517   0.03063   0.01930  -0.0114   1.0000   0.0717
  -6.250  -0.6296   0.02914   0.01768  -0.0103   1.0000   0.0792
  -6.000  -0.6096   0.02773   0.01626  -0.0090   1.0000   0.0895
  -5.750  -0.5908   0.02634   0.01482  -0.0074   1.0000   0.1006
  -5.500  -0.5737   0.02502   0.01354  -0.0058   1.0000   0.1196
  -5.250  -0.5583   0.02355   0.01223  -0.0041   1.0000   0.1467
  -5.000  -0.5453   0.02184   0.01095  -0.0023   1.0000   0.2057
  -4.750  -0.5369   0.01995   0.01007   0.0003   1.0000   0.3414
  -4.500  -0.5264   0.01921   0.01005   0.0041   1.0000   0.4962
  -4.250  -0.5121   0.01902   0.01000   0.0075   1.0000   0.5760
  -4.000  -0.4963   0.01898   0.01001   0.0109   1.0000   0.6330
  -3.750  -0.4800   0.01899   0.00998   0.0143   1.0000   0.6789
  -3.500  -0.4616   0.01906   0.01003   0.0175   1.0000   0.7165
  -3.250  -0.4428   0.01907   0.00999   0.0203   1.0000   0.7489
  -3.000  -0.4217   0.01907   0.00990   0.0225   1.0000   0.7759
  -2.750  -0.3998   0.01900   0.00973   0.0243   1.0000   0.7993
  -2.500  -0.3756   0.01891   0.00951   0.0254   1.0000   0.8197
  -2.250  -0.3526   0.01877   0.00927   0.0263   1.0000   0.8396
  -2.000  -0.3237   0.01870   0.00909   0.0261   1.0000   0.8567
  -1.750  -0.2933   0.01863   0.00890   0.0254   1.0000   0.8735
  -1.500  -0.2610   0.01859   0.00876   0.0242   1.0000   0.8902
  -1.250  -0.2267   0.01856   0.00865   0.0223   1.0000   0.9068
  -1.000  -0.1904   0.01855   0.00857   0.0199   1.0000   0.9233
  -0.750  -0.1518   0.01856   0.00850   0.0168   1.0000   0.9400
  -0.500  -0.1104   0.01859   0.00847   0.0130   1.0000   0.9560
  -0.250  -0.0564   0.01861   0.00845   0.0067   0.9905   0.9675
   0.000   0.0000   0.01863   0.00845   0.0000   0.9788   0.9789
   0.250   0.0565   0.01861   0.00845  -0.0067   0.9675   0.9905
   0.500   0.1103   0.01858   0.00847  -0.0130   0.9560   1.0000
   0.750   0.1518   0.01856   0.00849  -0.0168   0.9400   1.0000
   1.000   0.1904   0.01855   0.00856  -0.0199   0.9234   1.0000
   1.250   0.2266   0.01856   0.00865  -0.0223   0.9068   1.0000
   1.500   0.2609   0.01858   0.00876  -0.0242   0.8902   1.0000
   1.750   0.2932   0.01863   0.00890  -0.0254   0.8735   1.0000
   2.000   0.3236   0.01869   0.00909  -0.0261   0.8567   1.0000
   2.250   0.3525   0.01877   0.00926  -0.0263   0.8396   1.0000
   2.500   0.3754   0.01891   0.00950  -0.0253   0.8197   1.0000
   2.750   0.3996   0.01899   0.00973  -0.0242   0.7993   1.0000
   3.000   0.4215   0.01906   0.00990  -0.0225   0.7759   1.0000
   3.250   0.4426   0.01907   0.00998  -0.0203   0.7490   1.0000
   3.500   0.4614   0.01906   0.01003  -0.0174   0.7167   1.0000
   3.750   0.4798   0.01899   0.00997  -0.0143   0.6791   1.0000
   4.000   0.4961   0.01898   0.01000  -0.0109   0.6332   1.0000
   4.250   0.5119   0.01902   0.01000  -0.0075   0.5763   1.0000
   4.500   0.5262   0.01921   0.01005  -0.0041   0.4967   1.0000
   4.750   0.5367   0.01995   0.01007  -0.0003   0.3407   1.0000
   5.000   0.5452   0.02184   0.01095   0.0023   0.2054   1.0000
   5.250   0.5583   0.02355   0.01223   0.0041   0.1469   1.0000
   5.500   0.5738   0.02502   0.01353   0.0058   0.1197   1.0000
   5.750   0.5908   0.02634   0.01482   0.0074   0.1004   1.0000
   6.000   0.6097   0.02774   0.01626   0.0090   0.0894   1.0000
   6.250   0.6297   0.02914   0.01768   0.0102   0.0791   1.0000
   6.500   0.6519   0.03063   0.01930   0.0113   0.0717   1.0000
   6.750   0.6761   0.03252   0.02132   0.0123   0.0663   1.0000
   7.000   0.6981   0.03423   0.02326   0.0133   0.0602   1.0000
   7.250   0.7196   0.03649   0.02548   0.0139   0.0565   1.0000
   7.500   0.7402   0.03910   0.02863   0.0152   0.0546   1.0000
   7.750   0.7568   0.04188   0.03192   0.0166   0.0520   1.0000
   8.000   0.7706   0.04450   0.03492   0.0180   0.0492   1.0000
   8.250   0.7833   0.04695   0.03762   0.0192   0.0467   1.0000
   8.500   0.7933   0.05002   0.04095   0.0204   0.0455   1.0000
   8.750   0.7997   0.05361   0.04482   0.0217   0.0448   1.0000
   9.000   0.8013   0.05745   0.04896   0.0231   0.0445   1.0000
   9.250   0.7981   0.06125   0.05311   0.0245   0.0444   1.0000
   9.500   0.7907   0.06518   0.05733   0.0257   0.0443   1.0000
   9.750   0.7798   0.06909   0.06147   0.0266   0.0443   1.0000
  10.000   0.7650   0.07284   0.06536   0.0274   0.0444   1.0000
  10.250   0.7500   0.07692   0.06953   0.0271   0.0445   1.0000
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