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S1223 (s1223-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: S1223 (s1223-il)
Reynolds number: 50,000
Max Cl/Cd: 33.06 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1223-il-50000.txt
Download as CSV file: xf-s1223-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S1223                                           
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -4.250  -0.0828   0.11564   0.11083  -0.0447   0.9014   0.1609
  -4.000  -0.0475   0.11002   0.10518  -0.0465   0.8871   0.1690
  -3.750  -0.0409   0.11072   0.10591  -0.0538   0.8647   0.1762
  -3.500  -0.0071   0.10406   0.09923  -0.0546   0.8541   0.1816
  -3.250   0.0129   0.10130   0.09647  -0.0570   0.8379   0.1907
  -3.000   0.0288   0.09977   0.09497  -0.0641   0.8207   0.1972
  -2.750   0.0461   0.09568   0.09088  -0.0612   0.8071   0.2035
  -2.500   0.0733   0.09527   0.09045  -0.0735   0.7904   0.2159
  -2.250   0.0961   0.08932   0.08448  -0.0686   0.7834   0.2224
  -2.000   0.1225   0.08867   0.08381  -0.0790   0.7675   0.2362
  -1.750   0.1506   0.08297   0.07808  -0.0754   0.7629   0.2462
  -1.500   0.1642   0.08205   0.07718  -0.0799   0.7474   0.2583
  -1.250   0.1763   0.08012   0.07525  -0.0783   0.7365   0.2700
  -1.000   0.2092   0.07659   0.07168  -0.0819   0.7280   0.2854
  -0.750   0.2604   0.07211   0.06711  -0.0888   0.7225   0.3060
  -0.500   0.2752   0.07178   0.06679  -0.0925   0.7092   0.3186
  -0.250   0.3430   0.06612   0.06098  -0.1007   0.7058   0.3412
   0.000   1.0370   0.03358   0.02536  -0.2494   0.6974   0.3628
   0.250   1.0579   0.03407   0.02590  -0.2458   0.6878   0.4698
   0.500   1.0660   0.03492   0.02683  -0.2413   0.6781   0.5138
   0.750   1.1061   0.03496   0.02662  -0.2430   0.6690   0.5588
   1.000   1.1377   0.03567   0.02717  -0.2444   0.6595   0.5859
   1.250   1.1734   0.03607   0.02739  -0.2460   0.6508   0.6117
   1.500   1.2120   0.03675   0.02786  -0.2485   0.6431   0.6377
   1.750   1.2302   0.03803   0.02916  -0.2477   0.6354   0.6598
   2.000   1.2700   0.03842   0.02938  -0.2499   0.6290   0.6930
   2.250   1.2815   0.03995   0.03101  -0.2478   0.6225   0.7201
   2.500   1.2896   0.04167   0.03285  -0.2454   0.6157   0.7528
   2.750   1.3134   0.04213   0.03333  -0.2446   0.6105   0.8101
   3.000   1.3201   0.04244   0.03376  -0.2407   0.6067   0.9341
   3.250   1.2851   0.04798   0.03967  -0.2347   0.5998   1.0000
   3.500   1.3406   0.05048   0.04176  -0.2427   0.5928   1.0000
   3.750   1.4128   0.04999   0.04075  -0.2493   0.5882   1.0000
   4.000   1.1095   0.07727   0.06975  -0.2200   0.5846   0.8296
   4.250   1.0692   0.08693   0.07957  -0.2200   0.5843   0.8182
   4.500   1.0523   0.09182   0.08467  -0.2186   0.5853   0.8746
   4.750   1.0753   0.09731   0.08997  -0.2242   0.5868   1.0000
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