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S1221 w/ 4 deg flap (s1221-flap-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: S1221 w/ 4 deg flap (s1221-flap-il)
Reynolds number: 50,000
Max Cl/Cd: 31.03 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1221-flap-il-50000.txt
Download as CSV file: xf-s1221-flap-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S1221  w/ 4 deg flap                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -5.500  -0.2854   0.10201   0.09672  -0.0057   1.0000   0.2589
  -5.250  -0.2604   0.09631   0.09103  -0.0034   1.0000   0.2633
  -5.000  -0.2612   0.09443   0.08927  -0.0017   1.0000   0.2694
  -4.750  -0.2857   0.09550   0.09058  -0.0038   1.0000   0.2758
  -4.500  -0.2748   0.09128   0.08644  -0.0004   1.0000   0.2784
  -4.250  -0.2732   0.08908   0.08437   0.0016   1.0000   0.2822
  -4.000  -0.2810   0.08803   0.08349   0.0027   1.0000   0.2870
  -3.750  -0.1084   0.04699   0.04103  -0.0922   1.0000   0.1176
  -3.500  -0.0247   0.03944   0.03282  -0.1099   1.0000   0.1150
  -3.250   0.0601   0.03447   0.02675  -0.1252   1.0000   0.1182
  -3.000   0.1008   0.03331   0.02558  -0.1305   0.9957   0.1228
  -2.750   0.1949   0.03092   0.02287  -0.1429   0.9770   0.1473
  -2.500   0.2807   0.03078   0.02395  -0.1515   0.9444   0.3128
  -2.250   0.2969   0.03316   0.02677  -0.1432   0.9017   0.4174
  -2.000   0.3105   0.03389   0.02763  -0.1352   0.8683   0.4712
  -1.750   0.3442   0.03317   0.02676  -0.1323   0.8440   0.5194
  -1.500   0.3905   0.03181   0.02512  -0.1339   0.8160   0.5514
  -1.250   0.4417   0.03027   0.02329  -0.1360   0.7930   0.5738
  -1.000   0.5116   0.02862   0.02118  -0.1424   0.7695   0.5947
  -0.750   0.5773   0.02770   0.01973  -0.1490   0.7434   0.6106
  -0.500   0.6343   0.02745   0.01903  -0.1547   0.7184   0.6238
  -0.250   0.6894   0.02752   0.01868  -0.1606   0.6966   0.6363
   0.000   0.7259   0.02779   0.01876  -0.1624   0.6785   0.6470
   0.250   0.7657   0.02811   0.01888  -0.1653   0.6624   0.6574
   0.500   0.8029   0.02856   0.01923  -0.1680   0.6480   0.6704
   0.750   0.8315   0.02904   0.01965  -0.1685   0.6359   0.6827
   1.000   0.8653   0.02950   0.01999  -0.1699   0.6251   0.6963
   1.250   0.8951   0.03007   0.02053  -0.1708   0.6144   0.7114
   1.500   0.9240   0.03071   0.02115  -0.1714   0.6057   0.7294
   1.750   0.9523   0.03135   0.02178  -0.1720   0.5970   0.7498
   2.000   0.9804   0.03206   0.02250  -0.1726   0.5892   0.7731
   2.250   1.0040   0.03275   0.02326  -0.1724   0.5812   0.8006
   2.500   1.0278   0.03312   0.02358  -0.1715   0.5754   0.8346
   2.750   1.0363   0.03395   0.02468  -0.1690   0.5690   0.8727
   3.000   1.0456   0.03421   0.02515  -0.1661   0.5634   1.0000
   3.250   1.1102   0.03606   0.02681  -0.1758   0.5550   1.0000
   3.500   1.1480   0.03812   0.02887  -0.1801   0.5475   1.0000
   3.750   1.1873   0.03952   0.03012  -0.1830   0.5418   1.0000
   4.000   1.2052   0.04196   0.03269  -0.1834   0.5359   1.0000
   4.250   1.2241   0.04422   0.03502  -0.1836   0.5298   1.0000
   4.500   1.2528   0.04579   0.03654  -0.1843   0.5249   1.0000
   4.750   1.2618   0.04894   0.03988  -0.1836   0.5202   1.0000
   5.000   1.2589   0.05302   0.04418  -0.1823   0.5149   1.0000
   5.250   1.2734   0.05572   0.04694  -0.1820   0.5098   1.0000
   5.500   1.3032   0.05734   0.04852  -0.1823   0.5053   1.0000
   5.750   1.2447   0.06675   0.05827  -0.1793   0.5027   1.0000
   6.000   1.1709   0.07805   0.06973  -0.1775   0.5036   1.0000
   6.250   1.1371   0.08701   0.07875  -0.1791   0.5065   1.0000
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