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


Details Polar file
Airfoil: S1221 w/o flap (s1221-il)
Reynolds number: 50,000
Max Cl/Cd: 27.07 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1221-il-50000.txt
Download as CSV file: xf-s1221-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S1221  w/o 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.2539   0.11445   0.10789  -0.0166   1.0000   0.2135
  -7.250  -0.2475   0.11203   0.10555  -0.0161   1.0000   0.2215
  -7.000  -0.2738   0.11430   0.10803  -0.0170   1.0000   0.2254
  -6.750  -0.2374   0.10665   0.10036  -0.0153   1.0000   0.2330
  -6.500  -0.2514   0.10696   0.10082  -0.0152   1.0000   0.2404
  -6.250  -0.2397   0.10277   0.09671  -0.0145   1.0000   0.2443
  -6.000  -0.2299   0.09975   0.09377  -0.0134   1.0000   0.2511
  -5.750  -0.2559   0.10127   0.09551  -0.0122   1.0000   0.2572
  -5.500  -0.2373   0.09624   0.09054  -0.0115   1.0000   0.2610
  -5.250  -0.2301   0.09348   0.08788  -0.0101   1.0000   0.2672
  -5.000  -0.2589   0.09477   0.08942  -0.0079   1.0000   0.2738
  -4.750  -0.2546   0.09140   0.08618  -0.0066   1.0000   0.2766
  -4.500  -0.2462   0.08835   0.08323  -0.0047   1.0000   0.2805
  -4.250  -0.2509   0.08676   0.08182  -0.0031   1.0000   0.2858
  -4.000  -0.2802   0.08779   0.08315  -0.0043   1.0000   0.2919
  -3.750  -0.2854   0.08564   0.08116  -0.0009   1.0000   0.2932
  -3.500  -0.3030   0.08534   0.08106   0.0026   1.0000   0.2944
  -3.250  -0.1867   0.04851   0.04304  -0.0753   1.0000   0.1199
  -3.000  -0.0716   0.03786   0.03125  -0.1006   0.9913   0.1128
  -2.750   0.0376   0.03214   0.02422  -0.1187   0.9736   0.1187
  -2.500   0.1167   0.02893   0.02082  -0.1281   0.9463   0.1310
  -2.250   0.1938   0.02615   0.01829  -0.1367   0.9173   0.1716
  -2.000   0.2497   0.02791   0.02125  -0.1363   0.8842   0.3953
  -1.750   0.2742   0.02928   0.02286  -0.1289   0.8583   0.4704
  -1.500   0.3063   0.02922   0.02264  -0.1255   0.8295   0.5215
  -1.250   0.3519   0.02846   0.02150  -0.1269   0.8009   0.5540
  -1.000   0.4026   0.02778   0.02033  -0.1307   0.7736   0.5774
  -0.750   0.4415   0.02749   0.01970  -0.1326   0.7478   0.5928
  -0.500   0.4806   0.02731   0.01917  -0.1347   0.7262   0.6060
  -0.250   0.5196   0.02726   0.01882  -0.1372   0.7071   0.6186
   0.000   0.5596   0.02732   0.01859  -0.1403   0.6896   0.6306
   0.250   0.5888   0.02749   0.01864  -0.1408   0.6741   0.6410
   0.500   0.6208   0.02773   0.01876  -0.1421   0.6597   0.6525
   0.750   0.6554   0.02806   0.01894  -0.1442   0.6469   0.6650
   1.000   0.6867   0.02839   0.01912  -0.1450   0.6366   0.6789
   1.250   0.7136   0.02884   0.01955  -0.1453   0.6258   0.6933
   1.500   0.7416   0.02933   0.02000  -0.1457   0.6163   0.7093
   1.750   0.7698   0.02980   0.02042  -0.1460   0.6074   0.7286
   2.000   0.7938   0.03046   0.02112  -0.1458   0.5996   0.7510
   2.250   0.8158   0.03105   0.02176  -0.1450   0.5921   0.7773
   2.500   0.8378   0.03151   0.02223  -0.1439   0.5858   0.8110
   2.750   0.8479   0.03221   0.02316  -0.1414   0.5787   0.8529
   3.000   0.8592   0.03215   0.02326  -0.1384   0.5731   1.0000
   3.250   0.9238   0.03413   0.02499  -0.1494   0.5646   1.0000
   3.500   0.9599   0.03566   0.02634  -0.1526   0.5575   1.0000
   3.750   0.9924   0.03693   0.02742  -0.1540   0.5519   1.0000
   4.000   1.0090   0.03934   0.02997  -0.1546   0.5456   1.0000
   4.250   1.0327   0.04112   0.03171  -0.1551   0.5403   1.0000
   4.500   1.0622   0.04239   0.03287  -0.1555   0.5355   1.0000
   4.750   1.0648   0.04612   0.03683  -0.1555   0.5297   1.0000
   5.000   1.0769   0.04905   0.03986  -0.1557   0.5248   1.0000
   5.250   1.1028   0.05069   0.04147  -0.1559   0.5205   1.0000
   5.500   1.1025   0.05498   0.04589  -0.1559   0.5163   1.0000
   5.750   1.0720   0.06226   0.05337  -0.1561   0.5136   1.0000
   6.000   1.0353   0.07003   0.06125  -0.1566   0.5145   1.0000
   6.250   1.0131   0.07669   0.06795  -0.1577   0.5165   1.0000
   6.500   1.0064   0.08222   0.07353  -0.1592   0.5196   1.0000
   6.750   0.8163   0.10358   0.09525  -0.1712   0.6768   1.0000
   7.000   0.8280   0.10643   0.09810  -0.1711   0.6622   1.0000
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