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NASA SC(2)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il)
Reynolds number: 50,000
Max Cl/Cd: 24.61 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20610-il-50000-n5.txt
Download as CSV file: xf-sc20610-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0610 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.6210   0.09866   0.09051  -0.0345   1.0000   0.0608
 -11.000  -0.6326   0.09153   0.08342  -0.0388   1.0000   0.0606
 -10.750  -0.6489   0.08486   0.07675  -0.0428   1.0000   0.0603
 -10.500  -0.6683   0.07899   0.07087  -0.0457   1.0000   0.0600
 -10.250  -0.6902   0.07413   0.06599  -0.0470   1.0000   0.0597
 -10.000  -0.7102   0.06947   0.06123  -0.0483   1.0000   0.0596
  -9.750  -0.7250   0.06486   0.05642  -0.0493   1.0000   0.0598
  -9.500  -0.7342   0.06038   0.05162  -0.0499   1.0000   0.0602
  -9.250  -0.7371   0.05596   0.04676  -0.0505   1.0000   0.0609
  -9.000  -0.7312   0.05212   0.04260  -0.0506   1.0000   0.0617
  -8.750  -0.7193   0.04900   0.03928  -0.0505   1.0000   0.0627
  -8.500  -0.7048   0.04616   0.03618  -0.0504   1.0000   0.0641
  -8.250  -0.6883   0.04371   0.03350  -0.0503   1.0000   0.0665
  -8.000  -0.6700   0.04110   0.03049  -0.0504   1.0000   0.0699
  -7.750  -0.6493   0.03839   0.02717  -0.0504   1.0000   0.0729
  -7.500  -0.6287   0.03628   0.02501  -0.0499   1.0000   0.0756
  -7.250  -0.6073   0.03471   0.02330  -0.0494   1.0000   0.0802
  -7.000  -0.5845   0.03303   0.02127  -0.0488   1.0000   0.0854
  -6.750  -0.5631   0.03154   0.01978  -0.0477   1.0000   0.0899
  -6.500  -0.5411   0.03038   0.01851  -0.0467   1.0000   0.0969
  -6.250  -0.5199   0.02922   0.01727  -0.0451   1.0000   0.1031
  -6.000  -0.4986   0.02828   0.01630  -0.0437   1.0000   0.1117
  -5.750  -0.4776   0.02731   0.01537  -0.0424   1.0000   0.1207
  -5.500  -0.4559   0.02641   0.01444  -0.0413   1.0000   0.1322
  -5.250  -0.4334   0.02545   0.01353  -0.0407   1.0000   0.1476
  -5.000  -0.4096   0.02440   0.01265  -0.0408   1.0000   0.1698
  -4.750  -0.3840   0.02312   0.01170  -0.0415   1.0000   0.2098
  -4.500  -0.3542   0.02115   0.01077  -0.0441   1.0000   0.3407
  -4.250  -0.3498   0.02127   0.01208  -0.0370   1.0000   0.5103
  -4.000  -0.3319   0.02192   0.01271  -0.0336   1.0000   0.6087
  -3.750  -0.3106   0.02254   0.01316  -0.0315   1.0000   0.6592
  -3.500  -0.2948   0.02322   0.01375  -0.0277   1.0000   0.6920
  -3.250  -0.2797   0.02385   0.01426  -0.0239   1.0000   0.7225
  -3.000  -0.2710   0.02436   0.01475  -0.0182   1.0000   0.7474
  -2.750  -0.2589   0.02468   0.01500  -0.0138   1.0000   0.7739
  -2.500  -0.2483   0.02474   0.01502  -0.0092   1.0000   0.7958
  -2.250  -0.2324   0.02463   0.01481  -0.0063   1.0000   0.8143
  -2.000  -0.2144   0.02439   0.01449  -0.0043   1.0000   0.8277
  -1.750  -0.1929   0.02414   0.01415  -0.0034   1.0000   0.8380
  -1.500  -0.1679   0.02395   0.01388  -0.0034   1.0000   0.8475
  -1.250  -0.1462   0.02368   0.01354  -0.0026   1.0000   0.8551
  -1.000  -0.1195   0.02355   0.01335  -0.0032   1.0000   0.8628
  -0.750  -0.0967   0.02332   0.01309  -0.0027   1.0000   0.8691
  -0.500  -0.0702   0.02322   0.01296  -0.0033   1.0000   0.8750
  -0.250  -0.0437   0.02314   0.01287  -0.0039   1.0000   0.8801
   0.000  -0.0185   0.02305   0.01279  -0.0042   1.0000   0.8852
   0.250   0.0090   0.02307   0.01284  -0.0051   1.0000   0.8903
   0.500   0.0349   0.02307   0.01288  -0.0055   1.0000   0.8951
   0.750   0.0607   0.02310   0.01298  -0.0061   1.0000   0.8999
   1.000   0.0881   0.02323   0.01319  -0.0070   1.0000   0.9048
   1.250   0.1139   0.02335   0.01340  -0.0076   1.0000   0.9093
   1.500   0.1395   0.02351   0.01368  -0.0083   1.0000   0.9140
   1.750   0.1826   0.02383   0.01416  -0.0123   0.9890   0.9177
   2.000   0.2368   0.02412   0.01465  -0.0180   0.9717   0.9207
   2.250   0.2891   0.02401   0.01477  -0.0226   0.9461   0.9236
   2.500   0.3460   0.02329   0.01434  -0.0268   0.9061   0.9258
   2.750   0.3969   0.02212   0.01343  -0.0287   0.8559   0.9281
   3.000   0.4318   0.02115   0.01269  -0.0278   0.7926   0.9315
   3.250   0.4958   0.02015   0.01091  -0.0294   0.5320   0.9321
   3.500   0.5100   0.02185   0.01121  -0.0268   0.3133   0.9361
   3.750   0.5289   0.02327   0.01197  -0.0259   0.2161   0.9410
   4.000   0.5529   0.02445   0.01282  -0.0258   0.1734   0.9460
   4.250   0.5797   0.02541   0.01370  -0.0259   0.1478   0.9509
   4.500   0.6086   0.02643   0.01469  -0.0264   0.1310   0.9559
   5.000   0.6732   0.02872   0.01712  -0.0282   0.1068   0.9664
   5.750   0.7697   0.03322   0.02184  -0.0315   0.0838   1.0000
   6.000   0.8008   0.03531   0.02437  -0.0322   0.0787   1.0000
   6.250   0.8289   0.03720   0.02651  -0.0329   0.0735   1.0000
   6.500   0.8565   0.03956   0.02892  -0.0338   0.0703   1.0000
   6.750   0.8804   0.04248   0.03249  -0.0336   0.0675   1.0000
   7.000   0.9016   0.04536   0.03588  -0.0335   0.0642   1.0000
   7.250   0.9217   0.04812   0.03897  -0.0334   0.0616   1.0000
   7.500   0.9400   0.05124   0.04244  -0.0333   0.0603   1.0000
   7.750   0.9554   0.05468   0.04618  -0.0331   0.0593   1.0000
   8.000   0.9669   0.05856   0.05035  -0.0327   0.0584   1.0000
   8.250   0.9678   0.06317   0.05555  -0.0315   0.0577   1.0000
   8.500   0.9632   0.06804   0.06093  -0.0305   0.0570   1.0000
   8.750   0.9535   0.07307   0.06636  -0.0297   0.0566   1.0000
   9.000   0.9385   0.07823   0.07182  -0.0294   0.0564   1.0000
   9.250   0.9193   0.08340   0.07721  -0.0296   0.0565   1.0000
   9.500   0.8993   0.08946   0.08342  -0.0317   0.0568   1.0000
   9.750   0.8811   0.09645   0.09053  -0.0359   0.0571   1.0000
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