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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 1,000,000
Max Cl/Cd: 65.63 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20606-il-1000000-n5.txt
Download as CSV file: xf-sc20606-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0606 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6636   0.08106   0.07945  -0.0137   1.0000   0.0046
  -9.750  -0.6755   0.07418   0.07261  -0.0183   1.0000   0.0046
  -9.500  -0.6959   0.06157   0.05997  -0.0336   1.0000   0.0045
  -9.250  -0.7447   0.02779   0.02491  -0.0509   1.0000   0.0050
  -9.000  -0.7219   0.02262   0.01918  -0.0533   1.0000   0.0052
  -8.750  -0.6958   0.01992   0.01611  -0.0547   1.0000   0.0054
  -8.500  -0.6688   0.01791   0.01380  -0.0557   1.0000   0.0056
  -8.250  -0.6415   0.01643   0.01210  -0.0565   1.0000   0.0057
  -8.000  -0.6146   0.01552   0.01104  -0.0569   1.0000   0.0059
  -7.750  -0.5877   0.01477   0.01017  -0.0572   1.0000   0.0061
  -7.500  -0.5602   0.01386   0.00909  -0.0576   1.0000   0.0062
  -7.250  -0.5311   0.01256   0.00760  -0.0584   1.0000   0.0064
  -7.000  -0.5004   0.01122   0.00606  -0.0597   1.0000   0.0070
  -6.750  -0.4725   0.01070   0.00549  -0.0601   1.0000   0.0075
  -6.500  -0.4448   0.01031   0.00505  -0.0603   1.0000   0.0079
  -6.250  -0.4170   0.00993   0.00460  -0.0606   1.0000   0.0082
  -6.000  -0.3890   0.00956   0.00419  -0.0609   1.0000   0.0086
  -5.750  -0.3609   0.00923   0.00380  -0.0611   1.0000   0.0090
  -5.500  -0.3328   0.00893   0.00346  -0.0614   1.0000   0.0093
  -5.250  -0.3049   0.00867   0.00316  -0.0616   1.0000   0.0097
  -5.000  -0.2770   0.00848   0.00293  -0.0618   0.9999   0.0102
  -4.750  -0.2452   0.00819   0.00259  -0.0628   0.9995   0.0114
  -4.500  -0.2135   0.00795   0.00236  -0.0638   0.9991   0.0150
  -4.250  -0.1817   0.00773   0.00219  -0.0648   0.9985   0.0224
  -4.000  -0.1497   0.00757   0.00205  -0.0659   0.9978   0.0302
  -3.750  -0.1174   0.00741   0.00191  -0.0670   0.9971   0.0384
  -3.500  -0.0852   0.00723   0.00180  -0.0681   0.9964   0.0532
  -3.250  -0.0525   0.00691   0.00169  -0.0695   0.9959   0.1048
  -3.000  -0.0186   0.00657   0.00159  -0.0712   0.9954   0.1715
  -2.750   0.0175   0.00625   0.00147  -0.0733   0.9945   0.2368
  -2.500   0.0535   0.00588   0.00134  -0.0754   0.9922   0.3149
  -2.250   0.0932   0.00530   0.00121  -0.0785   0.9888   0.4505
  -2.000   0.1355   0.00485   0.00112  -0.0820   0.9846   0.5641
  -1.750   0.1690   0.00473   0.00107  -0.0833   0.9783   0.5894
  -1.500   0.2046   0.00458   0.00101  -0.0850   0.9672   0.6158
  -1.250   0.2422   0.00449   0.00093  -0.0870   0.9423   0.6310
  -1.000   0.2691   0.00466   0.00091  -0.0864   0.8643   0.6451
  -0.750   0.2868   0.00611   0.00109  -0.0842   0.5406   0.6580
  -0.500   0.3124   0.00683   0.00123  -0.0841   0.3735   0.6725
  -0.250   0.3397   0.00719   0.00135  -0.0841   0.2890   0.6867
   0.000   0.3669   0.00756   0.00148  -0.0841   0.2093   0.6981
   0.250   0.3942   0.00792   0.00162  -0.0841   0.1368   0.7078
   0.500   0.4214   0.00834   0.00178  -0.0841   0.0652   0.7160
   0.750   0.4492   0.00854   0.00191  -0.0841   0.0442   0.7242
   1.000   0.4771   0.00870   0.00207  -0.0841   0.0342   0.7316
   1.250   0.5049   0.00888   0.00223  -0.0841   0.0252   0.7390
   1.500   0.5326   0.00908   0.00240  -0.0841   0.0174   0.7454
   1.750   0.5604   0.00929   0.00261  -0.0840   0.0133   0.7523
   2.000   0.5879   0.00957   0.00292  -0.0838   0.0105   0.7582
   2.250   0.6154   0.00984   0.00326  -0.0837   0.0098   0.7642
   2.500   0.6429   0.01007   0.00353  -0.0836   0.0094   0.7699
   2.750   0.6702   0.01037   0.00388  -0.0834   0.0090   0.7761
   3.000   0.6973   0.01070   0.00428  -0.0832   0.0087   0.7824
   3.250   0.7242   0.01106   0.00471  -0.0830   0.0083   0.7885
   3.500   0.7510   0.01145   0.00518  -0.0827   0.0080   0.7946
   3.750   0.7777   0.01185   0.00565  -0.0824   0.0076   0.8001
   4.000   0.8043   0.01227   0.00613  -0.0822   0.0073   0.8059
   4.250   0.8306   0.01276   0.00669  -0.0818   0.0069   0.8114
   4.500   0.8558   0.01360   0.00765  -0.0813   0.0064   0.8173
   5.000   0.9044   0.01616   0.01065  -0.0797   0.0059   0.8302
   5.250   0.9301   0.01685   0.01146  -0.0793   0.0058   0.8364
   5.500   0.9553   0.01759   0.01235  -0.0788   0.0056   0.8417
   5.750   0.9796   0.01870   0.01365  -0.0781   0.0054   0.8472
   6.000   1.0029   0.02020   0.01540  -0.0773   0.0053   0.8525
   6.250   1.0249   0.02202   0.01752  -0.0762   0.0051   0.8587
   6.500   1.0452   0.02436   0.02021  -0.0750   0.0050   0.8645
   6.750   1.0616   0.02780   0.02411  -0.0732   0.0047   0.8703
   7.000   1.0268   0.05128   0.04904  -0.0655   0.0041   0.8750
   7.250   1.0253   0.05951   0.05758  -0.0642   0.0039   0.8810
   7.500   1.0232   0.06674   0.06505  -0.0640   0.0038   0.8880
   7.750   1.0169   0.07416   0.07266  -0.0648   0.0037   0.8954
   8.250   0.9857   0.08953   0.08828  -0.0703   0.0037   0.9189
   8.500   0.9590   0.09841   0.09724  -0.0780   0.0038   0.9539
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