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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 200,000
Max Cl/Cd: 46.19 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20606-il-200000-n5.txt
Download as CSV file: xf-sc20606-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6194   0.08566   0.08201  -0.0137   1.0000   0.0160
  -9.250  -0.6225   0.08048   0.07688  -0.0169   1.0000   0.0160
  -9.000  -0.6276   0.07474   0.07119  -0.0214   1.0000   0.0159
  -8.750  -0.6288   0.06713   0.06354  -0.0305   1.0000   0.0157
  -8.500  -0.6281   0.06065   0.05694  -0.0357   1.0000   0.0156
  -8.250  -0.6237   0.05446   0.05056  -0.0395   1.0000   0.0153
  -8.000  -0.6145   0.04832   0.04414  -0.0428   1.0000   0.0151
  -7.750  -0.5998   0.04219   0.03762  -0.0456   1.0000   0.0148
  -7.500  -0.5796   0.03623   0.03116  -0.0482   1.0000   0.0146
  -7.250  -0.5544   0.03086   0.02519  -0.0505   1.0000   0.0146
  -7.000  -0.5265   0.02659   0.02031  -0.0522   1.0000   0.0149
  -6.750  -0.4978   0.02346   0.01664  -0.0532   1.0000   0.0154
  -6.500  -0.4696   0.02121   0.01397  -0.0538   1.0000   0.0160
  -6.250  -0.4423   0.02007   0.01252  -0.0539   1.0000   0.0170
  -6.000  -0.4146   0.01804   0.01030  -0.0546   1.0000   0.0188
  -5.750  -0.3870   0.01678   0.00890  -0.0548   1.0000   0.0195
  -5.500  -0.3590   0.01569   0.00771  -0.0551   1.0000   0.0206
  -5.250  -0.3303   0.01474   0.00668  -0.0556   1.0000   0.0222
  -5.000  -0.3011   0.01394   0.00577  -0.0562   1.0000   0.0244
  -4.750  -0.2708   0.01313   0.00485  -0.0570   1.0000   0.0287
  -4.500  -0.2421   0.01269   0.00439  -0.0575   1.0000   0.0381
  -4.250  -0.2122   0.01214   0.00385  -0.0583   1.0000   0.0533
  -4.000  -0.1822   0.01162   0.00347  -0.0592   1.0000   0.0847
  -3.750  -0.1498   0.01084   0.00309  -0.0609   1.0000   0.1838
  -3.500  -0.1147   0.00985   0.00280  -0.0636   1.0000   0.3687
  -3.250  -0.0826   0.00922   0.00285  -0.0649   1.0000   0.5519
  -3.000  -0.0553   0.00910   0.00291  -0.0648   1.0000   0.6180
  -2.750  -0.0288   0.00907   0.00297  -0.0643   1.0000   0.6653
  -2.500  -0.0031   0.00909   0.00307  -0.0636   1.0000   0.7013
  -2.250   0.0229   0.00913   0.00314  -0.0631   1.0000   0.7240
  -2.000   0.0490   0.00917   0.00321  -0.0626   1.0000   0.7411
  -1.750   0.0744   0.00922   0.00331  -0.0620   1.0000   0.7578
  -1.500   0.1004   0.00927   0.00338  -0.0615   1.0000   0.7713
  -1.250   0.1274   0.00932   0.00345  -0.0613   1.0000   0.7823
  -1.000   0.1534   0.00936   0.00355  -0.0609   1.0000   0.7902
  -0.750   0.1800   0.00942   0.00366  -0.0607   1.0000   0.7989
  -0.500   0.2069   0.00949   0.00379  -0.0606   1.0000   0.8076
  -0.250   0.2330   0.00955   0.00395  -0.0603   1.0000   0.8146
   0.000   0.2600   0.00964   0.00412  -0.0602   1.0000   0.8228
   0.250   0.3104   0.00942   0.00403  -0.0649   0.9837   0.8280
   0.500   0.3649   0.00893   0.00367  -0.0699   0.9365   0.8339
   0.750   0.4217   0.00913   0.00311  -0.0743   0.6775   0.8386
   1.000   0.4360   0.01077   0.00344  -0.0715   0.3810   0.8457
   1.250   0.4578   0.01185   0.00383  -0.0707   0.1983   0.8529
   1.500   0.4813   0.01273   0.00421  -0.0701   0.0891   0.8606
   1.750   0.5062   0.01325   0.00460  -0.0696   0.0549   0.8680
   2.000   0.5318   0.01375   0.00508  -0.0690   0.0397   0.8761
   2.250   0.5563   0.01419   0.00557  -0.0682   0.0297   0.8836
   2.500   0.5810   0.01493   0.00635  -0.0675   0.0247   0.8919
   2.750   0.6047   0.01557   0.00716  -0.0664   0.0225   0.9003
   3.000   0.6282   0.01637   0.00806  -0.0653   0.0208   0.9096
   3.250   0.6519   0.01728   0.00908  -0.0643   0.0196   0.9196
   3.500   0.6750   0.01830   0.01022  -0.0631   0.0188   0.9304
   3.750   0.6982   0.01943   0.01149  -0.0619   0.0181   0.9441
   4.000   0.7206   0.02133   0.01363  -0.0608   0.0164   0.9675
   4.250   0.7491   0.02253   0.01512  -0.0608   0.0155   1.0000
   4.500   0.7763   0.02477   0.01775  -0.0605   0.0149   1.0000
   4.750   0.8018   0.02766   0.02115  -0.0599   0.0144   1.0000
   5.000   0.8248   0.03132   0.02537  -0.0588   0.0142   1.0000
   5.250   0.8449   0.03577   0.03038  -0.0575   0.0141   1.0000
   5.500   0.8620   0.04076   0.03590  -0.0561   0.0143   1.0000
   5.750   0.8764   0.04612   0.04171  -0.0549   0.0145   1.0000
   6.000   0.8882   0.05168   0.04766  -0.0539   0.0147   1.0000
   6.250   0.8986   0.05755   0.05388  -0.0533   0.0143   1.0000
   6.500   0.9068   0.06351   0.06014  -0.0534   0.0135   1.0000
   6.750   0.9119   0.06936   0.06621  -0.0541   0.0130   1.0000
   7.000   0.9126   0.07569   0.07271  -0.0555   0.0131   1.0000
   7.250   0.9101   0.08211   0.07927  -0.0578   0.0133   1.0000
   7.500   0.9046   0.08845   0.08568  -0.0610   0.0135   1.0000
   7.750   0.8948   0.09445   0.09171  -0.0648   0.0137   1.0000
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