Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0614 AIRFOIL (sc20614-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0614 AIRFOIL (sc20614-il)
Reynolds number: 100,000
Max Cl/Cd: 31.71 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20614-il-100000-n5.txt
Download as CSV file: xf-sc20614-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0614 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.000  -0.8101   0.07460   0.06760  -0.0645   1.0000   0.0516
 -13.750  -0.8474   0.06704   0.05976  -0.0680   1.0000   0.0516
 -13.500  -0.8715   0.06184   0.05433  -0.0694   1.0000   0.0519
 -13.250  -0.8895   0.05774   0.05001  -0.0697   1.0000   0.0523
 -13.000  -0.9044   0.05431   0.04636  -0.0692   1.0000   0.0528
 -12.750  -0.9181   0.05136   0.04318  -0.0679   1.0000   0.0535
 -12.500  -0.9323   0.04881   0.04040  -0.0657   1.0000   0.0541
 -12.250  -0.9476   0.04661   0.03795  -0.0626   1.0000   0.0548
 -12.000  -0.9447   0.04508   0.03639  -0.0608   1.0000   0.0556
 -11.750  -0.9388   0.04381   0.03510  -0.0593   1.0000   0.0564
 -11.500  -0.9329   0.04246   0.03369  -0.0578   1.0000   0.0574
 -11.250  -0.9262   0.04101   0.03212  -0.0563   1.0000   0.0585
 -11.000  -0.9183   0.03948   0.03043  -0.0548   1.0000   0.0598
 -10.750  -0.9090   0.03792   0.02865  -0.0533   1.0000   0.0612
 -10.500  -0.8976   0.03652   0.02707  -0.0518   1.0000   0.0626
 -10.250  -0.8838   0.03558   0.02618  -0.0505   1.0000   0.0640
 -10.000  -0.8698   0.03463   0.02520  -0.0493   1.0000   0.0658
  -9.750  -0.8551   0.03360   0.02406  -0.0481   1.0000   0.0679
  -9.500  -0.8394   0.03253   0.02278  -0.0469   1.0000   0.0702
  -9.250  -0.8230   0.03160   0.02188  -0.0457   1.0000   0.0720
  -9.000  -0.8061   0.03079   0.02111  -0.0447   1.0000   0.0741
  -8.750  -0.7885   0.03000   0.02026  -0.0437   1.0000   0.0770
  -8.500  -0.7700   0.02923   0.01934  -0.0427   1.0000   0.0800
  -8.250  -0.7518   0.02846   0.01866  -0.0419   1.0000   0.0825
  -8.000  -0.7324   0.02779   0.01800  -0.0412   1.0000   0.0859
  -7.750  -0.7121   0.02718   0.01729  -0.0405   1.0000   0.0897
  -7.500  -0.6918   0.02648   0.01666  -0.0400   1.0000   0.0931
  -7.250  -0.6686   0.02588   0.01608  -0.0400   0.9995   0.0977
  -7.000  -0.6371   0.02526   0.01542  -0.0415   0.9969   0.1032
  -6.750  -0.6051   0.02461   0.01486  -0.0432   0.9944   0.1093
  -6.500  -0.5729   0.02407   0.01431  -0.0448   0.9921   0.1165
  -6.250  -0.5420   0.02349   0.01380  -0.0463   0.9893   0.1246
  -6.000  -0.5098   0.02293   0.01332  -0.0480   0.9865   0.1350
  -5.750  -0.4760   0.02240   0.01286  -0.0500   0.9839   0.1488
  -5.500  -0.4406   0.02183   0.01244  -0.0525   0.9817   0.1691
  -5.250  -0.4057   0.02119   0.01202  -0.0550   0.9795   0.2018
  -5.000  -0.3708   0.02039   0.01155  -0.0579   0.9769   0.2574
  -4.750  -0.3313   0.01933   0.01114  -0.0621   0.9753   0.3658
  -4.500  -0.3063   0.01953   0.01215  -0.0607   0.9717   0.4807
  -4.250  -0.2767   0.02014   0.01289  -0.0602   0.9683   0.5423
  -4.000  -0.2431   0.02038   0.01305  -0.0616   0.9648   0.5766
  -3.750  -0.2119   0.02066   0.01326  -0.0624   0.9608   0.5979
  -3.500  -0.1808   0.02109   0.01363  -0.0628   0.9574   0.6145
  -3.250  -0.1482   0.02159   0.01409  -0.0633   0.9544   0.6301
  -3.000  -0.1278   0.02231   0.01483  -0.0607   0.9491   0.6404
  -2.750  -0.1027   0.02284   0.01536  -0.0595   0.9441   0.6513
  -2.500  -0.0718   0.02329   0.01578  -0.0596   0.9407   0.6639
  -2.250  -0.0438   0.02389   0.01638  -0.0586   0.9377   0.6738
  -2.000  -0.0254   0.02427   0.01679  -0.0563   0.9308   0.6819
  -1.750   0.0041   0.02445   0.01695  -0.0564   0.9267   0.6879
  -1.500   0.0428   0.02435   0.01682  -0.0592   0.9240   0.6944
  -1.250   0.0688   0.02450   0.01699  -0.0585   0.9197   0.6976
  -1.000   0.0945   0.02457   0.01707  -0.0582   0.9136   0.7015
  -0.750   0.1339   0.02441   0.01690  -0.0609   0.9093   0.7064
  -0.500   0.1779   0.02417   0.01667  -0.0642   0.9060   0.7107
  -0.250   0.1999   0.02417   0.01671  -0.0628   0.8960   0.7131
   0.000   0.2402   0.02390   0.01648  -0.0650   0.8913   0.7154
   0.250   0.2721   0.02367   0.01629  -0.0657   0.8819   0.7180
   0.500   0.3257   0.02282   0.01546  -0.0696   0.8729   0.7210
   0.750   0.3684   0.02204   0.01470  -0.0720   0.8567   0.7247
   1.000   0.4001   0.02158   0.01431  -0.0719   0.8413   0.7268
   1.250   0.4316   0.02112   0.01389  -0.0716   0.8241   0.7288
   1.500   0.4653   0.02059   0.01339  -0.0717   0.8025   0.7308
   1.750   0.4935   0.02027   0.01312  -0.0712   0.7747   0.7329
   2.000   0.5264   0.01990   0.01275  -0.0715   0.7410   0.7352
   2.250   0.5620   0.01955   0.01226  -0.0722   0.6808   0.7379
   2.500   0.5981   0.01972   0.01159  -0.0729   0.5207   0.7411
   2.750   0.6148   0.02068   0.01182  -0.0707   0.3961   0.7431
   3.000   0.6325   0.02154   0.01218  -0.0690   0.3046   0.7449
   3.250   0.6532   0.02228   0.01256  -0.0680   0.2423   0.7468
   3.500   0.6763   0.02291   0.01294  -0.0675   0.2006   0.7489
   3.750   0.7008   0.02347   0.01334  -0.0672   0.1718   0.7511
   4.000   0.7264   0.02404   0.01378  -0.0672   0.1518   0.7536
   4.250   0.7533   0.02462   0.01428  -0.0675   0.1371   0.7563
   4.500   0.7794   0.02523   0.01482  -0.0676   0.1265   0.7590
   4.750   0.8010   0.02577   0.01535  -0.0665   0.1183   0.7609
   5.000   0.8232   0.02637   0.01596  -0.0656   0.1115   0.7631
   5.250   0.8464   0.02696   0.01655  -0.0649   0.1054   0.7655
   5.500   0.8703   0.02768   0.01724  -0.0645   0.1006   0.7679
   5.750   0.8958   0.02834   0.01794  -0.0644   0.0956   0.7707
   6.000   0.9215   0.02913   0.01866  -0.0645   0.0914   0.7738
   6.250   0.9468   0.02987   0.01949  -0.0643   0.0874   0.7763
   6.500   0.9693   0.03057   0.02023  -0.0635   0.0837   0.7782
   6.750   0.9918   0.03139   0.02100  -0.0629   0.0807   0.7803
   7.000   1.0165   0.03224   0.02200  -0.0625   0.0775   0.7826
   7.250   1.0409   0.03310   0.02293  -0.0622   0.0743   0.7851
   7.500   1.0654   0.03403   0.02384  -0.0621   0.0719   0.7879
   7.750   1.0917   0.03520   0.02508  -0.0623   0.0695   0.7911
   8.000   1.1141   0.03625   0.02634  -0.0616   0.0669   0.7935
   8.250   1.1345   0.03721   0.02740  -0.0606   0.0647   0.7959
   8.500   1.1556   0.03822   0.02839  -0.0600   0.0630   0.7985
   8.750   1.1774   0.03965   0.03002  -0.0594   0.0613   0.8011
   9.000   1.1981   0.04122   0.03185  -0.0587   0.0595   0.8039
   9.250   1.2181   0.04268   0.03349  -0.0581   0.0578   0.8067
   9.500   1.2365   0.04395   0.03485  -0.0573   0.0565   0.8094
   9.750   1.2530   0.04513   0.03604  -0.0560   0.0554   0.8118
  10.000   1.2631   0.04708   0.03835  -0.0539   0.0540   0.8147
  10.250   1.2709   0.04922   0.04085  -0.0518   0.0528   0.8178
  10.500   1.2777   0.05139   0.04330  -0.0497   0.0518   0.8210
  10.750   1.2805   0.05351   0.04566  -0.0472   0.0510   0.8242
  11.000   1.2814   0.05561   0.04799  -0.0446   0.0504   0.8267
  11.250   1.2818   0.05776   0.05033  -0.0422   0.0498   0.8293
  11.500   1.2825   0.05995   0.05267  -0.0401   0.0493   0.8320
  11.750   1.2841   0.06217   0.05502  -0.0385   0.0488   0.8349
  12.000   1.2768   0.06529   0.05835  -0.0367   0.0484   0.8377
  12.250   1.2516   0.07015   0.06363  -0.0349   0.0480   0.8400
  12.500   1.2224   0.07581   0.06969  -0.0343   0.0478   0.8419
  12.750   1.1886   0.08257   0.07681  -0.0353   0.0477   0.8435
  13.000   1.1485   0.09130   0.08589  -0.0388   0.0478   0.8447
  13.250   1.0977   0.10389   0.09882  -0.0468   0.0480   0.8454
<< Back to NASA SC(2)-0614 AIRFOIL (sc20614-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0614 AIRFOIL (sc20614-il)