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)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 50,000
Max Cl/Cd: 25.29 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20706-il-50000-n5.txt
Download as CSV file: xf-sc20706-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0706 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4817   0.10788   0.10088  -0.0133   1.0000   0.1244
 -10.250  -0.5012   0.09868   0.09173  -0.0207   1.0000   0.0657
  -9.750  -0.6046   0.09807   0.09080  -0.0209   1.0000   0.0512
  -9.500  -0.6040   0.09385   0.08663  -0.0218   1.0000   0.0509
  -9.250  -0.6036   0.08942   0.08225  -0.0234   1.0000   0.0506
  -9.000  -0.6027   0.08450   0.07737  -0.0266   1.0000   0.0502
  -8.750  -0.6014   0.07960   0.07248  -0.0298   1.0000   0.0497
  -8.500  -0.5993   0.07467   0.06751  -0.0329   1.0000   0.0492
  -8.250  -0.5952   0.06965   0.06241  -0.0360   1.0000   0.0486
  -8.000  -0.5883   0.06450   0.05712  -0.0391   1.0000   0.0477
  -7.750  -0.5775   0.05920   0.05158  -0.0425   1.0000   0.0468
  -7.500  -0.5621   0.05385   0.04588  -0.0459   1.0000   0.0459
  -7.250  -0.5416   0.04865   0.04020  -0.0492   1.0000   0.0453
  -7.000  -0.5169   0.04387   0.03487  -0.0521   1.0000   0.0451
  -6.750  -0.4892   0.03961   0.03001  -0.0545   1.0000   0.0453
  -6.500  -0.4600   0.03592   0.02565  -0.0562   1.0000   0.0461
  -6.250  -0.4301   0.03303   0.02212  -0.0574   1.0000   0.0494
  -6.000  -0.4020   0.03048   0.01916  -0.0580   1.0000   0.0535
  -5.750  -0.3755   0.02839   0.01683  -0.0579   1.0000   0.0566
  -5.500  -0.3498   0.02657   0.01475  -0.0570   1.0000   0.0601
  -5.250  -0.3250   0.02505   0.01303  -0.0560   1.0000   0.0667
  -5.000  -0.2998   0.02383   0.01158  -0.0553   1.0000   0.0777
  -4.750  -0.2741   0.02243   0.01016  -0.0549   1.0000   0.0876
  -4.500  -0.2447   0.02108   0.00886  -0.0558   1.0000   0.1134
  -4.250  -0.2043   0.01842   0.00729  -0.0605   1.0000   0.2780
  -4.000  -0.1903   0.01753   0.00797  -0.0562   1.0000   0.6256
  -3.750  -0.1784   0.01784   0.00831  -0.0510   1.0000   0.7153
  -3.500  -0.1736   0.01812   0.00856  -0.0439   1.0000   0.7755
  -3.250  -0.1677   0.01812   0.00851  -0.0374   1.0000   0.8203
  -3.000  -0.1562   0.01788   0.00814  -0.0330   1.0000   0.8538
  -2.750  -0.1409   0.01749   0.00762  -0.0298   1.0000   0.8807
  -2.500  -0.1214   0.01706   0.00704  -0.0279   1.0000   0.9026
  -2.250  -0.0986   0.01663   0.00648  -0.0270   1.0000   0.9223
  -2.000  -0.0735   0.01622   0.00591  -0.0266   1.0000   0.9421
  -1.750  -0.0446   0.01585   0.00545  -0.0273   1.0000   0.9629
  -1.500  -0.0177   0.01552   0.00509  -0.0277   1.0000   0.9931
  -1.250   0.0127   0.01548   0.00497  -0.0291   1.0000   1.0000
  -1.000   0.0452   0.01553   0.00495  -0.0310   1.0000   1.0000
  -0.750   0.0778   0.01561   0.00500  -0.0327   1.0000   1.0000
  -0.500   0.1102   0.01571   0.00510  -0.0345   1.0000   1.0000
  -0.250   0.1424   0.01584   0.00527  -0.0361   1.0000   1.0000
   0.000   0.1743   0.01599   0.00550  -0.0376   1.0000   1.0000
   0.250   0.2058   0.01616   0.00577  -0.0390   1.0000   1.0000
   0.500   0.2370   0.01635   0.00609  -0.0403   1.0000   1.0000
   0.750   0.2678   0.01657   0.00648  -0.0415   1.0000   1.0000
   1.000   0.2981   0.01682   0.00698  -0.0426   1.0000   1.0000
   1.250   0.3281   0.01709   0.00750  -0.0436   1.0000   1.0000
   1.500   0.3577   0.01739   0.00810  -0.0446   1.0000   1.0000
   1.750   0.3870   0.01772   0.00880  -0.0454   1.0000   1.0000
   2.000   0.5013   0.01982   0.00808  -0.0574   0.1595   1.0000
   2.250   0.5299   0.02173   0.00957  -0.0578   0.1030   1.0000
   2.500   0.5607   0.02315   0.01106  -0.0584   0.0806   1.0000
   2.750   0.5937   0.02483   0.01290  -0.0590   0.0714   1.0000
   3.000   0.6254   0.02657   0.01475  -0.0597   0.0607   1.0000
   3.250   0.6581   0.02849   0.01701  -0.0599   0.0551   1.0000
   3.500   0.6895   0.03080   0.01959  -0.0601   0.0522   1.0000
   3.750   0.7183   0.03343   0.02243  -0.0603   0.0493   1.0000
   4.000   0.7461   0.03609   0.02567  -0.0599   0.0459   1.0000
   4.250   0.7722   0.03903   0.02922  -0.0593   0.0434   1.0000
   4.500   0.7960   0.04252   0.03336  -0.0585   0.0429   1.0000
   4.750   0.8173   0.04633   0.03773  -0.0576   0.0429   1.0000
   5.000   0.8359   0.05046   0.04241  -0.0567   0.0431   1.0000
   5.250   0.8518   0.05489   0.04733  -0.0559   0.0436   1.0000
   5.500   0.8652   0.05954   0.05242  -0.0552   0.0442   1.0000
   5.750   0.8762   0.06435   0.05760  -0.0547   0.0451   1.0000
   6.000   0.8849   0.06926   0.06279  -0.0545   0.0460   1.0000
   6.250   0.8915   0.07423   0.06799  -0.0545   0.0467   1.0000
   6.500   0.8960   0.07923   0.07319  -0.0547   0.0472   1.0000
   6.750   0.8985   0.08428   0.07840  -0.0551   0.0475   1.0000
   7.000   0.8991   0.08936   0.08361  -0.0558   0.0477   1.0000
   7.250   0.8988   0.09443   0.08877  -0.0566   0.0480   1.0000
   7.500   0.8989   0.09953   0.09392  -0.0574   0.0483   1.0000
   8.000   0.7336   0.10502   0.09977  -0.0545   0.0578   1.0000
   8.250   0.7388   0.10958   0.10432  -0.0537   0.0608   1.0000
<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)