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)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 50,000
Max Cl/Cd: 28.31 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc21006-il-50000.txt
Download as CSV file: xf-sc21006-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-1006 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4974   0.09686   0.08995  -0.0089   0.9999   0.2713
  -9.750  -0.5045   0.09472   0.08793  -0.0074   0.9999   0.2820
  -9.500  -0.5118   0.09275   0.08608  -0.0062   0.9999   0.2943
  -9.000  -0.4990   0.08636   0.07976  -0.0025   0.9999   0.3260
  -8.750  -0.5020   0.08433   0.07782  -0.0014   0.9999   0.3473
  -8.500  -0.4914   0.08088   0.07439   0.0016   0.9999   0.3693
  -8.250  -0.4944   0.07868   0.07229   0.0028   0.9999   0.3923
  -8.000  -0.4832   0.07526   0.06888   0.0055   0.9999   0.4114
  -7.750  -0.3774   0.04469   0.03658  -0.0718   0.9999   0.1368
  -7.500  -0.3248   0.03783   0.02863  -0.0809   0.9999   0.1239
  -7.250  -0.2840   0.03341   0.02359  -0.0849   0.9999   0.1205
  -7.000  -0.2440   0.03000   0.01947  -0.0878   0.9999   0.1220
  -6.750  -0.2098   0.02760   0.01668  -0.0894   0.9999   0.1345
  -6.500  -0.1770   0.02535   0.01413  -0.0900   0.9999   0.1451
  -6.250  -0.1468   0.02353   0.01221  -0.0900   0.9999   0.1692
  -6.000  -0.1127   0.02140   0.01044  -0.0909   0.9999   0.2295
  -5.750  -0.0849   0.01941   0.01070  -0.0885   0.9999   0.6135
  -5.500  -0.0845   0.01991   0.01135  -0.0797   0.9999   0.7019
  -5.250  -0.0854   0.02006   0.01146  -0.0710   0.9999   0.7629
  -5.000  -0.0843   0.01975   0.01109  -0.0634   0.9999   0.8184
  -4.750  -0.0828   0.01897   0.01024  -0.0565   0.9999   0.8754
  -4.500  -0.0814   0.01755   0.00878  -0.0502   0.9999   0.9457
  -4.250  -0.0512   0.01679   0.00778  -0.0523   0.9999   1.0001
  -4.000  -0.0031   0.01683   0.00740  -0.0586   0.9999   1.0001
  -3.750   0.0402   0.01693   0.00714  -0.0634   0.9999   1.0001
  -3.500   0.0802   0.01706   0.00701  -0.0672   0.9999   1.0001
  -3.250   0.1176   0.01723   0.00698  -0.0703   0.9999   1.0001
  -3.000   0.1531   0.01741   0.00699  -0.0728   0.9999   1.0001
  -2.750   0.1871   0.01762   0.00708  -0.0749   0.9999   1.0001
  -2.500   0.2198   0.01784   0.00723  -0.0766   0.9999   1.0001
  -2.250   0.2516   0.01809   0.00743  -0.0781   0.9999   1.0001
  -2.000   0.2825   0.01835   0.00770  -0.0794   0.9999   1.0001
  -1.750   0.3126   0.01863   0.00801  -0.0804   0.9999   1.0001
  -1.500   0.3421   0.01894   0.00836  -0.0813   0.9999   1.0001
  -1.250   0.3710   0.01927   0.00878  -0.0821   0.9999   1.0001
  -1.000   0.3994   0.01962   0.00929  -0.0828   0.9999   1.0001
  -0.750   0.4273   0.02000   0.00983  -0.0834   0.9999   1.0001
  -0.500   0.4548   0.02040   0.01043  -0.0839   0.9999   1.0001
  -0.250   0.4818   0.02085   0.01110  -0.0843   0.9999   1.0001
   0.000   0.5084   0.02133   0.01185  -0.0847   0.9999   1.0001
   0.250   0.5346   0.02185   0.01275  -0.0850   0.9999   1.0001
   0.500   0.5603   0.02243   0.01370  -0.0852   0.9999   1.0001
   0.750   0.5856   0.02306   0.01480  -0.0854   0.9999   1.0001
   1.000   0.7408   0.02617   0.01452  -0.0965   0.1389   1.0001
   1.250   0.7811   0.02841   0.01686  -0.0975   0.1188   1.0001
   1.500   0.8197   0.03146   0.01989  -0.0984   0.1123   1.0001
   1.750   0.8536   0.03433   0.02319  -0.0981   0.1104   1.0001
   2.000   0.8834   0.03705   0.02650  -0.0972   0.1069   1.0001
   2.250   0.9115   0.04031   0.03036  -0.0962   0.1063   1.0001
   2.500   0.9380   0.04442   0.03488  -0.0953   0.1104   1.0001
   2.750   0.9628   0.04852   0.03992  -0.0934   0.1218   1.0001
   3.000   0.9871   0.05386   0.04582  -0.0921   0.1382   1.0001
   4.750   0.9441   0.10539   0.10042  -0.1297   0.4284   1.0001
   5.000   0.9468   0.10913   0.10413  -0.1271   0.3974   1.0001
<< Back to NASA SC(2)-1006 AIRFOIL (sc21006-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-1006 AIRFOIL (sc21006-il)