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=9


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 50,000
Max Cl/Cd: 24.04 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20706-il-50000.txt
Download as CSV file: xf-sc20706-il-50000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5680   0.10208   0.09491   0.0040   1.0000   0.2911
  -8.750  -0.5704   0.09949   0.09239   0.0046   1.0000   0.3073
  -8.500  -0.5796   0.09752   0.09054   0.0054   1.0000   0.3230
  -8.250  -0.5551   0.09253   0.08550   0.0074   1.0000   0.3434
  -8.000  -0.5651   0.09087   0.08397   0.0090   1.0000   0.3643
  -7.750  -0.5451   0.08662   0.07969   0.0109   1.0000   0.3879
  -7.500  -0.5408   0.08408   0.07720   0.0136   1.0000   0.4182
  -7.250  -0.5458   0.08238   0.07561   0.0170   1.0000   0.4486
  -6.750  -0.5198   0.05112   0.04300  -0.0500   1.0000   0.1560
  -6.500  -0.4869   0.04437   0.03571  -0.0540   1.0000   0.1311
  -6.250  -0.4519   0.03960   0.03023  -0.0575   1.0000   0.1244
  -6.000  -0.4178   0.03563   0.02564  -0.0600   1.0000   0.1240
  -5.750  -0.3837   0.03206   0.02147  -0.0616   1.0000   0.1218
  -5.500  -0.3512   0.02917   0.01806  -0.0622   1.0000   0.1227
  -5.250  -0.3214   0.02685   0.01540  -0.0623   1.0000   0.1317
  -5.000  -0.2940   0.02492   0.01323  -0.0615   1.0000   0.1440
  -4.750  -0.2699   0.02317   0.01146  -0.0596   1.0000   0.1579
  -4.500  -0.2449   0.02148   0.00996  -0.0586   1.0000   0.1948
  -4.250  -0.2195   0.01765   0.00910  -0.0570   1.0000   0.6103
  -4.000  -0.2348   0.01873   0.01037  -0.0423   1.0000   0.7634
  -3.750  -0.2447   0.01897   0.01055  -0.0307   1.0000   0.8283
  -3.500  -0.2530   0.01848   0.00998  -0.0201   1.0000   0.8849
  -3.250  -0.1729   0.01710   0.00783  -0.0256   1.0000   1.0000
  -3.000  -0.1632   0.01663   0.00720  -0.0238   1.0000   1.0000
  -2.750  -0.1516   0.01621   0.00663  -0.0222   1.0000   1.0000
  -2.500  -0.1336   0.01590   0.00614  -0.0218   1.0000   1.0000
  -2.250  -0.1095   0.01568   0.00574  -0.0225   1.0000   1.0000
  -2.000  -0.0814   0.01555   0.00540  -0.0238   1.0000   1.0000
  -1.750  -0.0510   0.01548   0.00518  -0.0255   1.0000   1.0000
  -1.500  -0.0194   0.01546   0.00504  -0.0273   1.0000   1.0000
  -1.250   0.0129   0.01547   0.00497  -0.0292   1.0000   1.0000
  -1.000   0.0455   0.01553   0.00495  -0.0310   1.0000   1.0000
  -0.750   0.0780   0.01560   0.00500  -0.0328   1.0000   1.0000
  -0.500   0.1104   0.01571   0.00510  -0.0345   1.0000   1.0000
  -0.250   0.1426   0.01584   0.00527  -0.0361   1.0000   1.0000
   0.000   0.1745   0.01599   0.00550  -0.0377   1.0000   1.0000
   0.250   0.2060   0.01616   0.00577  -0.0391   1.0000   1.0000
   0.500   0.2372   0.01635   0.00610  -0.0404   1.0000   1.0000
   0.750   0.2679   0.01657   0.00649  -0.0416   1.0000   1.0000
   1.000   0.2983   0.01682   0.00698  -0.0427   1.0000   1.0000
   1.250   0.3283   0.01709   0.00750  -0.0437   1.0000   1.0000
   1.500   0.3579   0.01739   0.00810  -0.0446   1.0000   1.0000
   1.750   0.3871   0.01772   0.00880  -0.0454   1.0000   1.0000
   2.000   0.4160   0.01810   0.00963  -0.0462   1.0000   1.0000
   2.250   0.4449   0.01851   0.01074  -0.0467   1.0000   1.0000
   2.500   0.5760   0.02453   0.01309  -0.0578   0.1425   1.0000
   2.750   0.6150   0.02704   0.01573  -0.0587   0.1310   1.0000
   3.000   0.6481   0.02952   0.01847  -0.0591   0.1194   1.0000
   3.250   0.6806   0.03217   0.02157  -0.0590   0.1153   1.0000
   3.500   0.7120   0.03537   0.02533  -0.0586   0.1169   1.0000
   3.750   0.7409   0.03905   0.02951  -0.0582   0.1209   1.0000
   4.000   0.7675   0.04313   0.03390  -0.0578   0.1230   1.0000
   4.250   0.7929   0.04696   0.03879  -0.0567   0.1334   1.0000
   4.500   0.8191   0.05257   0.04527  -0.0567   0.1613   1.0000
   5.750   0.8506   0.09100   0.08600  -0.1010   0.4286   1.0000
   6.000   0.8739   0.09505   0.09007  -0.0959   0.3881   1.0000
   6.250   0.8723   0.09872   0.09368  -0.0940   0.3593   1.0000
   6.500   0.8812   0.10287   0.09781  -0.0915   0.3318   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)