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=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 100,000
Max Cl/Cd: 48.87 at α=-0.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc21006-il-100000-n5.txt
Download as CSV file: xf-sc21006-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5094   0.09765   0.09252  -0.0246   0.9999   0.0343
 -11.000  -0.5120   0.09414   0.08907  -0.0246   0.9999   0.0322
 -10.750  -0.5184   0.09054   0.08555  -0.0247   0.9999   0.0306
 -10.500  -0.5300   0.08667   0.08178  -0.0248   0.9999   0.0289
 -10.250  -0.5338   0.07998   0.07516  -0.0310   0.9999   0.0260
 -10.000  -0.5342   0.07563   0.07085  -0.0328   0.9999   0.0255
  -9.750  -0.5301   0.06993   0.06516  -0.0382   0.9999   0.0251
  -9.500  -0.5215   0.06324   0.05839  -0.0451   0.9999   0.0247
  -9.250  -0.5076   0.05633   0.05130  -0.0521   0.9999   0.0243
  -9.000  -0.4856   0.04910   0.04373  -0.0598   0.9999   0.0240
  -8.750  -0.4533   0.04195   0.03596  -0.0679   0.9999   0.0247
  -8.500  -0.4142   0.03626   0.02935  -0.0749   0.9999   0.0265
  -8.250  -0.3772   0.03145   0.02389  -0.0799   0.9999   0.0269
  -8.000  -0.3408   0.02785   0.01970  -0.0834   0.9999   0.0273
  -7.750  -0.3062   0.02516   0.01657  -0.0858   0.9999   0.0280
  -7.500  -0.2731   0.02315   0.01424  -0.0875   0.9999   0.0292
  -7.250  -0.2408   0.02154   0.01239  -0.0888   0.9999   0.0310
  -7.000  -0.2097   0.02043   0.01102  -0.0898   0.9999   0.0353
  -6.750  -0.1766   0.01916   0.00963  -0.0914   0.9999   0.0401
  -6.500  -0.1424   0.01816   0.00854  -0.0931   0.9999   0.0458
  -6.250  -0.1081   0.01737   0.00769  -0.0949   0.9999   0.0593
  -6.000  -0.0708   0.01646   0.00682  -0.0974   0.9999   0.0903
  -5.750  -0.0296   0.01516   0.00627  -0.1015   0.9999   0.2529
  -5.500   0.0061   0.01448   0.00627  -0.1039   0.9999   0.4414
  -5.250   0.0350   0.01430   0.00634  -0.1041   0.9999   0.5361
  -5.000   0.0618   0.01427   0.00639  -0.1038   0.9999   0.5906
  -4.750   0.0875   0.01430   0.00643  -0.1032   0.9999   0.6333
  -4.500   0.1124   0.01437   0.00653  -0.1024   0.9999   0.6732
  -4.250   0.1364   0.01446   0.00665  -0.1014   0.9999   0.7079
  -4.000   0.1605   0.01456   0.00675  -0.1005   0.9999   0.7359
  -3.750   0.1849   0.01465   0.00684  -0.0997   0.9999   0.7597
  -3.500   0.2095   0.01475   0.00692  -0.0990   0.9999   0.7814
  -3.250   0.2335   0.01483   0.00702  -0.0981   0.9999   0.8004
  -3.000   0.2578   0.01492   0.00713  -0.0974   0.9999   0.8194
  -2.750   0.2821   0.01500   0.00725  -0.0968   0.9999   0.8384
  -2.500   0.3053   0.01506   0.00738  -0.0958   0.9999   0.8568
  -2.250   0.3286   0.01512   0.00751  -0.0950   0.9999   0.8770
  -2.000   0.3503   0.01514   0.00761  -0.0938   0.9999   0.8994
  -1.750   0.3702   0.01512   0.00771  -0.0923   0.9999   0.9281
  -1.500   0.3840   0.01520   0.00790  -0.0899   0.9999   1.0001
  -1.250   0.4141   0.01551   0.00832  -0.0910   0.9999   1.0001
  -1.000   0.4435   0.01584   0.00876  -0.0918   0.9999   1.0001
  -0.750   0.4813   0.01613   0.00919  -0.0944   0.9953   1.0001
  -0.500   0.5386   0.01606   0.00937  -0.1004   0.9773   1.0001
  -0.250   0.6583   0.01347   0.00531  -0.1114   0.4578   1.0001
   0.000   0.6767   0.01625   0.00624  -0.1104   0.1051   1.0001
   0.250   0.7023   0.01755   0.00720  -0.1101   0.0567   1.0001
   0.500   0.7286   0.01856   0.00835  -0.1096   0.0469   1.0001
   0.750   0.7538   0.01971   0.00953  -0.1091   0.0392   1.0001
   1.000   0.7796   0.02079   0.01072  -0.1086   0.0343   1.0001
   1.250   0.8052   0.02219   0.01215  -0.1081   0.0317   1.0001
   1.500   0.8317   0.02411   0.01409  -0.1077   0.0299   1.0001
   1.750   0.8604   0.02645   0.01657  -0.1075   0.0288   1.0001
   2.000   0.8890   0.02804   0.01856  -0.1070   0.0268   1.0001
   2.250   0.9165   0.02994   0.02081  -0.1065   0.0243   1.0001
   2.500   0.9429   0.03258   0.02388  -0.1057   0.0236   1.0001
   2.750   0.9673   0.03559   0.02737  -0.1044   0.0232   1.0001
   3.000   0.9892   0.03899   0.03132  -0.1028   0.0230   1.0001
   3.250   1.0083   0.04285   0.03573  -0.1009   0.0232   1.0001
   3.500   1.0244   0.04711   0.04052  -0.0987   0.0235   1.0001
   3.750   1.0375   0.05172   0.04563  -0.0964   0.0240   1.0001
   4.000   1.0474   0.05661   0.05097  -0.0941   0.0246   1.0001
   4.250   1.0549   0.06147   0.05622  -0.0920   0.0248   1.0001
   4.500   1.0599   0.06625   0.06142  -0.0900   0.0242   1.0001
   4.750   1.0623   0.07094   0.06640  -0.0883   0.0234   1.0001
   5.000   1.0619   0.07563   0.07132  -0.0869   0.0227   1.0001
   5.250   1.0572   0.08075   0.07666  -0.0857   0.0226   1.0001
   5.500   1.0480   0.08621   0.08230  -0.0849   0.0231   1.0001
   5.750   1.0342   0.09142   0.08763  -0.0844   0.0235   1.0001
   6.000   1.0186   0.09681   0.09311  -0.0849   0.0240   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)