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)-0410 AIRFOIL (sc20410-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0410 AIRFOIL (sc20410-il)
Reynolds number: 200,000
Max Cl/Cd: 35.26 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20410-il-200000.txt
Download as CSV file: xf-sc20410-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0410 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5223   0.09066   0.08693  -0.0229   1.0000   0.0776
 -10.000  -0.5291   0.08539   0.08167  -0.0247   1.0000   0.0792
  -9.750  -0.5435   0.07861   0.07493  -0.0279   1.0000   0.0806
  -9.500  -0.5785   0.06802   0.06436  -0.0350   1.0000   0.0812
  -9.250  -0.8131   0.05140   0.04630  -0.0424   1.0000   0.0516
  -9.000  -0.7952   0.04988   0.04488  -0.0420   1.0000   0.0530
  -8.750  -0.8120   0.03657   0.03007  -0.0418   1.0000   0.0428
  -8.500  -0.7948   0.03340   0.02655  -0.0414   1.0000   0.0429
  -8.250  -0.7750   0.03099   0.02391  -0.0411   1.0000   0.0435
  -8.000  -0.7535   0.02908   0.02184  -0.0408   1.0000   0.0444
  -7.750  -0.7310   0.02745   0.02003  -0.0404   1.0000   0.0458
  -7.500  -0.7077   0.02603   0.01838  -0.0401   1.0000   0.0479
  -7.250  -0.6833   0.02479   0.01684  -0.0396   1.0000   0.0499
  -7.000  -0.6610   0.02256   0.01451  -0.0390   1.0000   0.0520
  -6.750  -0.6373   0.02151   0.01347  -0.0385   1.0000   0.0548
  -6.500  -0.6130   0.02058   0.01244  -0.0380   1.0000   0.0583
  -6.250  -0.5893   0.01940   0.01118  -0.0374   1.0000   0.0626
  -6.000  -0.5650   0.01868   0.01052  -0.0371   1.0000   0.0676
  -5.750  -0.5399   0.01817   0.00990  -0.0367   1.0000   0.0724
  -5.500  -0.5161   0.01704   0.00889  -0.0364   1.0000   0.0786
  -5.250  -0.4907   0.01658   0.00839  -0.0361   1.0000   0.0850
  -5.000  -0.4652   0.01565   0.00756  -0.0362   1.0000   0.0921
  -4.750  -0.4390   0.01517   0.00704  -0.0361   1.0000   0.1000
  -4.500  -0.4111   0.01435   0.00635  -0.0367   1.0000   0.1109
  -4.250  -0.3823   0.01369   0.00578  -0.0375   1.0000   0.1277
  -4.000  -0.3499   0.01270   0.00512  -0.0394   1.0000   0.1774
  -3.750  -0.3051   0.01035   0.00489  -0.0451   1.0000   0.5769
  -3.500  -0.2810   0.01070   0.00531  -0.0438   1.0000   0.6450
  -3.250  -0.2572   0.01108   0.00566  -0.0425   1.0000   0.6743
  -3.000  -0.2359   0.01151   0.00608  -0.0406   1.0000   0.6928
  -2.750  -0.2147   0.01191   0.00649  -0.0387   1.0000   0.7083
  -2.500  -0.1922   0.01228   0.00682  -0.0373   1.0000   0.7223
  -2.250  -0.1751   0.01270   0.00729  -0.0344   1.0000   0.7308
  -2.000  -0.1535   0.01303   0.00762  -0.0329   1.0000   0.7419
  -1.750  -0.1340   0.01350   0.00809  -0.0308   1.0000   0.7557
  -1.500  -0.1232   0.01405   0.00873  -0.0263   1.0000   0.7664
  -1.250  -0.1081   0.01447   0.00918  -0.0232   1.0000   0.7776
  -1.000  -0.0864   0.01475   0.00945  -0.0220   1.0000   0.7884
  -0.750  -0.0711   0.01502   0.00977  -0.0192   1.0000   0.7956
  -0.500  -0.0463   0.01518   0.00993  -0.0190   1.0000   0.8035
  -0.250  -0.0175   0.01533   0.01011  -0.0191   0.9964   0.8091
   0.000   0.0302   0.01536   0.01014  -0.0234   0.9893   0.8155
   0.250   0.0764   0.01526   0.01006  -0.0272   0.9821   0.8184
   0.500   0.1257   0.01499   0.00983  -0.0314   0.9720   0.8198
   0.750   0.1730   0.01466   0.00954  -0.0350   0.9605   0.8216
   1.000   0.2235   0.01421   0.00915  -0.0389   0.9483   0.8236
   1.250   0.2718   0.01368   0.00869  -0.0422   0.9337   0.8260
   1.500   0.3132   0.01325   0.00831  -0.0443   0.9166   0.8287
   1.750   0.3509   0.01292   0.00801  -0.0459   0.8942   0.8310
   2.000   0.3808   0.01262   0.00773  -0.0454   0.8643   0.8329
   2.250   0.4072   0.01242   0.00749  -0.0441   0.8205   0.8348
   2.500   0.4330   0.01238   0.00727  -0.0428   0.7489   0.8366
   2.750   0.4542   0.01288   0.00710  -0.0407   0.5843   0.8386
   3.000   0.4693   0.01450   0.00739  -0.0387   0.3195   0.8411
   3.250   0.4917   0.01590   0.00790  -0.0386   0.1652   0.8439
   3.500   0.5205   0.01667   0.00842  -0.0394   0.1282   0.8462
   3.750   0.5456   0.01719   0.00888  -0.0389   0.1124   0.8480
   4.000   0.5704   0.01772   0.00938  -0.0384   0.1016   0.8503
   4.250   0.5951   0.01848   0.01003  -0.0379   0.0931   0.8527
   4.500   0.6231   0.01886   0.01050  -0.0380   0.0858   0.8552
   4.750   0.6502   0.01979   0.01134  -0.0382   0.0789   0.8576
   5.000   0.6799   0.02026   0.01188  -0.0388   0.0724   0.8600
   5.250   0.7070   0.02142   0.01296  -0.0391   0.0665   0.8622
   5.500   0.7330   0.02200   0.01370  -0.0385   0.0622   0.8644
   5.750   0.7586   0.02278   0.01448  -0.0382   0.0580   0.8672
   6.000   0.7846   0.02422   0.01603  -0.0380   0.0542   0.8702
   6.250   0.8124   0.02516   0.01714  -0.0380   0.0508   0.8733
   6.500   0.8408   0.02636   0.01842  -0.0384   0.0485   0.8761
   6.750   0.8647   0.02828   0.02041  -0.0380   0.0467   0.8786
   7.000   0.8870   0.03062   0.02311  -0.0371   0.0457   0.8814
   7.250   0.9090   0.03275   0.02569  -0.0361   0.0450   0.8845
   7.500   0.9300   0.03528   0.02867  -0.0351   0.0440   0.8879
   7.750   0.9490   0.03830   0.03216  -0.0343   0.0430   0.8912
   8.000   0.9595   0.04227   0.03663  -0.0321   0.0434   0.8943
   8.250   0.9669   0.04687   0.04169  -0.0301   0.0443   0.8980
   8.500   0.9735   0.05167   0.04679  -0.0288   0.0454   0.9021
   9.250   0.9010   0.08116   0.07804  -0.0249   0.0691   0.9141
   9.500   0.9089   0.08454   0.08141  -0.0244   0.0683   0.9184
   9.750   0.9141   0.09294   0.08968  -0.0252   0.0671   0.9217
  10.000   0.8888   0.09700   0.09392  -0.0249   0.0670   0.9279
  10.250   0.8620   0.10272   0.09978  -0.0284   0.0669   0.9341
  10.500   0.8306   0.11267   0.10984  -0.0391   0.0668   0.9403
<< Back to NASA SC(2)-0410 AIRFOIL (sc20410-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0410 AIRFOIL (sc20410-il)