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)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il)
Reynolds number: 100,000
Max Cl/Cd: 31.82 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20610-il-100000-n5.txt
Download as CSV file: xf-sc20610-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0610 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.6866   0.08095   0.07508  -0.0433   1.0000   0.0345
 -11.500  -0.7120   0.07277   0.06681  -0.0489   1.0000   0.0343
 -11.250  -0.7373   0.06644   0.06036  -0.0523   1.0000   0.0341
 -11.000  -0.7620   0.06162   0.05541  -0.0533   1.0000   0.0340
 -10.750  -0.7861   0.05765   0.05129  -0.0528   1.0000   0.0340
 -10.500  -0.8036   0.05317   0.04650  -0.0532   1.0000   0.0342
 -10.250  -0.8147   0.04861   0.04144  -0.0532   1.0000   0.0347
 -10.000  -0.8065   0.04634   0.03911  -0.0528   1.0000   0.0353
  -9.750  -0.7960   0.04428   0.03691  -0.0524   1.0000   0.0361
  -9.500  -0.7845   0.04191   0.03431  -0.0521   1.0000   0.0369
  -9.250  -0.7714   0.03918   0.03126  -0.0519   1.0000   0.0377
  -9.000  -0.7556   0.03652   0.02824  -0.0517   1.0000   0.0386
  -8.750  -0.7375   0.03404   0.02538  -0.0515   1.0000   0.0397
  -8.500  -0.7175   0.03177   0.02271  -0.0512   1.0000   0.0411
  -8.250  -0.6965   0.03009   0.02084  -0.0509   1.0000   0.0429
  -8.000  -0.6749   0.02903   0.01974  -0.0507   1.0000   0.0450
  -7.750  -0.6524   0.02780   0.01835  -0.0504   1.0000   0.0475
  -7.500  -0.6295   0.02646   0.01677  -0.0498   1.0000   0.0499
  -7.250  -0.6069   0.02524   0.01551  -0.0494   1.0000   0.0525
  -7.000  -0.5835   0.02441   0.01468  -0.0492   1.0000   0.0562
  -6.750  -0.5596   0.02353   0.01365  -0.0488   1.0000   0.0603
  -6.500  -0.5357   0.02257   0.01274  -0.0486   1.0000   0.0644
  -6.250  -0.5110   0.02186   0.01199  -0.0485   1.0000   0.0699
  -6.000  -0.4857   0.02105   0.01117  -0.0485   1.0000   0.0754
  -5.750  -0.4598   0.02041   0.01054  -0.0487   1.0000   0.0826
  -5.500  -0.4330   0.01972   0.00987  -0.0490   1.0000   0.0901
  -5.250  -0.4059   0.01919   0.00930  -0.0494   1.0000   0.1006
  -5.000  -0.3775   0.01855   0.00874  -0.0501   1.0000   0.1135
  -4.750  -0.3482   0.01793   0.00827  -0.0511   1.0000   0.1340
  -4.500  -0.3174   0.01726   0.00783  -0.0525   1.0000   0.1753
  -4.250  -0.2807   0.01614   0.00736  -0.0558   1.0000   0.2907
  -4.000  -0.2493   0.01527   0.00769  -0.0571   1.0000   0.5067
  -3.750  -0.2309   0.01566   0.00832  -0.0543   1.0000   0.5902
  -3.500  -0.2058   0.01596   0.00857  -0.0535   1.0000   0.6304
  -3.250  -0.1811   0.01629   0.00884  -0.0527   1.0000   0.6580
  -3.000  -0.1616   0.01677   0.00931  -0.0504   1.0000   0.6779
  -2.750  -0.1432   0.01728   0.00980  -0.0478   1.0000   0.6962
  -2.500  -0.1260   0.01781   0.01033  -0.0449   1.0000   0.7135
  -2.250  -0.1090   0.01827   0.01080  -0.0421   1.0000   0.7288
  -2.000  -0.0893   0.01863   0.01115  -0.0401   1.0000   0.7424
  -1.750  -0.0677   0.01886   0.01136  -0.0388   1.0000   0.7521
  -1.500  -0.0438   0.01899   0.01147  -0.0381   1.0000   0.7588
  -1.250  -0.0202   0.01910   0.01158  -0.0375   1.0000   0.7643
  -1.000   0.0151   0.01919   0.01162  -0.0397   0.9976   0.7707
  -0.750   0.0489   0.01929   0.01174  -0.0410   0.9940   0.7746
  -0.500   0.0821   0.01934   0.01181  -0.0424   0.9892   0.7781
  -0.250   0.1210   0.01941   0.01190  -0.0450   0.9846   0.7816
   0.000   0.1597   0.01943   0.01195  -0.0477   0.9786   0.7853
   0.250   0.2028   0.01942   0.01199  -0.0511   0.9721   0.7881
   0.500   0.2406   0.01930   0.01195  -0.0531   0.9616   0.7905
   0.750   0.2854   0.01900   0.01175  -0.0561   0.9473   0.7929
   1.000   0.3392   0.01836   0.01121  -0.0603   0.9271   0.7952
   1.250   0.3868   0.01761   0.01057  -0.0630   0.9011   0.7977
   1.500   0.4276   0.01699   0.01003  -0.0645   0.8674   0.8006
   1.750   0.4614   0.01657   0.00968  -0.0647   0.8213   0.8033
   2.000   0.5113   0.01607   0.00879  -0.0668   0.6712   0.8050
   2.250   0.5332   0.01734   0.00875  -0.0649   0.4180   0.8071
   2.500   0.5528   0.01857   0.00915  -0.0637   0.2596   0.8099
   3.000   0.6036   0.02023   0.01009  -0.0636   0.1339   0.8161
   3.250   0.6288   0.02088   0.01062  -0.0634   0.1140   0.8188
   3.500   0.6527   0.02144   0.01118  -0.0626   0.1008   0.8214
   3.750   0.6773   0.02205   0.01180  -0.0621   0.0908   0.8242
   4.000   0.7023   0.02280   0.01250  -0.0618   0.0828   0.8269
   4.250   0.7293   0.02346   0.01325  -0.0618   0.0755   0.8297
   4.500   0.7560   0.02438   0.01409  -0.0620   0.0694   0.8325
   4.750   0.7809   0.02511   0.01497  -0.0614   0.0641   0.8350
   5.000   0.8047   0.02591   0.01575  -0.0608   0.0595   0.8378
   5.250   0.8303   0.02698   0.01694  -0.0604   0.0558   0.8408
   5.500   0.8568   0.02791   0.01799  -0.0603   0.0517   0.8438
   5.750   0.8831   0.02904   0.01906  -0.0606   0.0486   0.8467
   6.000   0.9097   0.03050   0.02079  -0.0603   0.0464   0.8493
   6.250   0.9341   0.03195   0.02249  -0.0596   0.0441   0.8519
   6.500   0.9579   0.03316   0.02384  -0.0592   0.0419   0.8547
   6.750   0.9813   0.03441   0.02513  -0.0589   0.0401   0.8579
   7.000   1.0042   0.03659   0.02775  -0.0583   0.0385   0.8612
   7.250   1.0241   0.03909   0.03071  -0.0572   0.0373   0.8644
   7.500   1.0404   0.04167   0.03373  -0.0556   0.0364   0.8676
   7.750   1.0548   0.04448   0.03697  -0.0540   0.0356   0.8711
   8.000   1.0672   0.04747   0.04036  -0.0524   0.0350   0.8746
   8.250   1.0798   0.05010   0.04329  -0.0512   0.0343   0.8780
   8.500   1.0899   0.05225   0.04565  -0.0495   0.0336   0.8812
   8.750   1.0995   0.05451   0.04804  -0.0481   0.0330   0.8850
   9.000   1.0885   0.05954   0.05370  -0.0449   0.0325   0.8888
   9.250   1.0722   0.06473   0.05939  -0.0421   0.0321   0.8927
   9.500   1.0502   0.06922   0.06423  -0.0389   0.0321   0.8967
   9.750   1.0255   0.07415   0.06942  -0.0370   0.0321   0.9013
  10.000   1.0000   0.08002   0.07553  -0.0376   0.0322   0.9053
  10.250   0.9723   0.08735   0.08306  -0.0410   0.0324   0.9089
  10.500   0.9469   0.09680   0.09265  -0.0479   0.0326   0.9123
  10.750   0.9258   0.10907   0.10498  -0.0570   0.0329   0.9152
<< Back to NASA SC(2)-0610 AIRFOIL (sc20610-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0610 AIRFOIL (sc20610-il)