Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)
Reynolds number: 200,000
Max Cl/Cd: 76.32 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nlf0215f-il-200000.txt
Download as CSV file: xf-nlf0215f-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY NLF(1)-0215F 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2699   0.09969   0.09642  -0.0625   0.9880   0.0507
  -9.250  -0.2595   0.09344   0.09017  -0.0714   0.9820   0.0536
  -9.000  -0.2769   0.08197   0.07865  -0.0922   0.9658   0.0549
  -8.500  -0.2377   0.07345   0.07021  -0.0950   0.9532   0.0572
  -8.250  -0.2218   0.06974   0.06648  -0.0985   0.9435   0.0589
  -8.000  -0.2186   0.06417   0.06084  -0.1053   0.9310   0.0609
  -7.750  -0.2136   0.05757   0.05400  -0.1154   0.9176   0.0638
  -7.500  -0.2049   0.05032   0.04622  -0.1253   0.9069   0.0682
  -7.250  -0.1761   0.04730   0.04324  -0.1277   0.8970   0.0706
  -7.000  -0.1049   0.02699   0.02271  -0.1330   0.8624   0.0831
  -6.750  -0.0867   0.02482   0.02044  -0.1336   0.8514   0.0876
  -6.500  -0.0761   0.02168   0.01690  -0.1345   0.8410   0.0970
  -6.250  -0.0578   0.01984   0.01501  -0.1343   0.8311   0.1012
  -6.000  -0.0507   0.03345   0.02793  -0.1417   0.8440   0.1117
  -5.750  -0.0205   0.02449   0.01740  -0.1388   0.8350   0.0439
  -5.500   0.0089   0.02234   0.01493  -0.1392   0.8273   0.0420
  -5.250   0.0335   0.02087   0.01320  -0.1384   0.8183   0.0412
  -5.000   0.0622   0.01967   0.01178  -0.1384   0.8116   0.0410
  -4.750   0.0873   0.01897   0.01097  -0.1378   0.8042   0.0429
  -4.500   0.1142   0.01835   0.01021  -0.1375   0.7976   0.0446
  -4.250   0.1411   0.01771   0.00947  -0.1372   0.7919   0.0453
  -4.000   0.1623   0.01678   0.00862  -0.1360   0.7850   0.0470
  -3.750   0.1887   0.01620   0.00803  -0.1359   0.7796   0.0499
  -3.500   0.2147   0.01578   0.00758  -0.1359   0.7743   0.0541
  -3.250   0.2402   0.01527   0.00710  -0.1360   0.7687   0.0635
  -3.000   0.2716   0.01255   0.00622  -0.1400   0.7641   0.4776
  -2.750   0.2891   0.01401   0.00816  -0.1354   0.7598   0.6799
  -2.500   0.3072   0.01502   0.00915  -0.1319   0.7544   0.7048
  -2.250   0.3217   0.01617   0.01029  -0.1270   0.7499   0.7287
  -2.000   0.3335   0.01719   0.01129  -0.1210   0.7463   0.7497
  -1.750   0.3483   0.01779   0.01185  -0.1167   0.7425   0.7670
  -1.500   0.3635   0.01827   0.01231  -0.1131   0.7377   0.7874
  -1.250   0.3771   0.01849   0.01250  -0.1087   0.7336   0.8033
  -1.000   0.3960   0.01852   0.01246  -0.1058   0.7304   0.8140
  -0.750   0.4204   0.01849   0.01233  -0.1051   0.7273   0.8214
  -0.500   0.4389   0.01846   0.01230  -0.1034   0.7228   0.8274
  -0.250   0.4675   0.01854   0.01230  -0.1046   0.7187   0.8328
   0.000   0.4899   0.01835   0.01205  -0.1034   0.7155   0.8359
   0.250   0.5184   0.01830   0.01191  -0.1037   0.7126   0.8389
   0.500   0.5419   0.01833   0.01193  -0.1034   0.7088   0.8420
   0.750   0.5670   0.01838   0.01197  -0.1036   0.7047   0.8446
   1.000   0.5971   0.01845   0.01200  -0.1049   0.7011   0.8468
   1.250   0.6270   0.01843   0.01191  -0.1058   0.6980   0.8485
   1.500   0.6567   0.01842   0.01183  -0.1064   0.6953   0.8501
   1.750   0.6765   0.01850   0.01199  -0.1054   0.6910   0.8516
   2.000   0.7013   0.01857   0.01207  -0.1053   0.6869   0.8533
   2.250   0.7309   0.01861   0.01209  -0.1062   0.6834   0.8543
   2.500   0.7629   0.01866   0.01210  -0.1075   0.6805   0.8556
   2.750   0.7902   0.01881   0.01228  -0.1080   0.6767   0.8569
   3.000   0.8149   0.01897   0.01249  -0.1081   0.6718   0.8576
   3.250   0.8453   0.01904   0.01256  -0.1091   0.6677   0.8585
   3.500   0.8794   0.01906   0.01254  -0.1107   0.6641   0.8595
   3.750   0.9064   0.01923   0.01277  -0.1112   0.6593   0.8602
   4.000   0.9332   0.01934   0.01294  -0.1115   0.6536   0.8606
   4.250   0.9674   0.01928   0.01285  -0.1131   0.6489   0.8612
   4.500   0.9977   0.01936   0.01295  -0.1140   0.6435   0.8621
   4.750   1.0226   0.01939   0.01306  -0.1139   0.6368   0.8627
   5.000   1.0563   0.01925   0.01289  -0.1152   0.6318   0.8631
   5.250   1.0807   0.01930   0.01302  -0.1149   0.6249   0.8635
   5.500   1.1095   0.01917   0.01292  -0.1153   0.6180   0.8640
   5.750   1.1404   0.01905   0.01281  -0.1160   0.6119   0.8648
   6.000   1.1622   0.01904   0.01291  -0.1151   0.6040   0.8658
   6.250   1.1974   0.01882   0.01263  -0.1166   0.5977   0.8663
   6.500   1.2165   0.01885   0.01284  -0.1153   0.5886   0.8668
   6.750   1.2502   0.01861   0.01256  -0.1164   0.5812   0.8673
   7.000   1.2688   0.01860   0.01271  -0.1150   0.5711   0.8679
   7.250   1.2948   0.01847   0.01263  -0.1148   0.5618   0.8685
   7.500   1.3188   0.01832   0.01253  -0.1142   0.5512   0.8693
   7.750   1.3371   0.01828   0.01261  -0.1126   0.5390   0.8705
   8.000   1.3565   0.01822   0.01266  -0.1112   0.5258   0.8715
   8.250   1.3744   0.01820   0.01272  -0.1096   0.5102   0.8722
   8.500   1.3900   0.01825   0.01280  -0.1075   0.4920   0.8730
   8.750   1.4012   0.01836   0.01292  -0.1047   0.4709   0.8738
   9.000   1.4095   0.01871   0.01324  -0.1015   0.4437   0.8747
   9.250   1.4148   0.01935   0.01374  -0.0980   0.4114   0.8756
   9.500   1.4157   0.02031   0.01454  -0.0942   0.3756   0.8765
   9.750   1.4125   0.02156   0.01562  -0.0900   0.3414   0.8775
  10.000   1.4071   0.02303   0.01694  -0.0858   0.3105   0.8790
  10.250   1.4016   0.02472   0.01850  -0.0821   0.2831   0.8805
  10.500   1.3968   0.02660   0.02029  -0.0789   0.2579   0.8818
  10.750   1.3937   0.02861   0.02222  -0.0764   0.2348   0.8829
  11.000   1.3909   0.03078   0.02431  -0.0742   0.2143   0.8841
  11.250   1.3887   0.03308   0.02653  -0.0723   0.1966   0.8853
  11.500   1.3876   0.03543   0.02882  -0.0708   0.1807   0.8866
  11.750   1.3879   0.03778   0.03113  -0.0696   0.1662   0.8878
  12.000   1.3898   0.04009   0.03343  -0.0686   0.1533   0.8891
  12.250   1.3929   0.04239   0.03573  -0.0678   0.1417   0.8904
  12.500   1.3956   0.04475   0.03811  -0.0670   0.1318   0.8916
  12.750   1.3962   0.04729   0.04058  -0.0661   0.1235   0.8929
  13.000   1.4023   0.04939   0.04279  -0.0656   0.1151   0.8945
  13.250   1.4049   0.05188   0.04523  -0.0649   0.1084   0.8961
  13.500   1.4113   0.05406   0.04752  -0.0645   0.1017   0.8980
  13.750   1.4150   0.05651   0.04991  -0.0640   0.0960   0.9000
  14.000   1.4221   0.05873   0.05230  -0.0638   0.0905   0.9024
  14.250   1.4269   0.06120   0.05474  -0.0636   0.0857   0.9047
  14.500   1.4334   0.06345   0.05710  -0.0632   0.0811   0.9071
  14.750   1.4382   0.06593   0.05968  -0.0632   0.0768   0.9098
  15.000   1.4440   0.06825   0.06192  -0.0628   0.0723   0.9124
  15.250   1.4477   0.07102   0.06492  -0.0631   0.0688   0.9159
  15.500   1.4513   0.07377   0.06777  -0.0633   0.0653   0.9197
  15.750   1.4565   0.07609   0.07003  -0.0629   0.0615   0.9242
  16.000   1.4575   0.07910   0.07329  -0.0632   0.0589   0.9319
  16.250   1.4573   0.08187   0.07622  -0.0632   0.0561   0.9652
  16.500   1.4601   0.08498   0.07935  -0.0639   0.0533   1.0000
  16.750   1.4642   0.08805   0.08252  -0.0645   0.0504   1.0000
  17.000   1.4653   0.09174   0.08639  -0.0656   0.0478   1.0000
  17.250   1.4666   0.09535   0.09006  -0.0669   0.0453   1.0000
  17.500   1.4700   0.09849   0.09319  -0.0676   0.0426   1.0000
  17.750   1.4673   0.10292   0.09786  -0.0693   0.0405   1.0000
  18.000   1.4669   0.10695   0.10201  -0.0710   0.0384   1.0000
  18.250   1.4689   0.11040   0.10544  -0.0723   0.0364   1.0000
  18.500   1.4673   0.11455   0.10976  -0.0738   0.0345   1.0000
  18.750   1.4631   0.11936   0.11478  -0.0760   0.0329   1.0000
<< Back to NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)