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NASA/LANGLEY NLF(1)-0416 AIRFOIL (nlf416-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY NLF(1)-0416 AIRFOIL (nlf416-il)
Reynolds number: 50,000
Max Cl/Cd: 14.83 at α=0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nlf416-il-50000.txt
Download as CSV file: xf-nlf416-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY NLF(1)-0416 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4731   0.08982   0.08309  -0.0537   1.0000   0.1280
  -9.750  -0.4365   0.08975   0.08303  -0.0486   1.0000   0.1312
  -9.500  -0.5617   0.07613   0.06947  -0.0576   1.0000   0.1138
  -9.250  -0.5774   0.07229   0.06562  -0.0566   1.0000   0.1124
  -9.000  -0.5914   0.06846   0.06179  -0.0556   1.0000   0.1114
  -8.750  -0.6058   0.06478   0.05805  -0.0543   1.0000   0.1102
  -8.500  -0.6210   0.06109   0.05422  -0.0529   1.0000   0.1093
  -8.250  -0.6327   0.05744   0.05037  -0.0514   1.0000   0.1086
  -8.000  -0.6385   0.05392   0.04659  -0.0500   1.0000   0.1081
  -7.750  -0.6393   0.05055   0.04292  -0.0486   1.0000   0.1081
  -7.500  -0.6355   0.04734   0.03937  -0.0473   1.0000   0.1085
  -7.250  -0.6275   0.04434   0.03600  -0.0462   1.0000   0.1092
  -7.000  -0.6161   0.04166   0.03291  -0.0450   1.0000   0.1106
  -6.750  -0.6028   0.03933   0.03014  -0.0438   1.0000   0.1137
  -6.500  -0.5889   0.03759   0.02857  -0.0423   1.0000   0.1197
  -6.250  -0.5738   0.03600   0.02668  -0.0410   1.0000   0.1266
  -6.000  -0.5595   0.03455   0.02536  -0.0393   1.0000   0.1350
  -5.750  -0.5458   0.03333   0.02417  -0.0376   1.0000   0.1476
  -5.500  -0.5335   0.03228   0.02327  -0.0356   1.0000   0.1653
  -5.250  -0.5220   0.03118   0.02239  -0.0337   1.0000   0.1941
  -5.000  -0.5111   0.02970   0.02147  -0.0323   1.0000   0.2489
  -4.750  -0.4990   0.02762   0.02059  -0.0318   1.0000   0.3737
  -4.500  -0.4859   0.02874   0.02285  -0.0272   0.9937   0.5049
  -4.250  -0.4564   0.03363   0.02780  -0.0213   0.9753   0.5732
  -4.000  -0.4219   0.03774   0.03165  -0.0171   0.9594   0.6194
  -3.750  -0.3891   0.04058   0.03426  -0.0134   0.9452   0.6557
  -3.500  -0.3586   0.04221   0.03564  -0.0106   0.9317   0.6898
  -3.250  -0.3310   0.04307   0.03628  -0.0082   0.9190   0.7217
  -3.000  -0.3004   0.04348   0.03645  -0.0067   0.9072   0.7533
  -2.750  -0.2501   0.04362   0.03630  -0.0085   0.8981   0.7873
  -2.500  -0.2146   0.04336   0.03582  -0.0091   0.8870   0.8164
  -2.250  -0.1582   0.04291   0.03508  -0.0136   0.8778   0.8461
  -2.000  -0.0601   0.04202   0.03381  -0.0254   0.8698   0.8767
  -1.750   0.0035   0.04126   0.03281  -0.0327   0.8608   0.9009
  -1.500   0.1104   0.03973   0.03097  -0.0474   0.8534   0.9250
  -1.250   0.1824   0.03864   0.02967  -0.0566   0.8456   0.9439
  -1.000   0.2694   0.03721   0.02805  -0.0686   0.8369   0.9610
  -0.750   0.3495   0.03586   0.02653  -0.0795   0.8290   0.9761
  -0.500   0.4126   0.03493   0.02550  -0.0876   0.8193   0.9886
  -0.250   0.4670   0.03432   0.02481  -0.0943   0.8097   0.9991
   0.000   0.4946   0.03431   0.02473  -0.0957   0.8006   1.0000
   0.250   0.4887   0.03550   0.02592  -0.0919   0.7894   1.0000
   0.500   0.5229   0.03527   0.02563  -0.0939   0.7820   1.0000
   0.750   0.4980   0.03722   0.02761  -0.0875   0.7706   1.0000
   1.000   0.5400   0.03678   0.02710  -0.0903   0.7640   1.0000
   1.250   0.4933   0.03949   0.02987  -0.0811   0.7525   1.0000
   1.500   0.5188   0.03973   0.03007  -0.0815   0.7454   1.0000
   1.750   0.4636   0.04247   0.03287  -0.0714   0.7352   1.0000
   2.000   0.5017   0.04249   0.03285  -0.0735   0.7286   1.0000
   2.250   0.4090   0.04595   0.03636  -0.0590   0.7196   1.0000
   2.500   0.4686   0.04549   0.03587  -0.0634   0.7127   1.0000
   2.750   0.3782   0.04868   0.03909  -0.0501   0.7058   1.0000
   3.000   0.3803   0.04959   0.04000  -0.0478   0.6987   1.0000
   3.250   0.3695   0.05082   0.04122  -0.0440   0.6922   1.0000
   3.500   0.3366   0.05233   0.04274  -0.0380   0.6870   1.0000
   3.750   0.3671   0.05284   0.04323  -0.0386   0.6788   1.0000
   4.000   0.3303   0.05443   0.04483  -0.0327   0.6745   1.0000
   4.250   0.3169   0.05589   0.04629  -0.0299   0.6694   1.0000
   4.500   0.3506   0.05684   0.04724  -0.0312   0.6596   1.0000
   4.750   0.3394   0.05891   0.04934  -0.0301   0.6547   1.0000
   5.000   0.3829   0.05991   0.05034  -0.0323   0.6432   1.0000
   5.250   0.3716   0.06242   0.05289  -0.0320   0.6385   1.0000
   5.500   0.3890   0.06435   0.05485  -0.0333   0.6308   1.0000
   5.750   0.4054   0.06639   0.05692  -0.0345   0.6217   1.0000
   6.000   0.4127   0.06893   0.05952  -0.0358   0.6172   1.0000
   6.250   0.3917   0.07486   0.06556  -0.0396   0.6590   1.0000
   6.500   0.4192   0.07684   0.06758  -0.0412   0.6424   1.0000
   6.750   0.4651   0.07615   0.06690  -0.0407   0.5954   1.0000
   7.000   0.4763   0.07937   0.07019  -0.0428   0.5945   1.0000
   7.250   0.4942   0.08261   0.07350  -0.0450   0.5913   1.0000
   7.500   0.5179   0.08451   0.07545  -0.0460   0.5742   1.0000
   8.000   0.5716   0.08826   0.07934  -0.0481   0.5408   1.0000
   9.000   0.5407   0.10596   0.09737  -0.0572   0.5782   1.0000
   9.250   0.5464   0.10874   0.10021  -0.0582   0.5669   1.0000
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