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 LS(1)-0417MOD AIRFOIL (ls417mod-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0417MOD AIRFOIL (ls417mod-il)
Reynolds number: 50,000
Max Cl/Cd: 26.31 at α=7.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ls417mod-il-50000-n5.txt
Download as CSV file: xf-ls417mod-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY LS(1)-0417MOD 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.4816   0.10130   0.09327  -0.0492   1.0000   0.0632
 -11.750  -0.5185   0.08988   0.08178  -0.0563   1.0000   0.0623
 -11.500  -0.5692   0.07939   0.07113  -0.0624   1.0000   0.0612
 -11.250  -0.6120   0.07241   0.06395  -0.0648   1.0000   0.0606
 -11.000  -0.6439   0.06780   0.05917  -0.0646   1.0000   0.0603
 -10.750  -0.6717   0.06449   0.05571  -0.0624   1.0000   0.0602
 -10.500  -0.6952   0.06184   0.05294  -0.0594   1.0000   0.0602
 -10.250  -0.7096   0.05927   0.05022  -0.0568   1.0000   0.0605
 -10.000  -0.7193   0.05691   0.04770  -0.0541   1.0000   0.0608
  -9.750  -0.7254   0.05465   0.04526  -0.0514   1.0000   0.0613
  -9.500  -0.7285   0.05252   0.04293  -0.0488   1.0000   0.0621
  -9.250  -0.7296   0.05042   0.04057  -0.0462   1.0000   0.0632
  -9.000  -0.7290   0.04823   0.03800  -0.0438   1.0000   0.0646
  -8.750  -0.7195   0.04684   0.03662  -0.0420   1.0000   0.0661
  -8.500  -0.7094   0.04541   0.03510  -0.0401   1.0000   0.0677
  -8.250  -0.6983   0.04380   0.03327  -0.0383   1.0000   0.0695
  -8.000  -0.6848   0.04219   0.03140  -0.0366   1.0000   0.0712
  -7.750  -0.6696   0.04106   0.03031  -0.0350   1.0000   0.0730
  -7.500  -0.6543   0.03996   0.02913  -0.0334   1.0000   0.0757
  -7.250  -0.6380   0.03890   0.02797  -0.0318   1.0000   0.0788
  -7.000  -0.6215   0.03803   0.02712  -0.0302   1.0000   0.0821
  -6.750  -0.6037   0.03717   0.02615  -0.0285   1.0000   0.0858
  -6.500  -0.5870   0.03641   0.02547  -0.0269   1.0000   0.0897
  -6.250  -0.5701   0.03568   0.02471  -0.0253   1.0000   0.0951
  -6.000  -0.5535   0.03497   0.02397  -0.0238   1.0000   0.1016
  -5.750  -0.5374   0.03418   0.02326  -0.0225   1.0000   0.1090
  -5.500  -0.5177   0.03334   0.02249  -0.0219   0.9990   0.1202
  -5.250  -0.4848   0.03224   0.02158  -0.0241   0.9934   0.1410
  -5.000  -0.4501   0.03084   0.02055  -0.0271   0.9862   0.1801
  -4.750  -0.4106   0.02902   0.01951  -0.0317   0.9783   0.2726
  -4.500  -0.3760   0.02787   0.01943  -0.0337   0.9683   0.4192
  -4.250  -0.3496   0.02905   0.02127  -0.0298   0.9564   0.5243
  -4.000  -0.3115   0.03005   0.02216  -0.0300   0.9442   0.5942
  -3.750  -0.2681   0.03103   0.02292  -0.0311   0.9333   0.6369
  -3.500  -0.2392   0.03174   0.02349  -0.0296   0.9202   0.6623
  -3.250  -0.2069   0.03241   0.02400  -0.0286   0.9101   0.6848
  -3.000  -0.1752   0.03294   0.02440  -0.0275   0.8997   0.7050
  -2.750  -0.1449   0.03323   0.02455  -0.0266   0.8888   0.7232
  -2.500  -0.1083   0.03319   0.02437  -0.0274   0.8783   0.7401
  -2.250  -0.0785   0.03315   0.02422  -0.0264   0.8661   0.7510
  -2.000  -0.0390   0.03275   0.02366  -0.0279   0.8551   0.7629
  -1.750  -0.0134   0.03250   0.02333  -0.0267   0.8401   0.7728
  -1.500   0.0223   0.03202   0.02273  -0.0275   0.8275   0.7832
  -1.250   0.0528   0.03158   0.02220  -0.0270   0.8116   0.7920
  -1.000   0.0779   0.03114   0.02167  -0.0262   0.7932   0.8016
  -0.750   0.1062   0.03066   0.02112  -0.0252   0.7743   0.8091
  -0.500   0.1322   0.03016   0.02053  -0.0247   0.7529   0.8182
  -0.250   0.1587   0.02964   0.01995  -0.0235   0.7283   0.8243
   0.000   0.1850   0.02914   0.01935  -0.0227   0.7010   0.8313
   0.250   0.2097   0.02871   0.01881  -0.0220   0.6698   0.8380
   0.500   0.2388   0.02823   0.01816  -0.0214   0.6375   0.8431
   0.750   0.2677   0.02786   0.01754  -0.0211   0.6062   0.8489
   1.000   0.2962   0.02762   0.01701  -0.0211   0.5787   0.8544
   1.250   0.3245   0.02746   0.01656  -0.0208   0.5551   0.8586
   1.500   0.3510   0.02744   0.01629  -0.0206   0.5351   0.8633
   1.750   0.3768   0.02753   0.01617  -0.0205   0.5183   0.8679
   2.000   0.4037   0.02766   0.01607  -0.0207   0.5044   0.8721
   2.250   0.4297   0.02778   0.01607  -0.0206   0.4909   0.8760
   2.500   0.4571   0.02797   0.01611  -0.0208   0.4794   0.8799
   2.750   0.4843   0.02820   0.01625  -0.0212   0.4685   0.8838
   3.000   0.5129   0.02845   0.01639  -0.0218   0.4591   0.8876
   3.250   0.5391   0.02868   0.01661  -0.0219   0.4491   0.8914
   3.500   0.5691   0.02895   0.01675  -0.0226   0.4408   0.8949
   3.750   0.5953   0.02931   0.01714  -0.0230   0.4323   0.8985
   4.000   0.6225   0.02966   0.01748  -0.0234   0.4240   0.9023
   4.250   0.6527   0.02998   0.01767  -0.0240   0.4170   0.9060
   4.500   0.6764   0.03040   0.01820  -0.0240   0.4095   0.9103
   4.750   0.7025   0.03085   0.01869  -0.0243   0.4020   0.9143
   5.000   0.7323   0.03126   0.01902  -0.0251   0.3955   0.9176
   5.250   0.7568   0.03179   0.01964  -0.0251   0.3889   0.9214
   5.500   0.7804   0.03234   0.02031  -0.0251   0.3819   0.9261
   5.750   0.8074   0.03283   0.02080  -0.0255   0.3756   0.9308
   6.000   0.8358   0.03335   0.02127  -0.0261   0.3701   0.9354
   6.250   0.8555   0.03412   0.02227  -0.0257   0.3633   0.9407
   6.500   0.8800   0.03476   0.02301  -0.0259   0.3569   0.9455
   6.750   0.9095   0.03529   0.02349  -0.0267   0.3516   0.9502
   7.000   0.9311   0.03619   0.02456  -0.0267   0.3457   0.9562
   7.250   0.9517   0.03711   0.02570  -0.0267   0.3391   0.9630
   7.500   0.9789   0.03780   0.02643  -0.0274   0.3334   0.9704
   7.750   1.0111   0.03843   0.02702  -0.0288   0.3286   0.9787
   8.000   1.0219   0.03984   0.02882  -0.0280   0.3218   1.0000
   8.250   1.0427   0.04096   0.03006  -0.0282   0.3160   1.0000
   8.500   1.0729   0.04163   0.03069  -0.0293   0.3111   1.0000
   8.750   1.0865   0.04324   0.03252  -0.0290   0.3055   1.0000
   9.000   1.0960   0.04504   0.03459  -0.0283   0.2994   1.0000
   9.250   1.1205   0.04598   0.03558  -0.0289   0.2942   1.0000
   9.500   1.1496   0.04675   0.03633  -0.0299   0.2897   1.0000
   9.750   1.1352   0.04996   0.03996  -0.0275   0.2836   1.0000
  10.000   1.1406   0.05185   0.04202  -0.0265   0.2785   1.0000
  10.250   1.1769   0.05206   0.04218  -0.0279   0.2740   1.0000
  10.500   1.1524   0.05597   0.04639  -0.0253   0.2693   1.0000
  10.750   1.0918   0.06368   0.05447  -0.0235   0.2635   1.0000
  11.000   1.1047   0.06547   0.05633  -0.0239   0.2589   1.0000
  11.250   1.1600   0.06337   0.05414  -0.0244   0.2554   1.0000
  11.500   0.9584   0.09241   0.08368  -0.0330   0.2415   1.0000
  11.750   0.9980   0.09006   0.08136  -0.0315   0.2398   1.0000
<< Back to NASA/LANGLEY LS(1)-0417MOD AIRFOIL (ls417mod-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY LS(1)-0417MOD AIRFOIL (ls417mod-il)