NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Reynolds number: 50,000 Max Cl/Cd: 25.87 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ls013-il-50000-n5.txt Download as CSV file: xf-ls013-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY LS(1)-0013 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 -0.7957 0.09361 0.08615 -0.0140 1.0000 0.0718
-11.750 -0.8295 0.08470 0.07712 -0.0194 1.0000 0.0713
-11.500 -0.8606 0.07768 0.06992 -0.0226 1.0000 0.0709
-11.250 -0.8874 0.07209 0.06411 -0.0238 1.0000 0.0708
-11.000 -0.9098 0.06750 0.05927 -0.0234 1.0000 0.0709
-10.750 -0.9268 0.06349 0.05496 -0.0220 1.0000 0.0713
-10.500 -0.9370 0.05956 0.05064 -0.0207 1.0000 0.0721
-10.250 -0.9429 0.05583 0.04641 -0.0192 1.0000 0.0734
-10.000 -0.9358 0.05284 0.04318 -0.0183 1.0000 0.0752
-9.750 -0.9214 0.05060 0.04084 -0.0178 1.0000 0.0775
-9.500 -0.9078 0.04813 0.03813 -0.0171 1.0000 0.0798
-9.250 -0.8929 0.04553 0.03517 -0.0164 1.0000 0.0822
-9.000 -0.8774 0.04310 0.03226 -0.0156 1.0000 0.0856
-8.750 -0.8566 0.04105 0.03011 -0.0154 1.0000 0.0890
-8.500 -0.8342 0.03928 0.02825 -0.0152 1.0000 0.0926
-8.250 -0.8115 0.03755 0.02623 -0.0149 1.0000 0.0973
-8.000 -0.7869 0.03589 0.02445 -0.0147 1.0000 0.1023
-7.750 -0.7615 0.03449 0.02303 -0.0147 1.0000 0.1074
-7.500 -0.7355 0.03321 0.02149 -0.0144 1.0000 0.1144
-7.250 -0.7102 0.03198 0.02039 -0.0143 1.0000 0.1214
-7.000 -0.6837 0.03088 0.01916 -0.0140 1.0000 0.1298
-6.750 -0.6612 0.02982 0.01818 -0.0135 1.0000 0.1398
-6.500 -0.6406 0.02878 0.01718 -0.0127 1.0000 0.1515
-6.250 -0.6227 0.02776 0.01623 -0.0115 1.0000 0.1654
-6.000 -0.6085 0.02676 0.01538 -0.0100 1.0000 0.1841
-5.750 -0.5979 0.02579 0.01460 -0.0078 1.0000 0.2087
-5.500 -0.5913 0.02478 0.01389 -0.0051 1.0000 0.2440
-5.250 -0.5860 0.02373 0.01322 -0.0022 1.0000 0.2979
-5.000 -0.5807 0.02267 0.01269 0.0009 1.0000 0.3769
-4.750 -0.5722 0.02212 0.01293 0.0048 1.0000 0.4812
-4.500 -0.5582 0.02251 0.01371 0.0089 1.0000 0.5758
-4.250 -0.5453 0.02286 0.01405 0.0122 1.0000 0.6325
-4.000 -0.5321 0.02323 0.01432 0.0154 1.0000 0.6737
-3.750 -0.5068 0.02384 0.01480 0.0169 0.9949 0.7109
-3.500 -0.4787 0.02456 0.01539 0.0182 0.9890 0.7420
-3.250 -0.4493 0.02530 0.01599 0.0194 0.9837 0.7706
-3.000 -0.4194 0.02587 0.01642 0.0203 0.9787 0.7954
-2.750 -0.3896 0.02615 0.01655 0.0204 0.9733 0.8152
-2.500 -0.3535 0.02632 0.01656 0.0192 0.9692 0.8292
-2.250 -0.3186 0.02641 0.01651 0.0180 0.9646 0.8409
-2.000 -0.2850 0.02640 0.01637 0.0166 0.9596 0.8522
-1.750 -0.2499 0.02636 0.01621 0.0148 0.9552 0.8640
-1.500 -0.2117 0.02641 0.01617 0.0126 0.9510 0.8744
-1.250 -0.1752 0.02642 0.01609 0.0107 0.9462 0.8849
-1.000 -0.1399 0.02637 0.01598 0.0086 0.9417 0.8954
-0.750 -0.0997 0.02634 0.01589 0.0056 0.9374 0.9013
-0.500 -0.0723 0.02628 0.01579 0.0048 0.9309 0.9093
-0.250 -0.0311 0.02625 0.01574 0.0015 0.9266 0.9142
0.000 0.0000 0.02625 0.01573 0.0000 0.9204 0.9204
0.250 0.0311 0.02625 0.01574 -0.0015 0.9142 0.9266
0.500 0.0723 0.02628 0.01579 -0.0048 0.9093 0.9309
0.750 0.0997 0.02634 0.01589 -0.0056 0.9013 0.9374
1.000 0.1399 0.02637 0.01598 -0.0086 0.8954 0.9417
1.250 0.1752 0.02641 0.01609 -0.0107 0.8850 0.9463
1.500 0.2117 0.02641 0.01616 -0.0126 0.8744 0.9510
1.750 0.2499 0.02636 0.01621 -0.0148 0.8640 0.9552
2.000 0.2850 0.02639 0.01636 -0.0166 0.8522 0.9596
2.250 0.3186 0.02640 0.01651 -0.0180 0.8409 0.9646
2.500 0.3535 0.02632 0.01656 -0.0192 0.8292 0.9692
2.750 0.3896 0.02615 0.01655 -0.0204 0.8152 0.9733
3.000 0.4194 0.02587 0.01642 -0.0203 0.7954 0.9787
3.250 0.4493 0.02530 0.01598 -0.0194 0.7706 0.9837
3.500 0.4787 0.02456 0.01538 -0.0182 0.7420 0.9890
3.750 0.5068 0.02384 0.01479 -0.0169 0.7109 0.9949
4.000 0.5320 0.02323 0.01431 -0.0154 0.6737 1.0000
4.250 0.5452 0.02286 0.01404 -0.0122 0.6327 1.0000
4.500 0.5581 0.02250 0.01370 -0.0088 0.5760 1.0000
4.750 0.5720 0.02211 0.01293 -0.0048 0.4814 1.0000
5.000 0.5806 0.02266 0.01268 -0.0009 0.3771 1.0000
5.250 0.5859 0.02372 0.01321 0.0022 0.2981 1.0000
5.500 0.5912 0.02478 0.01389 0.0051 0.2441 1.0000
5.750 0.5978 0.02578 0.01459 0.0078 0.2087 1.0000
6.000 0.6085 0.02676 0.01537 0.0100 0.1841 1.0000
6.250 0.6228 0.02776 0.01623 0.0115 0.1654 1.0000
6.500 0.6407 0.02878 0.01718 0.0127 0.1515 1.0000
6.750 0.6613 0.02982 0.01818 0.0135 0.1398 1.0000
7.000 0.6838 0.03088 0.01916 0.0140 0.1298 1.0000
7.250 0.7103 0.03198 0.02039 0.0143 0.1214 1.0000
7.500 0.7356 0.03321 0.02149 0.0144 0.1144 1.0000
7.750 0.7617 0.03449 0.02303 0.0146 0.1074 1.0000
8.000 0.7870 0.03590 0.02445 0.0147 0.1023 1.0000
8.250 0.8116 0.03755 0.02623 0.0149 0.0973 1.0000
8.500 0.8344 0.03929 0.02825 0.0152 0.0926 1.0000
8.750 0.8568 0.04105 0.03012 0.0154 0.0890 1.0000
9.000 0.8776 0.04310 0.03226 0.0156 0.0856 1.0000
9.250 0.8931 0.04553 0.03517 0.0164 0.0822 1.0000
9.500 0.9080 0.04813 0.03813 0.0171 0.0798 1.0000
9.750 0.9216 0.05061 0.04085 0.0178 0.0775 1.0000
10.000 0.9361 0.05285 0.04318 0.0183 0.0752 1.0000
10.250 0.9431 0.05585 0.04643 0.0191 0.0734 1.0000
10.500 0.9372 0.05958 0.05066 0.0206 0.0721 1.0000
10.750 0.9271 0.06351 0.05499 0.0220 0.0713 1.0000
11.000 0.9101 0.06753 0.05930 0.0233 0.0709 1.0000
11.250 0.8878 0.07213 0.06416 0.0237 0.0708 1.0000
11.500 0.8610 0.07775 0.06999 0.0224 0.0709 1.0000
11.750 0.8299 0.08479 0.07721 0.0192 0.0713 1.0000
12.000 0.7962 0.09372 0.08626 0.0138 0.0718 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il)