NASA/LANGLEY LS(1)-0421MOD AIRFOIL (ls421mod-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY LS(1)-0421MOD AIRFOIL (ls421mod-il) Reynolds number: 50,000 Max Cl/Cd: 7.25 at α=0.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ls421mod-il-50000.txt Download as CSV file: xf-ls421mod-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY LS(1)-0421MOD 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.2531 0.13330 0.12560 -0.0269 1.0000 0.3345
-11.750 -0.6557 0.07524 0.06728 -0.0631 1.0000 0.1597
-11.500 -0.6924 0.07214 0.06420 -0.0594 1.0000 0.1590
-11.250 -0.7318 0.07004 0.06211 -0.0542 1.0000 0.1583
-11.000 -0.7690 0.06775 0.05977 -0.0493 1.0000 0.1578
-10.750 -0.8021 0.06534 0.05724 -0.0447 1.0000 0.1575
-10.500 -0.8299 0.06275 0.05445 -0.0407 1.0000 0.1578
-10.250 -0.8520 0.06003 0.05146 -0.0372 1.0000 0.1585
-10.000 -0.8687 0.05723 0.04826 -0.0343 1.0000 0.1596
-9.750 -0.8585 0.05536 0.04654 -0.0326 1.0000 0.1630
-9.500 -0.8532 0.05355 0.04467 -0.0308 1.0000 0.1664
-9.250 -0.8510 0.05135 0.04223 -0.0291 1.0000 0.1701
-9.000 -0.8493 0.04895 0.03931 -0.0278 1.0000 0.1745
-8.750 -0.8341 0.04757 0.03817 -0.0265 1.0000 0.1805
-8.500 -0.8226 0.04593 0.03633 -0.0255 1.0000 0.1875
-8.250 -0.8087 0.04442 0.03482 -0.0244 1.0000 0.1952
-8.000 -0.7947 0.04309 0.03333 -0.0236 1.0000 0.2054
-7.750 -0.7791 0.04200 0.03244 -0.0224 1.0000 0.2163
-7.500 -0.7636 0.04086 0.03133 -0.0214 1.0000 0.2301
-7.250 -0.7478 0.03984 0.03038 -0.0205 1.0000 0.2468
-7.000 -0.7321 0.03895 0.02968 -0.0193 1.0000 0.2668
-6.750 -0.7165 0.03816 0.02915 -0.0182 1.0000 0.2911
-6.500 -0.7002 0.03727 0.02847 -0.0174 1.0000 0.3231
-6.250 -0.6845 0.03652 0.02812 -0.0163 1.0000 0.3633
-6.000 -0.6537 0.03651 0.02887 -0.0165 0.9936 0.4273
-5.750 -0.6248 0.03853 0.03150 -0.0138 0.9845 0.4905
-5.500 -0.5923 0.04242 0.03558 -0.0100 0.9750 0.5363
-5.250 -0.5624 0.04612 0.03924 -0.0061 0.9648 0.5691
-5.000 -0.5297 0.05009 0.04312 -0.0021 0.9547 0.5946
-4.750 -0.4929 0.05315 0.04603 -0.0005 0.9450 0.6260
-4.500 -0.4644 0.05591 0.04869 0.0030 0.9351 0.6476
-4.250 -0.4016 0.05936 0.05194 0.0023 0.9251 0.6729
-4.000 -0.3776 0.06064 0.05312 0.0045 0.9152 0.6953
-3.750 -0.3132 0.06230 0.05457 0.0008 0.9055 0.7249
-3.500 -0.2894 0.06285 0.05503 0.0019 0.8954 0.7473
-3.250 -0.2577 0.06314 0.05518 0.0008 0.8864 0.7734
-3.000 -0.1550 0.06381 0.05564 -0.0099 0.8752 0.8009
-2.750 -0.0975 0.06376 0.05545 -0.0151 0.8655 0.8298
-2.500 -0.0190 0.06332 0.05485 -0.0240 0.8556 0.8578
-2.250 0.0154 0.06296 0.05441 -0.0266 0.8454 0.8790
-2.000 0.1121 0.06166 0.05293 -0.0391 0.8371 0.9079
-1.750 0.1083 0.06196 0.05323 -0.0363 0.8256 0.9194
-1.500 0.2076 0.06030 0.05143 -0.0497 0.8171 0.9434
-1.250 0.1874 0.06128 0.05241 -0.0446 0.8056 0.9491
-1.000 0.2555 0.06034 0.05139 -0.0532 0.7964 0.9640
-0.750 0.2655 0.06091 0.05194 -0.0530 0.7855 0.9709
-0.500 0.3102 0.06060 0.05159 -0.0579 0.7752 0.9803
-0.250 0.3443 0.06085 0.05180 -0.0614 0.7650 0.9881
0.000 0.3706 0.06121 0.05215 -0.0637 0.7538 0.9945
0.250 0.4185 0.06108 0.05198 -0.0687 0.7443 1.0000
0.500 0.3902 0.06290 0.05383 -0.0626 0.7334 1.0000
0.750 0.4484 0.06185 0.05270 -0.0674 0.7258 1.0000
1.000 0.3839 0.06501 0.05592 -0.0570 0.7148 1.0000
1.250 0.4097 0.06506 0.05594 -0.0576 0.7061 1.0000
1.500 0.3890 0.06679 0.05768 -0.0528 0.6969 1.0000
1.750 0.3810 0.06801 0.05890 -0.0496 0.6880 1.0000
2.000 0.4168 0.06792 0.05876 -0.0510 0.6794 1.0000
2.250 0.3703 0.07039 0.06128 -0.0438 0.6704 1.0000
2.500 0.4002 0.07043 0.06128 -0.0444 0.6613 1.0000
2.750 0.3734 0.07237 0.06324 -0.0395 0.6529 1.0000
3.000 0.3673 0.07351 0.06439 -0.0367 0.6439 1.0000
3.250 0.3885 0.07395 0.06480 -0.0361 0.6345 1.0000
3.500 0.3549 0.07599 0.06686 -0.0310 0.6266 1.0000
3.750 0.4101 0.07522 0.06605 -0.0331 0.6152 1.0000
4.000 0.3536 0.07805 0.06892 -0.0264 0.6085 1.0000
4.250 0.3661 0.07851 0.06937 -0.0249 0.5980 1.0000
4.500 0.3553 0.07988 0.07074 -0.0218 0.5890 1.0000
4.750 0.3428 0.08116 0.07203 -0.0187 0.5804 1.0000
5.000 0.2169 0.09076 0.08188 -0.0146 0.6783 1.0000
5.250 0.2379 0.09187 0.08296 -0.0141 0.6578 1.0000
5.500 0.2198 0.09283 0.08394 -0.0112 0.6564 1.0000
5.750 0.2458 0.09436 0.08544 -0.0114 0.6364 1.0000
6.000 0.2251 0.09527 0.08638 -0.0089 0.6335 1.0000
6.250 0.3335 0.08798 0.07886 -0.0065 0.5215 1.0000
6.500 0.3688 0.08841 0.07929 -0.0070 0.5073 1.0000
6.750 0.3545 0.09153 0.08246 -0.0069 0.4987 1.0000
7.000 0.3819 0.09277 0.08370 -0.0078 0.4849 1.0000
7.250 0.3883 0.09540 0.08637 -0.0088 0.4748 1.0000
7.500 0.4020 0.09771 0.08872 -0.0099 0.4628 1.0000
7.750 0.4554 0.09753 0.08853 -0.0110 0.4480 1.0000
8.000 0.4259 0.10318 0.09426 -0.0127 0.4426 1.0000
8.250 0.4753 0.10331 0.09439 -0.0137 0.4277 1.0000
8.500 0.4534 0.10903 0.10019 -0.0159 0.4245 1.0000
8.750 0.4478 0.11368 0.10491 -0.0181 0.4226 1.0000
9.000 0.4524 0.11849 0.10979 -0.0208 0.4249 1.0000
9.250 0.4728 0.12343 0.11479 -0.0240 0.4279 1.0000
9.500 0.5022 0.12477 0.11613 -0.0249 0.4086 1.0000
9.750 0.4604 0.13746 0.12903 -0.0325 0.4696 1.0000
10.000 0.4419 0.13893 0.13053 -0.0335 0.4648 1.0000
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Polar data table (+)
Polar graphs
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