FX 60-100 (126) AIRFOIL (fx601001-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 60-100 (126) AIRFOIL (fx601001-il) Reynolds number: 50,000 Max Cl/Cd: 42.51 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx601001-il-50000-n5.txt Download as CSV file: xf-fx601001-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-100 (126) 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3768 0.10042 0.09346 -0.0381 1.0000 0.0458
-9.000 -0.3770 0.09670 0.08983 -0.0391 1.0000 0.0450
-8.750 -0.3790 0.09288 0.08611 -0.0402 1.0000 0.0443
-8.500 -0.3829 0.08904 0.08238 -0.0413 1.0000 0.0435
-8.250 -0.3893 0.08514 0.07860 -0.0422 1.0000 0.0428
-8.000 -0.3996 0.08112 0.07473 -0.0431 1.0000 0.0421
-7.500 -0.4356 0.06710 0.06084 -0.0551 1.0000 0.0392
-7.250 -0.4393 0.06270 0.05640 -0.0580 1.0000 0.0390
-7.000 -0.4401 0.05805 0.05161 -0.0610 1.0000 0.0388
-6.750 -0.4359 0.05325 0.04655 -0.0643 1.0000 0.0388
-6.500 -0.4253 0.04849 0.04140 -0.0676 1.0000 0.0389
-6.250 -0.4083 0.04379 0.03614 -0.0707 1.0000 0.0392
-6.000 -0.3899 0.04050 0.03262 -0.0722 1.0000 0.0406
-5.750 -0.3698 0.03818 0.03002 -0.0731 1.0000 0.0435
-5.500 -0.3450 0.03529 0.02659 -0.0745 1.0000 0.0472
-5.250 -0.3184 0.03260 0.02325 -0.0753 1.0000 0.0497
-5.000 -0.2958 0.03041 0.02099 -0.0753 1.0000 0.0530
-4.750 -0.2719 0.02884 0.01916 -0.0751 1.0000 0.0602
-4.500 -0.2483 0.02737 0.01759 -0.0750 1.0000 0.0700
-4.250 -0.2236 0.02588 0.01600 -0.0749 1.0000 0.0835
-4.000 -0.1975 0.02467 0.01481 -0.0758 1.0000 0.1116
-3.750 -0.1686 0.02354 0.01366 -0.0771 0.9993 0.1432
-3.500 -0.1254 0.02222 0.01257 -0.0814 0.9936 0.2026
-3.250 -0.0861 0.02143 0.01270 -0.0843 0.9872 0.3549
-3.000 -0.0497 0.02205 0.01315 -0.0857 0.9781 0.4520
-2.750 -0.0144 0.02258 0.01340 -0.0870 0.9688 0.5079
-2.500 0.0139 0.02310 0.01394 -0.0863 0.9591 0.5558
-2.250 0.0455 0.02336 0.01410 -0.0865 0.9506 0.5888
-2.000 0.0792 0.02339 0.01390 -0.0877 0.9414 0.6086
-1.750 0.1124 0.02339 0.01372 -0.0888 0.9320 0.6238
-1.500 0.1506 0.02338 0.01352 -0.0909 0.9247 0.6380
-1.250 0.1819 0.02338 0.01335 -0.0918 0.9143 0.6507
-1.000 0.2147 0.02338 0.01325 -0.0928 0.9050 0.6639
-0.750 0.2513 0.02335 0.01311 -0.0945 0.8972 0.6781
-0.500 0.2821 0.02336 0.01305 -0.0952 0.8869 0.6900
-0.250 0.3176 0.02333 0.01294 -0.0968 0.8783 0.7000
0.000 0.3537 0.02330 0.01282 -0.0985 0.8697 0.7095
0.250 0.3841 0.02330 0.01280 -0.0992 0.8593 0.7176
0.500 0.4182 0.02329 0.01275 -0.1004 0.8502 0.7265
0.750 0.4529 0.02322 0.01268 -0.1017 0.8414 0.7356
1.000 0.4827 0.02325 0.01273 -0.1021 0.8303 0.7446
1.250 0.5144 0.02323 0.01275 -0.1028 0.8197 0.7548
1.500 0.5486 0.02308 0.01264 -0.1037 0.8099 0.7659
1.750 0.5798 0.02294 0.01257 -0.1039 0.7983 0.7781
2.000 0.6083 0.02285 0.01260 -0.1037 0.7853 0.7923
2.250 0.6360 0.02274 0.01262 -0.1034 0.7725 0.8100
2.500 0.6622 0.02258 0.01264 -0.1026 0.7599 0.8353
2.750 0.6886 0.02230 0.01262 -0.1017 0.7468 1.0000
3.000 0.7227 0.02241 0.01273 -0.1029 0.7337 1.0000
3.250 0.7553 0.02251 0.01285 -0.1036 0.7197 1.0000
3.500 0.7864 0.02261 0.01302 -0.1039 0.7044 1.0000
3.750 0.8164 0.02267 0.01311 -0.1039 0.6875 1.0000
4.000 0.8464 0.02268 0.01315 -0.1036 0.6696 1.0000
4.250 0.8756 0.02270 0.01319 -0.1032 0.6507 1.0000
4.500 0.9019 0.02289 0.01350 -0.1025 0.6299 1.0000
4.750 0.9300 0.02302 0.01368 -0.1019 0.6101 1.0000
5.000 0.9555 0.02330 0.01405 -0.1011 0.5883 1.0000
5.250 0.9819 0.02354 0.01437 -0.1003 0.5662 1.0000
5.500 1.0060 0.02393 0.01493 -0.0993 0.5417 1.0000
5.750 1.0302 0.02432 0.01541 -0.0982 0.5160 1.0000
6.000 1.0534 0.02478 0.01596 -0.0971 0.4888 1.0000
6.250 1.0755 0.02531 0.01659 -0.0959 0.4596 1.0000
6.500 1.0971 0.02589 0.01725 -0.0946 0.4299 1.0000
6.750 1.1177 0.02656 0.01792 -0.0931 0.3994 1.0000
7.000 1.1371 0.02736 0.01880 -0.0916 0.3692 1.0000
7.250 1.1560 0.02831 0.01974 -0.0902 0.3411 1.0000
7.500 1.1741 0.02936 0.02083 -0.0887 0.3146 1.0000
7.750 1.1880 0.03051 0.02200 -0.0868 0.2837 1.0000
8.000 1.1947 0.03178 0.02315 -0.0843 0.2442 1.0000
8.250 1.1974 0.03331 0.02449 -0.0818 0.1968 1.0000
8.500 1.2006 0.03516 0.02615 -0.0794 0.1513 1.0000
8.750 1.2011 0.03739 0.02814 -0.0769 0.1030 1.0000
9.000 1.1978 0.04026 0.03070 -0.0742 0.0755 1.0000
9.250 1.1960 0.04318 0.03355 -0.0719 0.0622 1.0000
9.500 1.1937 0.04618 0.03649 -0.0699 0.0547 1.0000
9.750 1.1959 0.04889 0.03939 -0.0683 0.0482 1.0000
10.000 1.1945 0.05195 0.04250 -0.0670 0.0445 1.0000
10.250 1.1982 0.05481 0.04558 -0.0657 0.0415 1.0000
10.500 1.2026 0.05771 0.04869 -0.0645 0.0392 1.0000
10.750 1.2062 0.06073 0.05184 -0.0636 0.0372 1.0000
11.000 1.2102 0.06387 0.05506 -0.0628 0.0353 1.0000
11.250 1.2159 0.06711 0.05866 -0.0620 0.0334 1.0000
11.500 1.2180 0.07068 0.06252 -0.0616 0.0317 1.0000
11.750 1.2181 0.07452 0.06659 -0.0615 0.0306 1.0000
12.000 1.2168 0.07865 0.07110 -0.0617 0.0298 1.0000
12.250 1.2135 0.08315 0.07584 -0.0623 0.0293 1.0000
12.500 1.2077 0.08809 0.08101 -0.0633 0.0290 1.0000
12.750 1.1993 0.09352 0.08669 -0.0649 0.0287 1.0000
13.000 1.1881 0.09957 0.09299 -0.0674 0.0286 1.0000
13.250 1.1737 0.10644 0.10011 -0.0708 0.0287 1.0000
13.500 1.1560 0.11441 0.10835 -0.0754 0.0289 1.0000
13.750 1.1350 0.12397 0.11818 -0.0816 0.0293 1.0000
14.000 1.1102 0.13584 0.13016 -0.0895 0.0301 1.0000
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Polar data table (+)
Polar graphs
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