FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-100 AIRFOIL (fx60100-il) Reynolds number: 100,000 Max Cl/Cd: 62.66 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60100-il-100000.txt Download as CSV file: xf-fx60100-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-100 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3514 0.09961 0.09477 -0.0345 1.0000 0.1037
-8.500 -0.3694 0.09778 0.09308 -0.0373 1.0000 0.1085
-8.250 -0.3983 0.09669 0.09218 -0.0393 1.0000 0.1092
-8.000 -0.3570 0.09152 0.08692 -0.0344 1.0000 0.1146
-7.750 -0.3678 0.08973 0.08525 -0.0340 1.0000 0.1204
-7.500 -0.3967 0.08874 0.08443 -0.0329 1.0000 0.1219
-7.250 -0.4286 0.08677 0.08263 -0.0365 1.0000 0.1227
-7.000 -0.3945 0.08407 0.07987 -0.0270 1.0000 0.1303
-6.750 -0.4133 0.08220 0.07813 -0.0270 1.0000 0.1338
-6.500 -0.4376 0.07790 0.07394 -0.0351 1.0000 0.1369
-6.250 -0.4218 0.07753 0.07360 -0.0240 1.0000 0.1435
-6.000 -0.4359 0.07228 0.06841 -0.0343 1.0000 0.1514
-5.750 -0.4289 0.07159 0.06777 -0.0262 1.0000 0.1555
-5.250 -0.3616 0.03899 0.03312 -0.0703 1.0000 0.0723
-5.000 -0.3244 0.03437 0.02748 -0.0731 1.0000 0.0585
-4.750 -0.2949 0.03021 0.02309 -0.0749 1.0000 0.0555
-4.500 -0.2625 0.02725 0.01959 -0.0763 1.0000 0.0528
-4.250 -0.2320 0.02510 0.01699 -0.0769 1.0000 0.0515
-4.000 -0.2042 0.02353 0.01518 -0.0769 1.0000 0.0517
-3.750 -0.1778 0.02229 0.01384 -0.0767 1.0000 0.0529
-3.500 -0.1518 0.02130 0.01281 -0.0765 1.0000 0.0556
-3.250 -0.1249 0.02038 0.01190 -0.0767 1.0000 0.0614
-3.000 -0.0964 0.01955 0.01109 -0.0774 1.0000 0.0755
-2.750 -0.0545 0.01723 0.01037 -0.0813 1.0000 0.4105
-2.500 -0.0290 0.01827 0.01160 -0.0808 0.9963 0.5032
-2.250 0.0088 0.01897 0.01224 -0.0824 0.9882 0.5438
-2.000 0.0479 0.01945 0.01262 -0.0845 0.9803 0.5738
-1.750 0.0883 0.01979 0.01288 -0.0868 0.9718 0.6014
-1.500 0.1257 0.01999 0.01300 -0.0887 0.9620 0.6267
-1.250 0.1660 0.02011 0.01308 -0.0911 0.9538 0.6464
-1.000 0.2040 0.02016 0.01307 -0.0931 0.9444 0.6658
-0.750 0.2400 0.02020 0.01307 -0.0948 0.9346 0.6852
-0.500 0.2845 0.02014 0.01300 -0.0978 0.9274 0.7046
-0.250 0.3191 0.02004 0.01289 -0.0990 0.9157 0.7221
0.000 0.3558 0.01993 0.01278 -0.1006 0.9048 0.7413
0.250 0.4026 0.01963 0.01252 -0.1036 0.8986 0.7626
0.500 0.4344 0.01952 0.01244 -0.1042 0.8871 0.7848
0.750 0.4686 0.01929 0.01228 -0.1050 0.8764 0.8076
1.000 0.5133 0.01869 0.01177 -0.1071 0.8699 0.8358
1.250 0.5404 0.01828 0.01149 -0.1062 0.8574 0.8716
1.500 0.5748 0.01774 0.01112 -0.1067 0.8453 1.0000
1.750 0.6194 0.01759 0.01091 -0.1101 0.8344 1.0000
2.000 0.6634 0.01725 0.01051 -0.1127 0.8244 1.0000
2.250 0.7020 0.01693 0.01015 -0.1140 0.8125 1.0000
2.500 0.7352 0.01675 0.00995 -0.1144 0.7985 1.0000
2.750 0.7678 0.01649 0.00967 -0.1144 0.7837 1.0000
3.000 0.7963 0.01632 0.00947 -0.1137 0.7649 1.0000
3.250 0.8256 0.01607 0.00923 -0.1129 0.7457 1.0000
3.500 0.8556 0.01583 0.00895 -0.1122 0.7272 1.0000
3.750 0.8827 0.01580 0.00890 -0.1114 0.7061 1.0000
4.000 0.9106 0.01573 0.00880 -0.1106 0.6846 1.0000
4.250 0.9367 0.01578 0.00885 -0.1096 0.6604 1.0000
4.500 0.9628 0.01586 0.00888 -0.1085 0.6348 1.0000
4.750 0.9883 0.01600 0.00896 -0.1075 0.6075 1.0000
5.000 1.0130 0.01624 0.00916 -0.1064 0.5791 1.0000
5.250 1.0370 0.01655 0.00942 -0.1052 0.5492 1.0000
5.500 1.0601 0.01692 0.00975 -0.1040 0.5174 1.0000
5.750 1.0825 0.01734 0.01010 -0.1027 0.4835 1.0000
6.000 1.1036 0.01784 0.01056 -0.1013 0.4453 1.0000
6.250 1.1232 0.01845 0.01104 -0.0997 0.4021 1.0000
6.500 1.1404 0.01926 0.01163 -0.0978 0.3521 1.0000
6.750 1.1558 0.02033 0.01241 -0.0958 0.2977 1.0000
7.000 1.1701 0.02166 0.01338 -0.0939 0.2501 1.0000
7.250 1.1852 0.02305 0.01452 -0.0921 0.2110 1.0000
7.500 1.2010 0.02449 0.01578 -0.0905 0.1821 1.0000
7.750 1.2180 0.02603 0.01718 -0.0890 0.1597 1.0000
8.000 1.2370 0.02754 0.01863 -0.0877 0.1435 1.0000
8.250 1.2580 0.02922 0.02026 -0.0868 0.1309 1.0000
8.500 1.2811 0.03109 0.02210 -0.0861 0.1207 1.0000
8.750 1.3049 0.03291 0.02380 -0.0857 0.1117 1.0000
9.000 1.3256 0.03451 0.02571 -0.0846 0.1048 1.0000
9.250 1.3522 0.03660 0.02778 -0.0846 0.0996 1.0000
9.500 1.3745 0.03903 0.03056 -0.0837 0.0960 1.0000
9.750 1.3947 0.04146 0.03337 -0.0826 0.0928 1.0000
10.000 1.4138 0.04386 0.03601 -0.0816 0.0895 1.0000
10.250 1.4310 0.04703 0.03932 -0.0808 0.0851 1.0000
10.500 1.4348 0.04897 0.04183 -0.0779 0.0808 1.0000
10.750 1.4471 0.04967 0.04236 -0.0769 0.0723 1.0000
11.000 1.4489 0.04966 0.04250 -0.0741 0.0650 1.0000
11.250 1.4494 0.05208 0.04520 -0.0716 0.0600 1.0000
11.500 1.4430 0.05399 0.04742 -0.0681 0.0550 1.0000
11.750 1.4383 0.05689 0.05033 -0.0657 0.0498 1.0000
12.000 1.4183 0.05983 0.05373 -0.0618 0.0474 1.0000
12.250 1.4051 0.06273 0.05687 -0.0594 0.0440 1.0000
12.500 1.4056 0.06621 0.06014 -0.0587 0.0397 1.0000
12.750 1.3842 0.07047 0.06481 -0.0573 0.0391 1.0000
13.000 1.3625 0.07556 0.07026 -0.0570 0.0387 1.0000
13.250 1.3400 0.08129 0.07630 -0.0578 0.0385 1.0000
13.500 1.3170 0.08767 0.08297 -0.0597 0.0385 1.0000
13.750 1.2940 0.09463 0.09017 -0.0626 0.0386 1.0000
14.000 1.2705 0.10229 0.09804 -0.0665 0.0389 1.0000
14.250 1.2485 0.11040 0.10633 -0.0711 0.0393 1.0000
14.500 1.2265 0.11924 0.11531 -0.0764 0.0397 1.0000
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Polar data table (+)
Polar graphs
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