WORTMANN FX 77-080 (fx77080-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: WORTMANN FX 77-080 (fx77080-il) Reynolds number: 50,000 Max Cl/Cd: 24.68 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx77080-il-50000.txt Download as CSV file: xf-fx77080-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: WORTMANN FX 77-080
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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5896 0.09411 0.08683 0.0215 1.0000 0.3537
-8.000 -0.5976 0.09091 0.08371 0.0212 1.0000 0.3641
-7.750 -0.7383 0.06117 0.05349 -0.0155 1.0000 0.1390
-7.500 -0.7323 0.05623 0.04835 -0.0150 1.0000 0.1373
-7.250 -0.7277 0.05130 0.04306 -0.0143 1.0000 0.1362
-7.000 -0.7216 0.04677 0.03801 -0.0131 1.0000 0.1371
-6.750 -0.7105 0.04317 0.03405 -0.0117 1.0000 0.1426
-6.500 -0.6949 0.04019 0.03077 -0.0102 1.0000 0.1490
-6.250 -0.6814 0.03669 0.02646 -0.0083 1.0000 0.1543
-6.000 -0.6620 0.03443 0.02410 -0.0070 1.0000 0.1673
-5.750 -0.6418 0.03197 0.02148 -0.0055 1.0000 0.1801
-5.500 -0.6221 0.02982 0.01904 -0.0039 1.0000 0.1993
-5.250 -0.6006 0.02792 0.01720 -0.0025 1.0000 0.2228
-5.000 -0.5802 0.02628 0.01555 -0.0009 1.0000 0.2535
-4.750 -0.5596 0.02478 0.01405 0.0008 1.0000 0.2887
-4.500 -0.5385 0.02347 0.01292 0.0024 1.0000 0.3260
-4.250 -0.5169 0.02226 0.01183 0.0040 1.0000 0.3661
-4.000 -0.4948 0.02109 0.01085 0.0056 1.0000 0.4113
-3.750 -0.4722 0.01991 0.00995 0.0072 1.0000 0.4685
-3.500 -0.4520 0.01871 0.00923 0.0097 1.0000 0.5474
-3.250 -0.4349 0.01762 0.00890 0.0140 1.0000 0.6667
-3.000 -0.3968 0.01767 0.00940 0.0173 1.0000 0.8287
-2.750 -0.2342 0.01909 0.00981 -0.0029 1.0000 0.9578
-2.500 -0.1358 0.01819 0.00835 -0.0181 1.0000 1.0000
-2.250 -0.1250 0.01758 0.00763 -0.0169 1.0000 1.0000
-2.000 -0.1128 0.01711 0.00707 -0.0154 1.0000 1.0000
-1.750 -0.0999 0.01675 0.00663 -0.0138 1.0000 1.0000
-1.500 -0.0863 0.01646 0.00625 -0.0120 1.0000 1.0000
-1.250 -0.0724 0.01623 0.00597 -0.0101 1.0000 1.0000
-1.000 -0.0583 0.01606 0.00574 -0.0081 1.0000 1.0000
-0.750 -0.0441 0.01592 0.00557 -0.0061 1.0000 1.0000
-0.500 -0.0296 0.01583 0.00545 -0.0040 1.0000 1.0000
-0.250 -0.0148 0.01578 0.00538 -0.0020 1.0000 1.0000
0.000 0.0000 0.01576 0.00536 0.0000 1.0000 1.0000
0.250 0.0149 0.01578 0.00538 0.0020 1.0000 1.0000
0.500 0.0296 0.01583 0.00545 0.0041 1.0000 1.0000
0.750 0.0442 0.01592 0.00557 0.0061 1.0000 1.0000
1.000 0.0584 0.01606 0.00574 0.0081 1.0000 1.0000
1.250 0.0725 0.01623 0.00596 0.0101 1.0000 1.0000
1.500 0.0864 0.01646 0.00625 0.0120 1.0000 1.0000
1.750 0.0999 0.01674 0.00662 0.0138 1.0000 1.0000
2.000 0.1130 0.01711 0.00707 0.0154 1.0000 1.0000
2.250 0.1251 0.01758 0.00763 0.0169 1.0000 1.0000
2.500 0.1359 0.01819 0.00835 0.0181 1.0000 1.0000
2.750 0.2344 0.01909 0.00981 0.0029 0.9576 1.0000
3.000 0.3973 0.01765 0.00938 -0.0173 0.8280 1.0000
3.250 0.4349 0.01762 0.00890 -0.0139 0.6666 1.0000
3.500 0.4520 0.01871 0.00923 -0.0097 0.5467 1.0000
3.750 0.4722 0.01991 0.00995 -0.0072 0.4682 1.0000
4.000 0.4948 0.02109 0.01085 -0.0056 0.4111 1.0000
4.250 0.5170 0.02226 0.01183 -0.0040 0.3659 1.0000
4.500 0.5385 0.02347 0.01292 -0.0024 0.3260 1.0000
4.750 0.5597 0.02479 0.01406 -0.0008 0.2888 1.0000
5.000 0.5803 0.02629 0.01556 0.0008 0.2540 1.0000
5.250 0.6006 0.02793 0.01720 0.0025 0.2228 1.0000
5.500 0.6221 0.02982 0.01905 0.0039 0.1993 1.0000
5.750 0.6418 0.03197 0.02147 0.0055 0.1802 1.0000
6.000 0.6619 0.03442 0.02411 0.0070 0.1671 1.0000
6.250 0.6811 0.03669 0.02649 0.0084 0.1540 1.0000
6.500 0.6948 0.04020 0.03079 0.0103 0.1488 1.0000
6.750 0.7104 0.04322 0.03411 0.0117 0.1427 1.0000
7.000 0.7215 0.04681 0.03806 0.0131 0.1371 1.0000
7.250 0.7273 0.05135 0.04312 0.0143 0.1361 1.0000
7.500 0.7322 0.05627 0.04839 0.0150 0.1373 1.0000
7.750 0.7380 0.06122 0.05355 0.0155 0.1389 1.0000
8.000 0.7417 0.06639 0.05889 0.0157 0.1397 1.0000
8.250 0.5906 0.09422 0.08694 -0.0215 0.3528 1.0000
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