FX 60-100 (126) AIRFOIL (fx601001-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 60-100 (126) AIRFOIL (fx601001-il) Reynolds number: 100,000 Max Cl/Cd: 60.97 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx601001-il-100000.txt Download as CSV file: xf-fx601001-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3722 0.10326 0.09828 -0.0325 1.0000 0.1031
-9.000 -0.3741 0.10050 0.09559 -0.0338 1.0000 0.1076
-8.750 -0.4056 0.09891 0.09420 -0.0394 1.0000 0.1104
-8.500 -0.3767 0.09403 0.08926 -0.0350 1.0000 0.1139
-8.250 -0.3714 0.09141 0.08669 -0.0342 1.0000 0.1185
-8.000 -0.3949 0.08951 0.08496 -0.0364 1.0000 0.1235
-7.750 -0.4351 0.08731 0.08299 -0.0421 1.0000 0.1246
-7.500 -0.3875 0.08358 0.07913 -0.0329 1.0000 0.1301
-7.250 -0.3995 0.08173 0.07740 -0.0313 1.0000 0.1347
-7.000 -0.4374 0.07763 0.07351 -0.0403 1.0000 0.1393
-6.750 -0.4183 0.07640 0.07228 -0.0306 1.0000 0.1428
-6.500 -0.4220 0.07405 0.06999 -0.0301 1.0000 0.1485
-6.250 -0.4323 0.07012 0.06615 -0.0336 1.0000 0.1558
-6.000 -0.4165 0.04261 0.03716 -0.0686 1.0000 0.0712
-5.750 -0.3930 0.03763 0.03177 -0.0707 1.0000 0.0634
-5.500 -0.3614 0.03217 0.02543 -0.0743 1.0000 0.0591
-5.250 -0.3362 0.02988 0.02284 -0.0752 1.0000 0.0624
-5.000 -0.3072 0.02730 0.01975 -0.0763 1.0000 0.0644
-4.750 -0.2778 0.02509 0.01707 -0.0769 1.0000 0.0662
-4.500 -0.2489 0.02311 0.01470 -0.0772 1.0000 0.0702
-4.250 -0.2230 0.02181 0.01336 -0.0774 1.0000 0.0809
-4.000 -0.1957 0.02024 0.01179 -0.0775 1.0000 0.0957
-3.500 -0.1425 0.01819 0.01017 -0.0787 1.0000 0.1611
-3.250 -0.1129 0.01716 0.00954 -0.0800 1.0000 0.2279
-3.000 -0.0892 0.01720 0.01056 -0.0797 1.0000 0.4687
-2.750 -0.0506 0.01822 0.01145 -0.0814 0.9917 0.5199
-2.500 -0.0089 0.01883 0.01191 -0.0840 0.9834 0.5541
-2.250 0.0281 0.01926 0.01230 -0.0855 0.9734 0.5828
-2.000 0.0631 0.01972 0.01280 -0.0863 0.9637 0.6131
-1.750 0.1011 0.02009 0.01318 -0.0876 0.9555 0.6427
-1.500 0.1341 0.02017 0.01323 -0.0885 0.9445 0.6603
-1.250 0.1718 0.02018 0.01315 -0.0905 0.9350 0.6740
-1.000 0.2154 0.02011 0.01297 -0.0937 0.9273 0.6864
-0.750 0.2498 0.02006 0.01288 -0.0951 0.9165 0.6965
-0.500 0.2891 0.01996 0.01274 -0.0973 0.9080 0.7060
-0.250 0.3310 0.01982 0.01255 -0.1000 0.8999 0.7167
0.000 0.3666 0.01975 0.01246 -0.1015 0.8898 0.7288
0.250 0.4124 0.01941 0.01212 -0.1044 0.8838 0.7401
0.500 0.4466 0.01917 0.01190 -0.1054 0.8721 0.7474
0.750 0.4839 0.01890 0.01161 -0.1068 0.8612 0.7552
1.000 0.5240 0.01845 0.01119 -0.1085 0.8530 0.7623
1.250 0.5584 0.01823 0.01099 -0.1094 0.8422 0.7702
1.500 0.5883 0.01806 0.01086 -0.1094 0.8305 0.7778
1.750 0.6209 0.01788 0.01070 -0.1099 0.8196 0.7863
2.000 0.6546 0.01755 0.01044 -0.1103 0.8099 0.7954
2.250 0.6847 0.01731 0.01024 -0.1101 0.7977 0.8057
2.500 0.7135 0.01702 0.01001 -0.1095 0.7834 0.8172
2.750 0.7413 0.01670 0.00975 -0.1086 0.7686 0.8313
3.000 0.7669 0.01640 0.00959 -0.1073 0.7538 0.8505
3.250 0.7901 0.01610 0.00945 -0.1056 0.7392 0.8863
3.500 0.8221 0.01598 0.00942 -0.1061 0.7229 1.0000
3.750 0.8545 0.01607 0.00951 -0.1071 0.7055 1.0000
4.000 0.8859 0.01616 0.00960 -0.1076 0.6878 1.0000
4.250 0.9162 0.01623 0.00963 -0.1076 0.6694 1.0000
4.500 0.9443 0.01635 0.00973 -0.1072 0.6484 1.0000
4.750 0.9716 0.01648 0.00986 -0.1066 0.6259 1.0000
5.000 0.9975 0.01670 0.01007 -0.1058 0.6007 1.0000
5.250 1.0229 0.01694 0.01029 -0.1048 0.5727 1.0000
5.500 1.0471 0.01722 0.01051 -0.1036 0.5408 1.0000
5.750 1.0700 0.01755 0.01079 -0.1022 0.5049 1.0000
6.000 1.0918 0.01793 0.01110 -0.1007 0.4659 1.0000
6.250 1.1131 0.01844 0.01152 -0.0993 0.4268 1.0000
6.500 1.1338 0.01911 0.01202 -0.0979 0.3904 1.0000
6.750 1.1525 0.01988 0.01260 -0.0963 0.3516 1.0000
7.000 1.1687 0.02067 0.01325 -0.0946 0.3087 1.0000
7.250 1.1856 0.02146 0.01397 -0.0931 0.2695 1.0000
7.500 1.2025 0.02231 0.01480 -0.0915 0.2312 1.0000
7.750 1.2108 0.02394 0.01601 -0.0891 0.1478 1.0000
8.000 1.2098 0.02721 0.01847 -0.0853 0.0777 1.0000
8.250 1.2178 0.02942 0.02063 -0.0825 0.0652 1.0000
8.500 1.2276 0.03159 0.02278 -0.0800 0.0584 1.0000
8.750 1.2409 0.03341 0.02462 -0.0782 0.0524 1.0000
9.000 1.2586 0.03605 0.02726 -0.0769 0.0486 1.0000
9.250 1.2824 0.03840 0.02979 -0.0759 0.0462 1.0000
9.500 1.3072 0.04114 0.03274 -0.0753 0.0444 1.0000
9.750 1.3281 0.04396 0.03577 -0.0743 0.0426 1.0000
10.000 1.3464 0.04706 0.03897 -0.0736 0.0404 1.0000
10.250 1.3621 0.05254 0.04470 -0.0730 0.0390 1.0000
10.500 1.3681 0.05612 0.04869 -0.0709 0.0388 1.0000
10.750 1.3704 0.06034 0.05331 -0.0687 0.0389 1.0000
11.000 1.3695 0.06464 0.05800 -0.0666 0.0391 1.0000
11.250 1.3627 0.06756 0.06125 -0.0636 0.0393 1.0000
11.500 1.3492 0.07024 0.06429 -0.0603 0.0397 1.0000
11.750 1.3311 0.07338 0.06776 -0.0576 0.0401 1.0000
12.000 1.3070 0.07750 0.07220 -0.0561 0.0407 1.0000
12.250 1.2751 0.08290 0.07798 -0.0563 0.0414 1.0000
12.500 1.2408 0.08969 0.08510 -0.0587 0.0420 1.0000
12.750 1.2067 0.09781 0.09350 -0.0630 0.0427 1.0000
13.000 1.1712 0.10761 0.10353 -0.0695 0.0434 1.0000
13.250 1.1323 0.12010 0.11620 -0.0788 0.0444 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 60-100 (126) AIRFOIL (fx601001-il)