GOE 619 AIRFOIL (goe619-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 619 AIRFOIL (goe619-il) Reynolds number: 50,000 Max Cl/Cd: 33.26 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe619-il-50000-n5.txt Download as CSV file: xf-goe619-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 619 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3214 0.10399 0.09677 -0.0437 1.0000 0.0853
-9.250 -0.3312 0.10123 0.09410 -0.0429 1.0000 0.0851
-9.000 -0.3436 0.09864 0.09162 -0.0418 1.0000 0.0853
-8.750 -0.3591 0.09613 0.08922 -0.0404 1.0000 0.0855
-8.500 -0.3777 0.09369 0.08690 -0.0388 1.0000 0.0858
-8.250 -0.3981 0.09138 0.08471 -0.0369 1.0000 0.0857
-8.000 -0.4217 0.08881 0.08226 -0.0354 0.9994 0.0859
-7.750 -0.4242 0.08181 0.07525 -0.0437 0.9885 0.0867
-7.500 -0.4313 0.07082 0.06414 -0.0561 0.9758 0.0873
-7.250 -0.4431 0.05847 0.05133 -0.0680 0.9629 0.0880
-7.000 -0.4425 0.04956 0.04163 -0.0749 0.9526 0.0896
-6.750 -0.4171 0.04762 0.03963 -0.0766 0.9457 0.0917
-6.500 -0.3962 0.04536 0.03714 -0.0779 0.9378 0.0948
-6.250 -0.3728 0.04211 0.03338 -0.0803 0.9308 0.1000
-6.000 -0.3503 0.03965 0.03052 -0.0815 0.9235 0.1050
-5.750 -0.3241 0.03840 0.02917 -0.0823 0.9165 0.1104
-5.500 -0.2920 0.03602 0.02623 -0.0845 0.9117 0.1186
-5.250 -0.2723 0.03553 0.02583 -0.0836 0.9028 0.1252
-5.000 -0.2407 0.03443 0.02448 -0.0850 0.8973 0.1360
-4.750 -0.2163 0.03376 0.02357 -0.0851 0.8898 0.1474
-4.500 -0.1883 0.03366 0.02353 -0.0855 0.8830 0.1598
-4.250 -0.1528 0.03344 0.02323 -0.0872 0.8785 0.1754
-4.000 -0.1347 0.03318 0.02292 -0.0859 0.8692 0.1841
-3.750 -0.1013 0.03249 0.02198 -0.0871 0.8639 0.1957
-3.500 -0.0766 0.03221 0.02168 -0.0867 0.8568 0.2040
-3.250 -0.0485 0.03168 0.02097 -0.0870 0.8499 0.2135
-3.000 -0.0120 0.03121 0.02034 -0.0885 0.8455 0.2270
-2.750 0.0070 0.03112 0.02028 -0.0872 0.8365 0.2361
-2.500 0.0407 0.03070 0.01976 -0.0882 0.8309 0.2499
-2.250 0.0706 0.03047 0.01945 -0.0885 0.8248 0.2662
-2.000 0.0939 0.03040 0.01933 -0.0878 0.8165 0.2827
-1.750 0.1318 0.03008 0.01899 -0.0893 0.8119 0.3037
-1.500 0.1507 0.03018 0.01905 -0.0879 0.8024 0.3203
-1.250 0.1855 0.02998 0.01883 -0.0889 0.7967 0.3431
-1.000 0.2109 0.02999 0.01882 -0.0884 0.7886 0.3631
-0.750 0.2409 0.02990 0.01871 -0.0886 0.7811 0.3864
-0.500 0.2757 0.02966 0.01852 -0.0894 0.7751 0.4142
-0.250 0.2968 0.02972 0.01862 -0.0882 0.7650 0.4396
0.000 0.3388 0.02915 0.01809 -0.0899 0.7602 0.4725
0.250 0.3559 0.02921 0.01824 -0.0881 0.7482 0.4940
0.500 0.3902 0.02869 0.01784 -0.0886 0.7407 0.5267
0.750 0.4155 0.02833 0.01772 -0.0877 0.7304 0.5654
1.000 0.4397 0.02775 0.01763 -0.0862 0.7205 0.6457
1.250 0.5032 0.02664 0.01693 -0.0912 0.7120 1.0000
1.500 0.5255 0.02681 0.01690 -0.0901 0.6996 1.0000
1.750 0.5590 0.02660 0.01648 -0.0903 0.6898 1.0000
2.250 0.6093 0.02666 0.01627 -0.0883 0.6654 1.0000
2.500 0.6364 0.02662 0.01610 -0.0876 0.6537 1.0000
2.750 0.6703 0.02633 0.01569 -0.0878 0.6432 1.0000
3.000 0.6896 0.02656 0.01584 -0.0860 0.6291 1.0000
3.250 0.7112 0.02671 0.01593 -0.0846 0.6153 1.0000
3.500 0.7351 0.02679 0.01593 -0.0834 0.6019 1.0000
3.750 0.7608 0.02679 0.01585 -0.0825 0.5884 1.0000
4.000 0.7873 0.02677 0.01575 -0.0816 0.5746 1.0000
4.250 0.8091 0.02692 0.01585 -0.0802 0.5594 1.0000
4.500 0.8291 0.02716 0.01603 -0.0786 0.5435 1.0000
4.750 0.8495 0.02740 0.01622 -0.0770 0.5274 1.0000
5.000 0.8701 0.02767 0.01642 -0.0755 0.5114 1.0000
5.250 0.8908 0.02796 0.01665 -0.0741 0.4955 1.0000
5.500 0.9114 0.02830 0.01691 -0.0726 0.4797 1.0000
5.750 0.9318 0.02870 0.01724 -0.0712 0.4647 1.0000
6.000 0.9520 0.02915 0.01761 -0.0699 0.4504 1.0000
6.250 0.9731 0.02963 0.01803 -0.0687 0.4374 1.0000
6.500 0.9972 0.03008 0.01836 -0.0679 0.4259 1.0000
6.750 1.0163 0.03073 0.01899 -0.0666 0.4142 1.0000
7.000 1.0353 0.03145 0.01969 -0.0653 0.4035 1.0000
7.250 1.0622 0.03194 0.02008 -0.0651 0.3943 1.0000
7.500 1.0766 0.03288 0.02109 -0.0633 0.3847 1.0000
7.750 1.1048 0.03346 0.02161 -0.0634 0.3772 1.0000
8.000 1.1186 0.03450 0.02277 -0.0617 0.3693 1.0000
8.250 1.1465 0.03512 0.02335 -0.0617 0.3626 1.0000
8.500 1.1595 0.03624 0.02459 -0.0599 0.3557 1.0000
8.750 1.1791 0.03714 0.02556 -0.0590 0.3493 1.0000
9.000 1.2057 0.03792 0.02635 -0.0590 0.3438 1.0000
9.250 1.2127 0.03932 0.02792 -0.0566 0.3380 1.0000
9.500 1.2323 0.04022 0.02888 -0.0557 0.3321 1.0000
9.750 1.2500 0.04122 0.02994 -0.0547 0.3263 1.0000
10.000 1.2521 0.04278 0.03170 -0.0520 0.3202 1.0000
10.250 1.2757 0.04341 0.03233 -0.0514 0.3140 1.0000
10.500 1.2768 0.04507 0.03415 -0.0489 0.3084 1.0000
10.750 1.2781 0.04669 0.03592 -0.0465 0.3021 1.0000
11.000 1.3040 0.04702 0.03621 -0.0460 0.2953 1.0000
11.250 1.2857 0.04983 0.03930 -0.0425 0.2899 1.0000
11.500 1.2945 0.05102 0.04055 -0.0410 0.2831 1.0000
11.750 1.2996 0.05255 0.04216 -0.0393 0.2771 1.0000
12.000 1.2782 0.05614 0.04600 -0.0369 0.2715 1.0000
12.250 1.3138 0.05507 0.04479 -0.0364 0.2633 1.0000
12.500 1.2709 0.06083 0.05090 -0.0342 0.2585 1.0000
12.750 1.2524 0.06491 0.05515 -0.0332 0.2523 1.0000
13.000 1.2752 0.06469 0.05492 -0.0322 0.2454 1.0000
13.250 1.1901 0.07763 0.06820 -0.0340 0.2402 1.0000
13.500 1.1948 0.07969 0.07034 -0.0336 0.2337 1.0000
13.750 1.0770 0.10193 0.09275 -0.0421 0.2221 1.0000
14.000 1.1016 0.10075 0.09167 -0.0405 0.2173 1.0000
14.250 1.1619 0.09358 0.08455 -0.0361 0.2147 1.0000
14.500 1.0728 0.11287 0.10395 -0.0446 0.2007 1.0000
15.000 1.0624 0.12186 0.11309 -0.0476 0.1846 1.0000
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Polar data table (+)
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