GOE 121 (MVA H.1) AIRFOIL (goe121-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 121 (MVA H.1) AIRFOIL (goe121-il) Reynolds number: 200,000 Max Cl/Cd: 74.94 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe121-il-200000-n5.txt Download as CSV file: xf-goe121-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 121 (MVA H.1) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3624 0.09898 0.09535 -0.0212 1.0000 0.0284
-8.250 -0.3590 0.09583 0.09224 -0.0223 1.0000 0.0284
-8.000 -0.3612 0.09309 0.08956 -0.0244 1.0000 0.0298
-7.750 -0.3643 0.09110 0.08764 -0.0279 1.0000 0.0306
-7.250 -0.3578 0.08390 0.08049 -0.0276 1.0000 0.0298
-7.000 -0.3544 0.08065 0.07729 -0.0289 1.0000 0.0297
-6.750 -0.3515 0.07740 0.07407 -0.0305 1.0000 0.0298
-6.500 -0.3451 0.07407 0.07078 -0.0328 0.9990 0.0320
-6.000 -0.2903 0.06831 0.06495 -0.0413 0.9895 0.0371
-5.750 -0.2559 0.06243 0.05898 -0.0508 0.9825 0.0356
-5.500 -0.2192 0.05713 0.05357 -0.0596 0.9757 0.0359
-5.250 -0.1832 0.05244 0.04874 -0.0672 0.9676 0.0377
-5.000 -0.1435 0.04697 0.04310 -0.0754 0.9613 0.0373
-4.750 -0.1039 0.04165 0.03756 -0.0826 0.9543 0.0372
-4.500 -0.0616 0.03626 0.03189 -0.0897 0.9486 0.0381
-4.250 -0.0198 0.03019 0.02541 -0.0961 0.9418 0.0386
-4.000 0.0252 0.02167 0.01595 -0.1029 0.9372 0.0387
-3.750 0.0613 0.01850 0.01213 -0.1055 0.9295 0.0392
-3.500 0.0984 0.01675 0.00995 -0.1076 0.9222 0.0398
-3.250 0.1321 0.01560 0.00850 -0.1088 0.9121 0.0404
-3.000 0.1651 0.01485 0.00752 -0.1098 0.9011 0.0413
-2.750 0.1965 0.01409 0.00655 -0.1103 0.8888 0.0421
-2.500 0.2258 0.01336 0.00566 -0.1104 0.8747 0.0425
-2.250 0.2538 0.01281 0.00498 -0.1102 0.8600 0.0431
-2.000 0.2815 0.01237 0.00446 -0.1099 0.8458 0.0440
-1.750 0.3089 0.01203 0.00405 -0.1096 0.8310 0.0452
-1.500 0.3363 0.01175 0.00371 -0.1092 0.8154 0.0468
-1.250 0.3637 0.01152 0.00341 -0.1089 0.7994 0.0491
-1.000 0.3911 0.01133 0.00314 -0.1085 0.7839 0.0522
-0.750 0.4184 0.01116 0.00299 -0.1081 0.7686 0.0596
-0.500 0.4456 0.01113 0.00298 -0.1077 0.7529 0.0792
-0.250 0.4726 0.01119 0.00293 -0.1072 0.7365 0.0956
0.000 0.4994 0.01124 0.00290 -0.1067 0.7189 0.1088
0.250 0.5260 0.01126 0.00285 -0.1063 0.7003 0.1190
0.500 0.5527 0.01126 0.00280 -0.1059 0.6805 0.1271
0.750 0.5792 0.01126 0.00275 -0.1054 0.6557 0.1320
1.000 0.6051 0.01129 0.00271 -0.1049 0.6161 0.1419
1.250 0.6293 0.01146 0.00266 -0.1040 0.5634 0.1512
1.500 0.6537 0.01170 0.00269 -0.1032 0.5251 0.1578
1.750 0.6791 0.01189 0.00279 -0.1026 0.4987 0.1693
2.000 0.7050 0.01202 0.00294 -0.1022 0.4775 0.2026
2.250 0.7309 0.01211 0.00310 -0.1019 0.4575 0.2673
3.250 0.8299 0.01157 0.00366 -0.0990 0.3901 1.0000
3.500 0.8559 0.01183 0.00384 -0.0986 0.3747 1.0000
3.750 0.8818 0.01210 0.00404 -0.0982 0.3605 1.0000
4.000 0.9077 0.01237 0.00428 -0.0978 0.3484 1.0000
4.250 0.9334 0.01266 0.00453 -0.0974 0.3376 1.0000
4.500 0.9595 0.01291 0.00479 -0.0970 0.3278 1.0000
4.750 0.9853 0.01319 0.00507 -0.0966 0.3186 1.0000
5.000 1.0109 0.01349 0.00538 -0.0962 0.3083 1.0000
5.250 1.0358 0.01384 0.00570 -0.0957 0.2874 1.0000
5.500 1.0594 0.01432 0.00602 -0.0950 0.2477 1.0000
6.000 1.1035 0.01573 0.00698 -0.0935 0.1708 1.0000
6.250 1.1263 0.01635 0.00754 -0.0927 0.1521 1.0000
6.500 1.1499 0.01686 0.00806 -0.0921 0.1221 1.0000
6.750 1.1670 0.01823 0.00904 -0.0907 0.0692 1.0000
7.000 1.1858 0.01937 0.00999 -0.0894 0.0335 1.0000
7.250 1.2056 0.02037 0.01104 -0.0881 0.0274 1.0000
7.500 1.2269 0.02110 0.01190 -0.0871 0.0245 1.0000
7.750 1.2469 0.02197 0.01290 -0.0859 0.0220 1.0000
8.000 1.2646 0.02304 0.01411 -0.0844 0.0203 1.0000
8.250 1.2777 0.02454 0.01576 -0.0824 0.0189 1.0000
8.500 1.2953 0.02545 0.01681 -0.0809 0.0179 1.0000
8.750 1.3112 0.02649 0.01797 -0.0793 0.0167 1.0000
9.000 1.3247 0.02771 0.01930 -0.0773 0.0158 1.0000
9.250 1.3364 0.02904 0.02074 -0.0752 0.0151 1.0000
9.500 1.3464 0.03046 0.02224 -0.0729 0.0145 1.0000
9.750 1.3532 0.03193 0.02380 -0.0702 0.0141 1.0000
10.000 1.3584 0.03360 0.02555 -0.0674 0.0136 1.0000
10.250 1.3614 0.03566 0.02772 -0.0646 0.0131 1.0000
10.500 1.3662 0.03777 0.02993 -0.0622 0.0127 1.0000
10.750 1.3742 0.03929 0.03162 -0.0602 0.0122 1.0000
11.000 1.3809 0.04107 0.03357 -0.0583 0.0117 1.0000
11.250 1.3863 0.04314 0.03579 -0.0565 0.0114 1.0000
11.500 1.3907 0.04540 0.03822 -0.0547 0.0111 1.0000
11.750 1.3937 0.04784 0.04084 -0.0531 0.0109 1.0000
12.000 1.3952 0.05046 0.04365 -0.0516 0.0106 1.0000
12.250 1.3952 0.05329 0.04667 -0.0504 0.0104 1.0000
12.500 1.3934 0.05637 0.04994 -0.0494 0.0103 1.0000
12.750 1.3898 0.05974 0.05352 -0.0487 0.0101 1.0000
13.000 1.3842 0.06342 0.05740 -0.0484 0.0100 1.0000
13.250 1.3770 0.06745 0.06167 -0.0485 0.0099 1.0000
13.500 1.3681 0.07184 0.06627 -0.0490 0.0099 1.0000
13.750 1.3576 0.07663 0.07127 -0.0500 0.0098 1.0000
14.000 1.3456 0.08186 0.07671 -0.0516 0.0098 1.0000
14.250 1.3326 0.08750 0.08254 -0.0537 0.0097 1.0000
14.500 1.3186 0.09361 0.08885 -0.0564 0.0097 1.0000
14.750 1.3036 0.10023 0.09568 -0.0597 0.0097 1.0000
15.000 1.2874 0.10747 0.10312 -0.0635 0.0097 1.0000
15.250 1.2706 0.11527 0.11111 -0.0680 0.0097 1.0000
15.500 1.2523 0.12387 0.11991 -0.0732 0.0098 1.0000
15.750 1.2310 0.13380 0.13005 -0.0795 0.0099 1.0000
16.000 1.2050 0.14592 0.14236 -0.0875 0.0101 1.0000
16.250 1.1739 0.16126 0.15791 -0.0977 0.0104 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 121 (MVA H.1) AIRFOIL (goe121-il)