GOE 121 (MVA H.1) AIRFOIL (goe121-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 121 (MVA H.1) AIRFOIL (goe121-il) Reynolds number: 100,000 Max Cl/Cd: 59.93 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe121-il-100000-n5.txt Download as CSV file: xf-goe121-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3455 0.09368 0.08878 -0.0226 1.0000 0.0617
-7.250 -0.3489 0.09232 0.08753 -0.0270 1.0000 0.0646
-7.000 -0.3447 0.09004 0.08530 -0.0335 1.0000 0.0651
-6.750 -0.3366 0.08693 0.08221 -0.0376 1.0000 0.0653
-6.500 -0.3335 0.08285 0.07821 -0.0342 1.0000 0.0659
-6.250 -0.3293 0.07982 0.07522 -0.0314 1.0000 0.0667
-6.000 -0.3247 0.07722 0.07266 -0.0304 1.0000 0.0676
-5.750 -0.3201 0.07496 0.07043 -0.0295 1.0000 0.0694
-5.500 -0.3136 0.07277 0.06826 -0.0302 1.0000 0.0732
-5.250 -0.2711 0.06848 0.06370 -0.0461 0.9971 0.0771
-5.000 -0.2331 0.05766 0.05260 -0.0556 0.9925 0.0536
-4.750 -0.2049 0.05408 0.04896 -0.0587 0.9874 0.0529
-4.250 -0.1173 0.04356 0.03787 -0.0742 0.9762 0.0533
-4.000 -0.0798 0.03963 0.03371 -0.0789 0.9695 0.0528
-3.750 -0.0373 0.03527 0.02900 -0.0845 0.9646 0.0524
-3.500 0.0032 0.03096 0.02425 -0.0891 0.9583 0.0524
-3.250 0.0471 0.02652 0.01908 -0.0939 0.9535 0.0538
-3.000 0.0837 0.02522 0.01766 -0.0964 0.9470 0.0558
-2.750 0.1228 0.02293 0.01491 -0.0992 0.9410 0.0563
-2.250 0.1985 0.01967 0.01096 -0.1033 0.9274 0.0581
-2.000 0.2373 0.01847 0.00950 -0.1052 0.9196 0.0600
-1.750 0.2739 0.01746 0.00830 -0.1066 0.9090 0.0625
-1.500 0.3080 0.01665 0.00737 -0.1074 0.8970 0.0658
-1.250 0.3405 0.01627 0.00699 -0.1080 0.8857 0.0705
-1.000 0.3725 0.01594 0.00663 -0.1085 0.8756 0.0803
-0.750 0.4026 0.01570 0.00635 -0.1086 0.8641 0.0976
-0.500 0.4310 0.01539 0.00595 -0.1084 0.8512 0.1158
-0.250 0.4588 0.01511 0.00570 -0.1081 0.8373 0.1306
0.000 0.4862 0.01489 0.00556 -0.1078 0.8221 0.1492
0.250 0.5136 0.01475 0.00546 -0.1073 0.8059 0.1698
0.500 0.5412 0.01459 0.00530 -0.1068 0.7888 0.1832
0.750 0.5688 0.01445 0.00512 -0.1063 0.7711 0.1922
1.000 0.5965 0.01434 0.00499 -0.1058 0.7529 0.2055
1.250 0.6239 0.01424 0.00491 -0.1053 0.7337 0.2316
1.500 0.6512 0.01413 0.00485 -0.1048 0.7135 0.2839
1.750 0.6710 0.01295 0.00490 -0.1029 0.6922 0.7636
2.000 0.6985 0.01286 0.00480 -0.1022 0.6677 1.0000
2.250 0.7244 0.01303 0.00484 -0.1015 0.6410 1.0000
2.500 0.7501 0.01322 0.00489 -0.1007 0.6104 1.0000
2.750 0.7753 0.01346 0.00491 -0.0998 0.5764 1.0000
3.000 0.7999 0.01379 0.00497 -0.0988 0.5449 1.0000
3.250 0.8246 0.01417 0.00514 -0.0980 0.5182 1.0000
3.500 0.8496 0.01455 0.00538 -0.0973 0.4953 1.0000
3.750 0.8745 0.01494 0.00565 -0.0967 0.4744 1.0000
4.000 0.8997 0.01529 0.00597 -0.0961 0.4543 1.0000
4.250 0.9248 0.01565 0.00628 -0.0955 0.4356 1.0000
4.500 0.9500 0.01600 0.00662 -0.0949 0.4194 1.0000
4.750 0.9751 0.01636 0.00698 -0.0944 0.4048 1.0000
5.000 1.0003 0.01673 0.00739 -0.0938 0.3916 1.0000
5.250 1.0252 0.01711 0.00780 -0.0932 0.3793 1.0000
5.500 1.0499 0.01752 0.00824 -0.0926 0.3675 1.0000
5.750 1.0747 0.01794 0.00872 -0.0920 0.3556 1.0000
6.000 1.0980 0.01841 0.00922 -0.0912 0.3369 1.0000
6.250 1.1197 0.01896 0.00969 -0.0902 0.3071 1.0000
6.500 1.1419 0.01948 0.01021 -0.0893 0.2749 1.0000
6.750 1.1641 0.02004 0.01075 -0.0885 0.2400 1.0000
7.000 1.1825 0.02100 0.01143 -0.0872 0.1886 1.0000
7.250 1.1997 0.02225 0.01243 -0.0858 0.1495 1.0000
7.500 1.2120 0.02421 0.01388 -0.0840 0.0734 1.0000
7.750 1.2279 0.02568 0.01528 -0.0824 0.0443 1.0000
8.000 1.2431 0.02715 0.01671 -0.0807 0.0375 1.0000
8.250 1.2582 0.02853 0.01825 -0.0788 0.0343 1.0000
8.500 1.2718 0.02995 0.01989 -0.0769 0.0314 1.0000
8.750 1.2817 0.03160 0.02170 -0.0747 0.0289 1.0000
9.000 1.2858 0.03364 0.02387 -0.0719 0.0273 1.0000
9.250 1.2938 0.03523 0.02563 -0.0694 0.0265 1.0000
9.500 1.2996 0.03687 0.02744 -0.0666 0.0258 1.0000
9.750 1.3049 0.03862 0.02933 -0.0639 0.0250 1.0000
10.000 1.3107 0.04045 0.03131 -0.0615 0.0240 1.0000
10.250 1.3165 0.04239 0.03338 -0.0593 0.0228 1.0000
10.500 1.3220 0.04444 0.03556 -0.0573 0.0217 1.0000
10.750 1.3276 0.04666 0.03793 -0.0555 0.0209 1.0000
11.000 1.3342 0.04906 0.04044 -0.0537 0.0204 1.0000
11.250 1.3412 0.05169 0.04319 -0.0521 0.0199 1.0000
11.500 1.3486 0.05466 0.04630 -0.0506 0.0195 1.0000
11.750 1.3550 0.05817 0.05000 -0.0491 0.0191 1.0000
12.000 1.3551 0.06176 0.05381 -0.0477 0.0189 1.0000
12.250 1.3493 0.06500 0.05735 -0.0465 0.0187 1.0000
12.500 1.3410 0.06858 0.06122 -0.0458 0.0185 1.0000
12.750 1.3312 0.07257 0.06549 -0.0456 0.0183 1.0000
13.000 1.3193 0.07699 0.07019 -0.0461 0.0182 1.0000
13.250 1.3061 0.08188 0.07534 -0.0472 0.0180 1.0000
13.500 1.2916 0.08726 0.08098 -0.0489 0.0179 1.0000
13.750 1.2758 0.09316 0.08713 -0.0514 0.0179 1.0000
14.000 1.2594 0.09960 0.09379 -0.0545 0.0179 1.0000
14.250 1.2422 0.10653 0.10093 -0.0583 0.0179 1.0000
14.500 1.2247 0.11403 0.10863 -0.0628 0.0180 1.0000
14.750 1.2074 0.12205 0.11683 -0.0678 0.0181 1.0000
15.000 1.1901 0.13063 0.12557 -0.0733 0.0182 1.0000
15.250 1.1732 0.13970 0.13478 -0.0793 0.0184 1.0000
15.500 1.1575 0.14909 0.14427 -0.0854 0.0186 1.0000
15.750 1.1433 0.15871 0.15398 -0.0916 0.0188 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 121 (MVA H.1) AIRFOIL (goe121-il)