GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 457 AIRFOIL (goe457-il) Reynolds number: 200,000 Max Cl/Cd: 73.57 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe457-il-200000.txt Download as CSV file: xf-goe457-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 457 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3619 0.08947 0.08615 -0.0206 1.0000 0.0396
-7.250 -0.3716 0.08777 0.08452 -0.0186 1.0000 0.0404
-7.000 -0.3798 0.08580 0.08261 -0.0173 1.0000 0.0414
-6.750 -0.3839 0.08337 0.08022 -0.0176 1.0000 0.0424
-6.500 -0.3842 0.08046 0.07733 -0.0242 1.0000 0.0444
-6.250 -0.3730 0.07665 0.07343 -0.0319 1.0000 0.0450
-6.000 -0.3775 0.07189 0.06876 -0.0286 1.0000 0.0459
-5.750 -0.3743 0.06972 0.06662 -0.0253 1.0000 0.0469
-5.500 -0.3673 0.06727 0.06418 -0.0245 1.0000 0.0482
-5.250 -0.3571 0.06440 0.06129 -0.0253 1.0000 0.0500
-5.000 -0.3175 0.05917 0.05565 -0.0376 1.0000 0.0562
-4.750 -0.3092 0.05361 0.05009 -0.0382 1.0000 0.0574
-4.500 -0.3000 0.05145 0.04801 -0.0364 0.9999 0.0587
-4.250 -0.2529 0.04832 0.04463 -0.0432 0.9955 0.0680
-4.000 -0.2164 0.04271 0.03881 -0.0489 0.9911 0.0716
-3.750 -0.1809 0.04014 0.03619 -0.0522 0.9863 0.0760
-3.500 -0.1352 0.03554 0.03110 -0.0581 0.9817 0.0857
-3.250 -0.0995 0.03354 0.02906 -0.0607 0.9758 0.0934
-3.000 -0.0564 0.03037 0.02559 -0.0650 0.9718 0.1031
-2.750 -0.0204 0.02804 0.02295 -0.0673 0.9647 0.1160
-2.500 0.0276 0.02183 0.01544 -0.0685 0.9615 0.0669
-2.250 0.0722 0.01950 0.01272 -0.0714 0.9580 0.0642
-2.000 0.1072 0.01818 0.01113 -0.0723 0.9495 0.0649
-1.750 0.1531 0.01699 0.00971 -0.0753 0.9448 0.0666
-1.500 0.1892 0.01597 0.00857 -0.0763 0.9359 0.0675
-1.250 0.2359 0.01503 0.00754 -0.0795 0.9308 0.0691
-1.000 0.2723 0.01391 0.00651 -0.0807 0.9213 0.0737
-0.750 0.3212 0.01306 0.00570 -0.0844 0.9156 0.0811
-0.500 0.3567 0.01223 0.00496 -0.0852 0.9029 0.0908
-0.250 0.3910 0.01103 0.00427 -0.0861 0.8882 0.2123
0.000 0.4510 0.00877 0.00398 -0.0923 0.8722 1.0000
0.250 0.4832 0.00859 0.00361 -0.0924 0.8405 1.0000
0.500 0.5152 0.00846 0.00324 -0.0924 0.7886 1.0000
0.750 0.5475 0.00856 0.00291 -0.0925 0.7258 1.0000
1.000 0.5722 0.00891 0.00286 -0.0914 0.6754 1.0000
1.250 0.5955 0.00927 0.00291 -0.0902 0.6381 1.0000
1.500 0.6189 0.00960 0.00300 -0.0890 0.6095 1.0000
1.750 0.6427 0.00993 0.00312 -0.0880 0.5867 1.0000
2.000 0.6669 0.01023 0.00327 -0.0871 0.5675 1.0000
2.250 0.6915 0.01054 0.00343 -0.0863 0.5510 1.0000
2.500 0.7162 0.01085 0.00362 -0.0856 0.5357 1.0000
2.750 0.7410 0.01115 0.00383 -0.0849 0.5213 1.0000
3.000 0.7659 0.01146 0.00404 -0.0842 0.5075 1.0000
3.250 0.7906 0.01175 0.00426 -0.0834 0.4936 1.0000
3.500 0.8151 0.01202 0.00446 -0.0827 0.4796 1.0000
3.750 0.8394 0.01228 0.00468 -0.0819 0.4653 1.0000
4.000 0.8635 0.01251 0.00486 -0.0810 0.4508 1.0000
4.250 0.8873 0.01273 0.00504 -0.0801 0.4361 1.0000
4.500 0.9110 0.01293 0.00524 -0.0792 0.4211 1.0000
4.750 0.9346 0.01314 0.00542 -0.0783 0.4062 1.0000
5.000 0.9584 0.01337 0.00563 -0.0774 0.3928 1.0000
5.250 0.9822 0.01363 0.00587 -0.0766 0.3803 1.0000
5.500 1.0061 0.01386 0.00615 -0.0758 0.3672 1.0000
5.750 1.0296 0.01410 0.00642 -0.0749 0.3535 1.0000
6.000 1.0527 0.01436 0.00669 -0.0739 0.3389 1.0000
6.250 1.0754 0.01463 0.00697 -0.0729 0.3232 1.0000
6.500 1.0977 0.01493 0.00728 -0.0719 0.3057 1.0000
6.750 1.1197 0.01522 0.00760 -0.0708 0.2840 1.0000
7.000 1.1399 0.01563 0.00794 -0.0695 0.2532 1.0000
7.250 1.1578 0.01629 0.00840 -0.0679 0.2081 1.0000
7.500 1.1741 0.01721 0.00906 -0.0661 0.1635 1.0000
7.750 1.1911 0.01812 0.00981 -0.0645 0.1409 1.0000
8.000 1.2086 0.01895 0.01059 -0.0629 0.1282 1.0000
8.250 1.2266 0.01971 0.01136 -0.0614 0.1188 1.0000
8.500 1.2429 0.02058 0.01219 -0.0597 0.1116 1.0000
8.750 1.2625 0.02115 0.01290 -0.0584 0.1058 1.0000
9.000 1.2797 0.02187 0.01363 -0.0569 0.0993 1.0000
9.250 1.2987 0.02242 0.01430 -0.0556 0.0938 1.0000
9.500 1.3175 0.02293 0.01488 -0.0543 0.0869 1.0000
9.750 1.3365 0.02342 0.01549 -0.0530 0.0805 1.0000
10.000 1.3554 0.02391 0.01605 -0.0518 0.0698 1.0000
10.250 1.3720 0.02453 0.01670 -0.0501 0.0564 1.0000
10.500 1.3826 0.02552 0.01762 -0.0476 0.0408 1.0000
10.750 1.3891 0.02674 0.01879 -0.0446 0.0317 1.0000
11.000 1.3946 0.02801 0.02012 -0.0416 0.0284 1.0000
11.250 1.3977 0.02946 0.02163 -0.0384 0.0264 1.0000
11.500 1.3983 0.03111 0.02341 -0.0353 0.0252 1.0000
11.750 1.3995 0.03277 0.02524 -0.0326 0.0243 1.0000
12.000 1.3987 0.03467 0.02729 -0.0300 0.0234 1.0000
12.250 1.3967 0.03678 0.02954 -0.0278 0.0225 1.0000
12.500 1.3926 0.03923 0.03213 -0.0259 0.0219 1.0000
12.750 1.3875 0.04196 0.03501 -0.0244 0.0215 1.0000
13.000 1.3812 0.04509 0.03827 -0.0235 0.0212 1.0000
13.250 1.3731 0.04869 0.04200 -0.0232 0.0208 1.0000
13.500 1.3671 0.05232 0.04576 -0.0232 0.0206 1.0000
13.750 1.3595 0.05634 0.04990 -0.0236 0.0203 1.0000
14.000 1.3533 0.06032 0.05400 -0.0239 0.0201 1.0000
14.250 1.3481 0.06425 0.05807 -0.0243 0.0199 1.0000
14.500 1.3439 0.06817 0.06211 -0.0246 0.0198 1.0000
14.750 1.3404 0.07213 0.06625 -0.0254 0.0197 1.0000
15.000 1.3360 0.07632 0.07061 -0.0264 0.0196 1.0000
15.250 1.3303 0.08084 0.07531 -0.0277 0.0196 1.0000
15.500 1.3235 0.08565 0.08029 -0.0292 0.0195 1.0000
15.750 1.3148 0.09088 0.08571 -0.0312 0.0195 1.0000
16.000 1.3053 0.09643 0.09145 -0.0335 0.0195 1.0000
16.250 1.2943 0.10240 0.09760 -0.0363 0.0196 1.0000
16.500 1.2818 0.10883 0.10422 -0.0395 0.0196 1.0000
16.750 1.2676 0.11585 0.11142 -0.0434 0.0196 1.0000
17.000 1.2528 0.12323 0.11898 -0.0477 0.0197 1.0000
17.250 1.2369 0.13113 0.12705 -0.0524 0.0197 1.0000
17.500 1.2228 0.13899 0.13510 -0.0574 0.0198 1.0000
17.750 1.2094 0.14712 0.14338 -0.0628 0.0200 1.0000
18.000 1.1911 0.15714 0.15358 -0.0697 0.0203 1.0000
18.250 1.0645 0.20810 0.20466 -0.0991 0.0253 1.0000
18.500 1.0533 0.22286 0.21935 -0.1052 0.0316 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 457 AIRFOIL (goe457-il)