GOE 411 AIRFOIL (goe411-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 411 AIRFOIL (goe411-il) Reynolds number: 100,000 Max Cl/Cd: 40.75 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe411-il-100000.txt Download as CSV file: xf-goe411-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 411 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.6285 0.10352 0.09861 -0.0379 1.0000 0.0841
-12.000 -0.6646 0.09045 0.08557 -0.0452 1.0000 0.0747
-11.750 -0.6939 0.08304 0.07818 -0.0498 1.0000 0.0750
-11.500 -0.7234 0.07705 0.07212 -0.0527 1.0000 0.0753
-11.250 -0.7552 0.07211 0.06707 -0.0537 1.0000 0.0753
-11.000 -0.7864 0.06828 0.06310 -0.0525 1.0000 0.0750
-10.750 -0.8217 0.06530 0.05994 -0.0488 1.0000 0.0745
-10.500 -0.8556 0.06341 0.05785 -0.0428 1.0000 0.0740
-10.250 -0.8575 0.05824 0.05248 -0.0401 1.0000 0.0651
-10.000 -0.8915 0.05541 0.04918 -0.0330 1.0000 0.0621
-9.750 -0.8960 0.05283 0.04623 -0.0284 1.0000 0.0577
-9.500 -0.8975 0.04885 0.04201 -0.0249 1.0000 0.0548
-9.250 -0.9061 0.04578 0.03840 -0.0195 1.0000 0.0515
-9.000 -0.9110 0.04528 0.03726 -0.0132 1.0000 0.0488
-8.750 -0.8999 0.04245 0.03415 -0.0106 1.0000 0.0482
-8.500 -0.8891 0.03984 0.03125 -0.0078 1.0000 0.0481
-8.250 -0.8737 0.03731 0.02843 -0.0056 1.0000 0.0480
-8.000 -0.8500 0.03453 0.02547 -0.0048 1.0000 0.0474
-7.750 -0.8218 0.03202 0.02282 -0.0044 1.0000 0.0467
-7.500 -0.7951 0.03002 0.02075 -0.0038 1.0000 0.0464
-7.250 -0.7742 0.02843 0.01914 -0.0023 1.0000 0.0465
-7.000 -0.7606 0.02714 0.01783 0.0003 1.0000 0.0470
-6.750 -0.7518 0.02603 0.01667 0.0038 1.0000 0.0477
-6.500 -0.7448 0.02507 0.01564 0.0076 1.0000 0.0488
-6.250 -0.7420 0.02388 0.01442 0.0118 1.0000 0.0521
-6.000 -0.7336 0.02305 0.01347 0.0153 1.0000 0.0554
-5.750 -0.7234 0.02230 0.01258 0.0186 1.0000 0.0584
-5.500 -0.7137 0.02143 0.01165 0.0220 1.0000 0.0638
-5.250 -0.7037 0.02053 0.01078 0.0253 1.0000 0.0776
-5.000 -0.7041 0.01874 0.00997 0.0298 1.0000 0.2150
-4.750 -0.6999 0.01764 0.00957 0.0338 1.0000 0.3264
-4.500 -0.6928 0.01705 0.00943 0.0377 1.0000 0.4158
-4.250 -0.6831 0.01674 0.00934 0.0412 1.0000 0.4779
-4.000 -0.6710 0.01656 0.00921 0.0443 1.0000 0.5233
-3.750 -0.6592 0.01640 0.00916 0.0475 1.0000 0.5679
-3.500 -0.6475 0.01625 0.00917 0.0509 1.0000 0.6141
-3.250 -0.6319 0.01613 0.00917 0.0535 1.0000 0.6500
-3.000 -0.6187 0.01616 0.00940 0.0569 1.0000 0.6972
-2.750 -0.6025 0.01626 0.00959 0.0596 1.0000 0.7361
-2.500 -0.5823 0.01631 0.00965 0.0610 1.0000 0.7637
-2.250 -0.5597 0.01642 0.00977 0.0619 1.0000 0.7920
-2.000 -0.5324 0.01667 0.01002 0.0620 1.0000 0.8213
-1.750 -0.4982 0.01711 0.01045 0.0606 1.0000 0.8499
-1.500 -0.4378 0.01805 0.01128 0.0541 0.9947 0.8758
-1.250 -0.3699 0.01919 0.01230 0.0466 0.9874 0.8999
-1.000 -0.2903 0.02043 0.01338 0.0368 0.9825 0.9144
-0.750 -0.2143 0.02115 0.01400 0.0271 0.9753 0.9223
-0.500 -0.1395 0.02171 0.01447 0.0175 0.9690 0.9320
-0.250 -0.0732 0.02197 0.01469 0.0094 0.9593 0.9430
0.000 -0.0001 0.02207 0.01477 0.0000 0.9510 0.9509
0.250 0.0732 0.02197 0.01469 -0.0094 0.9430 0.9593
0.500 0.1397 0.02171 0.01447 -0.0175 0.9319 0.9690
0.750 0.2145 0.02114 0.01400 -0.0271 0.9222 0.9752
1.000 0.2906 0.02043 0.01338 -0.0368 0.9144 0.9826
1.500 0.4380 0.01805 0.01127 -0.0542 0.8759 0.9947
1.750 0.4983 0.01711 0.01042 -0.0606 0.8498 1.0000
2.000 0.5325 0.01666 0.01002 -0.0620 0.8211 1.0000
2.250 0.5598 0.01642 0.00977 -0.0620 0.7921 1.0000
2.500 0.5824 0.01630 0.00965 -0.0610 0.7637 1.0000
2.750 0.6025 0.01625 0.00961 -0.0596 0.7359 1.0000
3.000 0.6189 0.01616 0.00940 -0.0569 0.6976 1.0000
3.250 0.6317 0.01613 0.00916 -0.0534 0.6487 1.0000
3.500 0.6479 0.01625 0.00917 -0.0509 0.6151 1.0000
3.750 0.6587 0.01640 0.00915 -0.0474 0.5657 1.0000
4.000 0.6709 0.01656 0.00921 -0.0443 0.5225 1.0000
4.250 0.6825 0.01675 0.00932 -0.0411 0.4747 1.0000
4.500 0.6929 0.01705 0.00943 -0.0377 0.4157 1.0000
4.750 0.6993 0.01767 0.00957 -0.0337 0.3211 1.0000
5.000 0.7037 0.01877 0.00997 -0.0297 0.2113 1.0000
5.250 0.7042 0.02048 0.01075 -0.0254 0.0790 1.0000
5.500 0.7135 0.02146 0.01168 -0.0219 0.0635 1.0000
5.750 0.7233 0.02231 0.01259 -0.0186 0.0582 1.0000
6.000 0.7334 0.02307 0.01349 -0.0153 0.0544 1.0000
6.250 0.7419 0.02389 0.01444 -0.0118 0.0518 1.0000
6.500 0.7448 0.02507 0.01564 -0.0076 0.0488 1.0000
6.750 0.7518 0.02603 0.01668 -0.0038 0.0476 1.0000
7.000 0.7606 0.02714 0.01783 -0.0003 0.0470 1.0000
7.250 0.7743 0.02841 0.01912 0.0023 0.0466 1.0000
7.500 0.7950 0.03001 0.02075 0.0038 0.0464 1.0000
7.750 0.8221 0.03206 0.02287 0.0044 0.0467 1.0000
8.000 0.8497 0.03448 0.02542 0.0048 0.0474 1.0000
8.250 0.8724 0.03716 0.02829 0.0059 0.0478 1.0000
8.500 0.8862 0.03940 0.03086 0.0084 0.0477 1.0000
8.750 0.8989 0.04230 0.03402 0.0108 0.0481 1.0000
9.000 0.9106 0.04555 0.03753 0.0132 0.0487 1.0000
9.250 0.9074 0.04571 0.03830 0.0193 0.0513 1.0000
9.500 0.8973 0.04891 0.04206 0.0250 0.0548 1.0000
9.750 0.8965 0.05274 0.04614 0.0283 0.0577 1.0000
10.000 0.8935 0.05537 0.04912 0.0327 0.0620 1.0000
10.250 0.8557 0.05842 0.05268 0.0403 0.0654 1.0000
10.500 0.8344 0.06140 0.05583 0.0453 0.0677 1.0000
10.750 0.8194 0.06532 0.05997 0.0490 0.0746 1.0000
11.000 0.7886 0.06816 0.06297 0.0524 0.0750 1.0000
11.250 0.7550 0.07216 0.06712 0.0537 0.0753 1.0000
11.500 0.7253 0.07693 0.07200 0.0527 0.0753 1.0000
11.750 0.6962 0.08281 0.07794 0.0499 0.0750 1.0000
12.000 0.6625 0.09093 0.08606 0.0447 0.0748 1.0000
12.250 0.6079 0.10806 0.10310 0.0330 0.0826 1.0000
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Polar data table (+)
Polar graphs
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