GOE 362 AIRFOIL (goe362-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: GOE 362 AIRFOIL (goe362-il) Reynolds number: 1,000,000 Max Cl/Cd: 131.44 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe362-il-1000000.txt Download as CSV file: xf-goe362-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 362 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2507 0.08513 0.08363 -0.0268 1.0000 0.0098
-8.500 -0.2491 0.08175 0.08026 -0.0272 1.0000 0.0099
-8.250 -0.3245 0.09278 0.09118 -0.0230 1.0000 0.0098
-8.000 -0.3218 0.09004 0.08847 -0.0236 1.0000 0.0098
-7.750 -0.3236 0.08725 0.08571 -0.0236 1.0000 0.0099
-7.500 -0.3280 0.08531 0.08381 -0.0221 1.0000 0.0098
-5.500 -0.0993 0.01768 0.01476 -0.1017 0.9504 0.0103
-5.250 -0.0733 0.01564 0.01240 -0.1021 0.9387 0.0107
-5.000 -0.0477 0.01426 0.01074 -0.1019 0.9263 0.0112
-4.750 -0.0223 0.01283 0.00900 -0.1016 0.9123 0.0116
-4.500 0.0034 0.01157 0.00744 -0.1011 0.8968 0.0120
-4.250 0.0293 0.01061 0.00622 -0.1005 0.8805 0.0124
-4.000 0.0558 0.01014 0.00556 -0.1000 0.8636 0.0131
-3.750 0.0817 0.00917 0.00431 -0.0994 0.8462 0.0140
-3.500 0.1079 0.00870 0.00370 -0.0989 0.8289 0.0150
-3.250 0.1346 0.00845 0.00334 -0.0984 0.8121 0.0159
-2.750 0.1883 0.00805 0.00271 -0.0975 0.7787 0.0179
-2.500 0.2150 0.00760 0.00214 -0.0971 0.7621 0.0204
-2.250 0.2420 0.00753 0.00200 -0.0967 0.7448 0.0228
-2.000 0.2689 0.00731 0.00167 -0.0962 0.7272 0.0255
-1.750 0.2959 0.00720 0.00149 -0.0959 0.7099 0.0298
-1.250 0.3499 0.00703 0.00118 -0.0951 0.6775 0.0413
-1.000 0.3770 0.00694 0.00105 -0.0948 0.6626 0.0523
-0.750 0.4041 0.00688 0.00102 -0.0945 0.6481 0.0797
-0.500 0.4313 0.00690 0.00101 -0.0942 0.6341 0.0960
0.000 0.4857 0.00691 0.00096 -0.0937 0.6057 0.1177
0.500 0.5400 0.00691 0.00094 -0.0931 0.5747 0.1507
0.750 0.5670 0.00688 0.00098 -0.0929 0.5563 0.2088
1.250 0.6186 0.00646 0.00115 -0.0922 0.5117 0.5199
1.500 0.6507 0.00561 0.00124 -0.0933 0.4874 1.0000
1.750 0.6774 0.00578 0.00129 -0.0930 0.4671 1.0000
2.000 0.7041 0.00595 0.00136 -0.0926 0.4511 1.0000
2.250 0.7309 0.00610 0.00143 -0.0923 0.4381 1.0000
2.500 0.7577 0.00624 0.00152 -0.0920 0.4271 1.0000
2.750 0.7845 0.00639 0.00160 -0.0917 0.4174 1.0000
3.000 0.8115 0.00651 0.00169 -0.0915 0.4079 1.0000
3.250 0.8383 0.00666 0.00179 -0.0912 0.3974 1.0000
3.500 0.8648 0.00683 0.00190 -0.0909 0.3864 1.0000
3.750 0.8917 0.00696 0.00200 -0.0906 0.3750 1.0000
4.000 0.9184 0.00710 0.00211 -0.0903 0.3643 1.0000
4.250 0.9450 0.00725 0.00222 -0.0900 0.3529 1.0000
4.500 0.9714 0.00742 0.00236 -0.0897 0.3416 1.0000
4.750 0.9977 0.00760 0.00250 -0.0894 0.3276 1.0000
5.000 1.0239 0.00779 0.00264 -0.0891 0.3131 1.0000
5.250 1.0500 0.00800 0.00280 -0.0887 0.2989 1.0000
5.500 1.0759 0.00822 0.00299 -0.0884 0.2831 1.0000
5.750 1.1014 0.00848 0.00319 -0.0880 0.2648 1.0000
6.000 1.1263 0.00881 0.00343 -0.0875 0.2439 1.0000
6.250 1.1507 0.00920 0.00371 -0.0869 0.2175 1.0000
6.500 1.1737 0.00975 0.00409 -0.0863 0.1805 1.0000
6.750 1.1950 0.01052 0.00458 -0.0853 0.1330 1.0000
7.000 1.2160 0.01132 0.00513 -0.0844 0.0914 1.0000
7.250 1.2391 0.01184 0.00557 -0.0837 0.0752 1.0000
7.500 1.2628 0.01225 0.00597 -0.0831 0.0656 1.0000
7.750 1.2862 0.01270 0.00638 -0.0824 0.0571 1.0000
8.000 1.3103 0.01304 0.00674 -0.0819 0.0522 1.0000
8.250 1.3334 0.01348 0.00716 -0.0812 0.0455 1.0000
8.500 1.3569 0.01387 0.00756 -0.0806 0.0404 1.0000
8.750 1.3795 0.01433 0.00800 -0.0799 0.0334 1.0000
9.000 1.3998 0.01504 0.00861 -0.0788 0.0213 1.0000
9.250 1.4191 0.01583 0.00936 -0.0776 0.0145 1.0000
9.500 1.4395 0.01646 0.01004 -0.0765 0.0125 1.0000
9.750 1.4575 0.01734 0.01099 -0.0751 0.0105 1.0000
10.000 1.4777 0.01791 0.01163 -0.0740 0.0099 1.0000
10.250 1.4966 0.01857 0.01237 -0.0728 0.0092 1.0000
10.500 1.5141 0.01932 0.01318 -0.0714 0.0086 1.0000
10.750 1.5277 0.02037 0.01432 -0.0694 0.0079 1.0000
11.000 1.5368 0.02169 0.01577 -0.0668 0.0074 1.0000
11.250 1.5518 0.02242 0.01659 -0.0651 0.0071 1.0000
11.500 1.5623 0.02325 0.01751 -0.0626 0.0069 1.0000
11.750 1.5705 0.02410 0.01844 -0.0599 0.0066 1.0000
12.000 1.5779 0.02502 0.01944 -0.0572 0.0063 1.0000
12.250 1.5833 0.02613 0.02063 -0.0546 0.0061 1.0000
12.500 1.5881 0.02735 0.02194 -0.0521 0.0058 1.0000
12.750 1.5901 0.02888 0.02355 -0.0498 0.0056 1.0000
13.000 1.5880 0.03087 0.02566 -0.0474 0.0055 1.0000
13.250 1.5805 0.03355 0.02847 -0.0453 0.0053 1.0000
13.500 1.5695 0.03688 0.03196 -0.0438 0.0051 1.0000
13.750 1.5533 0.04117 0.03642 -0.0429 0.0050 1.0000
14.000 1.5524 0.04404 0.03941 -0.0427 0.0050 1.0000
14.250 1.5522 0.04699 0.04247 -0.0428 0.0049 1.0000
14.500 1.5496 0.05040 0.04602 -0.0432 0.0048 1.0000
14.750 1.5425 0.05455 0.05030 -0.0440 0.0048 1.0000
15.000 1.5381 0.05851 0.05438 -0.0450 0.0047 1.0000
15.250 1.5270 0.06360 0.05961 -0.0465 0.0047 1.0000
15.500 1.5176 0.06865 0.06479 -0.0483 0.0046 1.0000
15.750 1.5099 0.07369 0.06995 -0.0502 0.0046 1.0000
16.000 1.4975 0.07965 0.07604 -0.0526 0.0045 1.0000
16.250 1.4863 0.08561 0.08214 -0.0552 0.0045 1.0000
16.500 1.4740 0.09196 0.08862 -0.0581 0.0045 1.0000
16.750 1.4625 0.09829 0.09507 -0.0611 0.0044 1.0000
17.000 1.4496 0.10505 0.10196 -0.0644 0.0044 1.0000
17.250 1.4369 0.11197 0.10900 -0.0678 0.0044 1.0000
17.500 1.4241 0.11903 0.11618 -0.0715 0.0043 1.0000
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Polar data table (+)
Polar graphs
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