GOE 677 (= M 6) AIRFOIL (goe677-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 677 (= M 6) AIRFOIL (goe677-il) Reynolds number: 100,000 Max Cl/Cd: 49.16 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe677-il-100000.txt Download as CSV file: xf-goe677-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 677 (= M 6) 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4985 0.09738 0.09245 -0.0124 1.0000 0.1373
-8.750 -0.5337 0.09209 0.08727 -0.0218 1.0000 0.1415
-8.500 -0.5887 0.08880 0.08384 -0.0267 1.0000 0.1426
-8.250 -0.5109 0.08470 0.07994 -0.0194 1.0000 0.1495
-8.000 -0.4069 0.07336 0.06888 -0.0250 1.0000 0.1645
-7.750 -0.4232 0.06884 0.06441 -0.0271 1.0000 0.1696
-7.500 -0.5005 0.06458 0.06006 -0.0287 1.0000 0.1740
-7.250 -0.5651 0.07018 0.06512 -0.0247 1.0000 0.1746
-7.000 -0.5384 0.06579 0.06097 -0.0232 1.0000 0.1779
-6.750 -0.5354 0.06343 0.05855 -0.0222 1.0000 0.1875
-6.500 -0.5316 0.06023 0.05539 -0.0205 1.0000 0.1948
-6.250 -0.5437 0.05840 0.05346 -0.0177 1.0000 0.2075
-6.000 -0.5414 0.05653 0.05176 -0.0139 1.0000 0.2117
-5.750 -0.5609 0.05598 0.05124 -0.0085 1.0000 0.2163
-5.500 -0.5494 0.05353 0.04873 -0.0085 0.9955 0.2300
-5.250 -0.5135 0.03829 0.03078 -0.0142 0.9841 0.1113
-5.000 -0.4788 0.03516 0.02728 -0.0165 0.9759 0.1108
-4.750 -0.4392 0.03224 0.02396 -0.0193 0.9699 0.1101
-4.500 -0.4041 0.02994 0.02127 -0.0210 0.9611 0.1097
-4.250 -0.3599 0.02805 0.01892 -0.0241 0.9558 0.1112
-4.000 -0.3230 0.02622 0.01692 -0.0263 0.9470 0.1151
-3.750 -0.2752 0.02467 0.01524 -0.0300 0.9411 0.1193
-3.500 -0.2362 0.02350 0.01388 -0.0319 0.9305 0.1241
-3.250 -0.1927 0.02218 0.01255 -0.0347 0.9231 0.1325
-3.000 -0.1588 0.02140 0.01166 -0.0355 0.9127 0.1409
-2.750 -0.1253 0.02040 0.01080 -0.0366 0.9045 0.1519
-2.500 -0.0948 0.01957 0.01007 -0.0369 0.8956 0.1672
-2.250 -0.0689 0.01888 0.00951 -0.0364 0.8861 0.1878
-2.000 -0.0426 0.01797 0.00897 -0.0358 0.8780 0.2351
-1.750 0.1581 0.01673 0.01002 -0.0630 0.8855 1.0000
-1.500 0.1807 0.01672 0.00983 -0.0621 0.8735 1.0000
-1.250 0.2026 0.01672 0.00968 -0.0611 0.8619 1.0000
-1.000 0.2253 0.01677 0.00960 -0.0604 0.8500 1.0000
-0.750 0.2475 0.01681 0.00952 -0.0595 0.8389 1.0000
-0.500 0.2691 0.01683 0.00941 -0.0581 0.8288 1.0000
-0.250 0.2921 0.01691 0.00941 -0.0574 0.8169 1.0000
0.000 0.3148 0.01700 0.00941 -0.0565 0.8060 1.0000
0.250 0.3368 0.01703 0.00933 -0.0551 0.7963 1.0000
0.500 0.3599 0.01714 0.00938 -0.0544 0.7840 1.0000
0.750 0.3828 0.01726 0.00944 -0.0535 0.7724 1.0000
1.000 0.4055 0.01734 0.00944 -0.0523 0.7621 1.0000
1.250 0.4286 0.01745 0.00949 -0.0513 0.7503 1.0000
1.500 0.4517 0.01762 0.00964 -0.0504 0.7377 1.0000
1.750 0.4748 0.01779 0.00976 -0.0494 0.7261 1.0000
2.000 0.4978 0.01790 0.00977 -0.0481 0.7160 1.0000
2.250 0.5208 0.01814 0.01002 -0.0474 0.7028 1.0000
2.500 0.5438 0.01839 0.01027 -0.0465 0.6905 1.0000
2.750 0.5670 0.01861 0.01045 -0.0455 0.6799 1.0000
3.000 0.5902 0.01880 0.01062 -0.0445 0.6690 1.0000
3.250 0.6129 0.01912 0.01098 -0.0437 0.6570 1.0000
3.500 0.6359 0.01937 0.01122 -0.0427 0.6466 1.0000
3.750 0.6592 0.01951 0.01134 -0.0416 0.6364 1.0000
4.000 0.6812 0.01971 0.01158 -0.0404 0.6233 1.0000
4.250 0.7033 0.01977 0.01164 -0.0391 0.6102 1.0000
4.500 0.7259 0.01973 0.01158 -0.0376 0.5974 1.0000
4.750 0.7495 0.01960 0.01140 -0.0363 0.5861 1.0000
5.000 0.7714 0.01969 0.01155 -0.0350 0.5730 1.0000
5.250 0.7933 0.01980 0.01173 -0.0338 0.5605 1.0000
5.500 0.8157 0.01983 0.01181 -0.0326 0.5485 1.0000
5.750 0.8388 0.01976 0.01175 -0.0313 0.5369 1.0000
6.000 0.8618 0.01965 0.01163 -0.0300 0.5243 1.0000
6.250 0.8835 0.01959 0.01161 -0.0286 0.5095 1.0000
6.500 0.9052 0.01958 0.01166 -0.0271 0.4946 1.0000
6.750 0.9267 0.01962 0.01172 -0.0257 0.4794 1.0000
7.000 0.9480 0.01971 0.01181 -0.0242 0.4633 1.0000
7.250 0.9679 0.01993 0.01210 -0.0226 0.4454 1.0000
7.500 0.9869 0.02019 0.01241 -0.0209 0.4253 1.0000
7.750 1.0053 0.02045 0.01266 -0.0190 0.4028 1.0000
8.000 1.0206 0.02078 0.01304 -0.0167 0.3743 1.0000
8.250 1.0331 0.02114 0.01336 -0.0140 0.3381 1.0000
8.500 1.0410 0.02180 0.01386 -0.0107 0.2904 1.0000
8.750 1.0470 0.02293 0.01466 -0.0072 0.2508 1.0000
9.000 1.0555 0.02416 0.01566 -0.0044 0.2249 1.0000
9.250 1.0658 0.02532 0.01664 -0.0019 0.2067 1.0000
9.500 1.0767 0.02640 0.01762 0.0004 0.1913 1.0000
9.750 1.0867 0.02743 0.01862 0.0027 0.1771 1.0000
10.000 1.0959 0.02843 0.01963 0.0050 0.1641 1.0000
10.250 1.1055 0.02949 0.02072 0.0073 0.1530 1.0000
10.500 1.1151 0.03063 0.02184 0.0094 0.1431 1.0000
10.750 1.1212 0.03175 0.02297 0.0119 0.1330 1.0000
11.000 1.1225 0.03296 0.02432 0.0150 0.1225 1.0000
11.250 1.1223 0.03453 0.02591 0.0177 0.1106 1.0000
11.500 1.1218 0.03657 0.02790 0.0201 0.0973 1.0000
11.750 1.1251 0.03896 0.03023 0.0221 0.0847 1.0000
12.000 1.1336 0.04130 0.03255 0.0236 0.0755 1.0000
12.250 1.1414 0.04322 0.03445 0.0250 0.0694 1.0000
12.500 1.1537 0.04563 0.03700 0.0262 0.0648 1.0000
12.750 1.1605 0.04791 0.03952 0.0275 0.0615 1.0000
13.000 1.1699 0.04995 0.04161 0.0284 0.0587 1.0000
13.250 1.1814 0.05262 0.04432 0.0293 0.0562 1.0000
13.500 1.1754 0.05554 0.04759 0.0305 0.0548 1.0000
13.750 1.1695 0.05870 0.05103 0.0313 0.0535 1.0000
14.000 1.1634 0.06203 0.05461 0.0318 0.0523 1.0000
14.250 1.1559 0.06566 0.05846 0.0320 0.0515 1.0000
14.500 1.1453 0.06976 0.06280 0.0318 0.0510 1.0000
14.750 1.1303 0.07444 0.06772 0.0311 0.0507 1.0000
15.000 1.1077 0.08015 0.07370 0.0295 0.0508 1.0000
15.250 1.0759 0.08739 0.08125 0.0266 0.0513 1.0000
15.500 1.0370 0.09640 0.09056 0.0220 0.0522 1.0000
15.750 0.9931 0.10759 0.10199 0.0154 0.0535 1.0000
16.000 0.9503 0.12027 0.11483 0.0077 0.0551 1.0000
16.250 0.9183 0.13209 0.12670 0.0011 0.0563 1.0000
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Polar data table (+)
Polar graphs
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