GOE 676 (= M 12) AIRFOIL (goe676-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 676 (= M 12) AIRFOIL (goe676-il) Reynolds number: 100,000 Max Cl/Cd: 48.84 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe676-il-100000.txt Download as CSV file: xf-goe676-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 676 (= M 12) 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4950 0.10907 0.10392 -0.0141 1.0000 0.1296
-9.750 -0.4895 0.10574 0.10060 -0.0148 1.0000 0.1343
-9.500 -0.5257 0.10130 0.09630 -0.0243 1.0000 0.1398
-9.250 -0.5170 0.09630 0.09132 -0.0236 1.0000 0.1421
-9.000 -0.4909 0.09405 0.08902 -0.0197 1.0000 0.1476
-8.750 -0.5102 0.08953 0.08460 -0.0253 1.0000 0.1541
-8.500 -0.5759 0.08468 0.07971 -0.0338 1.0000 0.1560
-8.250 -0.5056 0.08177 0.07691 -0.0261 1.0000 0.1632
-8.000 -0.5318 0.07745 0.07262 -0.0298 1.0000 0.1696
-7.750 -0.5476 0.07279 0.06795 -0.0310 1.0000 0.1745
-7.500 -0.5304 0.07028 0.06549 -0.0289 1.0000 0.1800
-7.250 -0.5633 0.06665 0.06172 -0.0291 1.0000 0.1896
-7.000 -0.5452 0.06381 0.05900 -0.0268 1.0000 0.1937
-6.750 -0.5685 0.06142 0.05644 -0.0245 1.0000 0.2056
-6.500 -0.5545 0.05875 0.05390 -0.0218 1.0000 0.2100
-6.250 -0.5871 0.04657 0.04010 -0.0205 1.0000 0.1240
-6.000 -0.5845 0.04121 0.03408 -0.0173 1.0000 0.1094
-5.750 -0.5736 0.03825 0.03099 -0.0153 1.0000 0.1074
-5.500 -0.5620 0.03552 0.02794 -0.0132 1.0000 0.1051
-5.250 -0.5480 0.03319 0.02524 -0.0113 1.0000 0.1042
-5.000 -0.5323 0.03144 0.02317 -0.0096 1.0000 0.1056
-4.750 -0.5147 0.02979 0.02119 -0.0082 1.0000 0.1067
-4.500 -0.4956 0.02827 0.01934 -0.0070 1.0000 0.1074
-4.250 -0.4751 0.02702 0.01777 -0.0059 1.0000 0.1087
-4.000 -0.4517 0.02586 0.01638 -0.0056 0.9994 0.1117
-3.750 -0.4069 0.02470 0.01521 -0.0094 0.9932 0.1177
-3.500 -0.3626 0.02370 0.01399 -0.0127 0.9860 0.1235
-3.250 -0.3192 0.02269 0.01305 -0.0161 0.9788 0.1325
-3.000 -0.2744 0.02182 0.01220 -0.0197 0.9716 0.1447
-2.750 -0.2344 0.02100 0.01149 -0.0223 0.9627 0.1610
-2.500 -0.1892 0.02006 0.01075 -0.0260 0.9561 0.1902
-2.250 -0.1565 0.01876 0.01006 -0.0274 0.9458 0.2622
-2.000 -0.1196 0.01715 0.01115 -0.0247 0.9446 0.9058
-1.750 0.0272 0.01790 0.01147 -0.0451 0.9570 0.9841
-1.500 0.1205 0.01753 0.01084 -0.0581 0.9560 1.0000
-1.250 0.1671 0.01725 0.01042 -0.0623 0.9430 1.0000
-1.000 0.2087 0.01700 0.01004 -0.0653 0.9293 1.0000
-0.750 0.2445 0.01678 0.00973 -0.0671 0.9147 1.0000
-0.500 0.2749 0.01661 0.00947 -0.0676 0.8998 1.0000
-0.250 0.3014 0.01648 0.00926 -0.0672 0.8849 1.0000
0.000 0.3255 0.01639 0.00909 -0.0663 0.8702 1.0000
0.250 0.3484 0.01631 0.00895 -0.0652 0.8559 1.0000
0.500 0.3707 0.01625 0.00882 -0.0638 0.8420 1.0000
0.750 0.3928 0.01619 0.00869 -0.0623 0.8286 1.0000
1.000 0.4145 0.01618 0.00862 -0.0608 0.8145 1.0000
1.250 0.4356 0.01621 0.00861 -0.0593 0.7998 1.0000
1.500 0.4570 0.01626 0.00862 -0.0579 0.7852 1.0000
1.750 0.4785 0.01631 0.00863 -0.0564 0.7706 1.0000
2.000 0.5002 0.01635 0.00864 -0.0550 0.7561 1.0000
2.250 0.5222 0.01639 0.00865 -0.0536 0.7416 1.0000
2.500 0.5443 0.01644 0.00866 -0.0522 0.7270 1.0000
2.750 0.5667 0.01648 0.00868 -0.0508 0.7123 1.0000
3.000 0.5891 0.01653 0.00868 -0.0494 0.6971 1.0000
3.250 0.6114 0.01658 0.00869 -0.0480 0.6811 1.0000
3.500 0.6337 0.01664 0.00870 -0.0465 0.6644 1.0000
3.750 0.6561 0.01674 0.00876 -0.0451 0.6476 1.0000
4.000 0.6785 0.01688 0.00886 -0.0438 0.6309 1.0000
4.250 0.7009 0.01706 0.00900 -0.0425 0.6145 1.0000
4.500 0.7232 0.01728 0.00921 -0.0412 0.5985 1.0000
4.750 0.7456 0.01754 0.00947 -0.0400 0.5836 1.0000
5.000 0.7678 0.01779 0.00971 -0.0388 0.5684 1.0000
5.250 0.7898 0.01800 0.00995 -0.0376 0.5530 1.0000
5.500 0.8117 0.01820 0.01015 -0.0362 0.5373 1.0000
5.750 0.8334 0.01837 0.01034 -0.0349 0.5213 1.0000
6.000 0.8553 0.01860 0.01059 -0.0336 0.5066 1.0000
6.250 0.8773 0.01885 0.01088 -0.0324 0.4923 1.0000
6.500 0.8992 0.01911 0.01117 -0.0312 0.4777 1.0000
6.750 0.9205 0.01937 0.01148 -0.0298 0.4618 1.0000
7.000 0.9410 0.01964 0.01180 -0.0283 0.4440 1.0000
7.250 0.9610 0.01986 0.01202 -0.0267 0.4232 1.0000
7.500 0.9785 0.02010 0.01226 -0.0247 0.3975 1.0000
7.750 0.9938 0.02035 0.01247 -0.0224 0.3663 1.0000
8.000 1.0076 0.02074 0.01282 -0.0199 0.3316 1.0000
8.250 1.0203 0.02135 0.01329 -0.0175 0.2959 1.0000
8.500 1.0321 0.02210 0.01398 -0.0151 0.2575 1.0000
8.750 1.0430 0.02302 0.01473 -0.0126 0.2243 1.0000
9.000 1.0526 0.02411 0.01565 -0.0102 0.1950 1.0000
9.250 1.0594 0.02548 0.01684 -0.0075 0.1674 1.0000
9.500 1.0626 0.02726 0.01841 -0.0044 0.1415 1.0000
9.750 1.0663 0.02945 0.02027 -0.0015 0.1207 1.0000
10.000 1.0761 0.03137 0.02213 0.0007 0.1051 1.0000
10.250 1.0894 0.03312 0.02375 0.0023 0.0940 1.0000
10.500 1.1013 0.03457 0.02541 0.0043 0.0863 1.0000
10.750 1.1183 0.03630 0.02707 0.0053 0.0798 1.0000
11.000 1.1296 0.03778 0.02877 0.0073 0.0747 1.0000
11.250 1.1516 0.03967 0.03053 0.0076 0.0703 1.0000
11.500 1.1666 0.04189 0.03303 0.0089 0.0672 1.0000
11.750 1.1763 0.04380 0.03516 0.0107 0.0641 1.0000
12.000 1.1875 0.04563 0.03707 0.0121 0.0615 1.0000
12.250 1.2103 0.04910 0.04057 0.0117 0.0593 1.0000
12.500 1.2075 0.05176 0.04361 0.0145 0.0586 1.0000
12.750 1.2007 0.05469 0.04689 0.0170 0.0580 1.0000
13.000 1.1904 0.05789 0.05042 0.0192 0.0574 1.0000
13.250 1.1768 0.06141 0.05426 0.0209 0.0570 1.0000
13.500 1.1603 0.06534 0.05848 0.0219 0.0567 1.0000
13.750 1.1408 0.06978 0.06321 0.0222 0.0565 1.0000
14.000 1.1184 0.07487 0.06856 0.0216 0.0566 1.0000
14.250 1.0932 0.08072 0.07465 0.0199 0.0569 1.0000
14.500 1.0658 0.08739 0.08154 0.0173 0.0573 1.0000
14.750 1.0374 0.09487 0.08921 0.0137 0.0579 1.0000
15.000 1.0093 0.10303 0.09752 0.0094 0.0585 1.0000
15.250 0.9837 0.11164 0.10623 0.0048 0.0591 1.0000
15.500 0.8646 0.14809 0.14281 -0.0181 0.0696 1.0000
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Polar data table (+)
Polar graphs
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