GOE 647 AIRFOIL (goe647-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 647 AIRFOIL (goe647-il) Reynolds number: 200,000 Max Cl/Cd: 58.99 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe647-il-200000.txt Download as CSV file: xf-goe647-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 647 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1637 0.09050 0.08723 -0.0639 0.9730 0.0915
-9.000 -0.1395 0.08685 0.08357 -0.0683 0.9689 0.0944
-8.750 -0.2732 0.04853 0.04470 -0.1106 0.9397 0.0707
-8.500 -0.3132 0.03562 0.03025 -0.1140 0.9231 0.0619
-8.250 -0.2857 0.03384 0.02855 -0.1149 0.9149 0.0631
-8.000 -0.2606 0.03132 0.02573 -0.1159 0.9072 0.0635
-7.750 -0.2402 0.02902 0.02308 -0.1156 0.8975 0.0637
-7.500 -0.2141 0.02703 0.02079 -0.1158 0.8890 0.0641
-7.250 -0.1920 0.02555 0.01908 -0.1150 0.8781 0.0646
-7.000 -0.1634 0.02413 0.01741 -0.1152 0.8695 0.0656
-6.750 -0.1417 0.02311 0.01618 -0.1139 0.8570 0.0669
-6.500 -0.1157 0.02209 0.01488 -0.1132 0.8458 0.0683
-6.250 -0.0901 0.02119 0.01370 -0.1124 0.8334 0.0693
-6.000 -0.0669 0.02002 0.01247 -0.1113 0.8200 0.0705
-5.750 -0.0409 0.01924 0.01166 -0.1107 0.8088 0.0722
-5.500 -0.0154 0.01862 0.01095 -0.1099 0.7974 0.0742
-5.250 0.0088 0.01813 0.01035 -0.1089 0.7856 0.0768
-5.000 0.0355 0.01744 0.00956 -0.1083 0.7757 0.0796
-4.750 0.0582 0.01693 0.00909 -0.1071 0.7633 0.0828
-4.500 0.0832 0.01651 0.00858 -0.1062 0.7524 0.0869
-4.250 0.1074 0.01592 0.00798 -0.1052 0.7417 0.0921
-4.000 0.1310 0.01559 0.00759 -0.1041 0.7299 0.0998
-3.750 0.1549 0.01504 0.00702 -0.1031 0.7198 0.1125
-3.500 0.1758 0.01436 0.00651 -0.1017 0.7083 0.1366
-3.250 0.1989 0.01392 0.00632 -0.1007 0.6976 0.1940
-3.000 0.2246 0.01380 0.00621 -0.1000 0.6871 0.2427
-2.750 0.2486 0.01375 0.00618 -0.0990 0.6751 0.2715
-2.500 0.2745 0.01375 0.00608 -0.0983 0.6643 0.2947
-2.250 0.2994 0.01372 0.00601 -0.0974 0.6529 0.3147
-2.000 0.3239 0.01370 0.00598 -0.0965 0.6413 0.3343
-1.750 0.3494 0.01369 0.00593 -0.0958 0.6306 0.3554
-1.500 0.3729 0.01367 0.00592 -0.0947 0.6181 0.3758
-1.250 0.3973 0.01365 0.00589 -0.0937 0.6062 0.3951
-1.000 0.4225 0.01365 0.00582 -0.0929 0.5950 0.4153
-0.750 0.4455 0.01360 0.00582 -0.0917 0.5821 0.4379
-0.500 0.4690 0.01357 0.00579 -0.0907 0.5695 0.4630
-0.250 0.4925 0.01351 0.00573 -0.0896 0.5573 0.4910
0.000 0.5144 0.01340 0.00569 -0.0882 0.5438 0.5257
0.250 0.5351 0.01323 0.00569 -0.0866 0.5304 0.5710
0.500 0.5542 0.01300 0.00572 -0.0846 0.5180 0.6477
0.750 0.6595 0.01263 0.00589 -0.0997 0.4947 0.9814
1.000 0.6940 0.01289 0.00594 -0.1012 0.4800 1.0000
1.250 0.7142 0.01313 0.00598 -0.0996 0.4682 1.0000
1.500 0.7345 0.01336 0.00606 -0.0981 0.4564 1.0000
1.750 0.7552 0.01360 0.00619 -0.0967 0.4458 1.0000
2.000 0.7763 0.01390 0.00629 -0.0953 0.4365 1.0000
2.250 0.7972 0.01414 0.00647 -0.0939 0.4269 1.0000
2.500 0.8188 0.01447 0.00660 -0.0927 0.4188 1.0000
2.750 0.8398 0.01471 0.00681 -0.0914 0.4105 1.0000
3.000 0.8616 0.01502 0.00698 -0.0902 0.4032 1.0000
3.250 0.8835 0.01533 0.00722 -0.0891 0.3964 1.0000
3.500 0.9052 0.01561 0.00744 -0.0879 0.3899 1.0000
3.750 0.9290 0.01601 0.00767 -0.0872 0.3844 1.0000
4.000 0.9507 0.01631 0.00797 -0.0860 0.3788 1.0000
4.250 0.9728 0.01661 0.00824 -0.0850 0.3735 1.0000
4.500 0.9965 0.01699 0.00850 -0.0843 0.3687 1.0000
4.750 1.0199 0.01739 0.00885 -0.0835 0.3639 1.0000
5.000 1.0410 0.01769 0.00917 -0.0823 0.3591 1.0000
5.250 1.0635 0.01803 0.00946 -0.0814 0.3546 1.0000
5.500 1.0890 0.01848 0.00979 -0.0811 0.3505 1.0000
5.750 1.1117 0.01891 0.01023 -0.0803 0.3468 1.0000
6.000 1.1327 0.01927 0.01063 -0.0792 0.3430 1.0000
6.250 1.1551 0.01965 0.01101 -0.0783 0.3393 1.0000
6.500 1.1791 0.02006 0.01137 -0.0778 0.3360 1.0000
6.750 1.2080 0.02067 0.01186 -0.0783 0.3327 1.0000
7.000 1.2273 0.02109 0.01238 -0.0769 0.3297 1.0000
7.250 1.2467 0.02151 0.01286 -0.0756 0.3265 1.0000
7.500 1.2677 0.02192 0.01330 -0.0746 0.3232 1.0000
7.750 1.2907 0.02236 0.01371 -0.0739 0.3201 1.0000
8.000 1.3169 0.02289 0.01419 -0.0739 0.3175 1.0000
8.250 1.3446 0.02363 0.01489 -0.0743 0.3148 1.0000
8.500 1.3604 0.02411 0.01552 -0.0725 0.3125 1.0000
8.750 1.3781 0.02464 0.01615 -0.0710 0.3100 1.0000
9.000 1.3973 0.02519 0.01676 -0.0699 0.3075 1.0000
9.250 1.4182 0.02572 0.01732 -0.0690 0.3050 1.0000
9.500 1.4415 0.02624 0.01784 -0.0686 0.3026 1.0000
9.750 1.4696 0.02687 0.01843 -0.0691 0.3003 1.0000
10.000 1.4932 0.02776 0.01935 -0.0689 0.2980 1.0000
10.250 1.5050 0.02840 0.02015 -0.0666 0.2961 1.0000
10.500 1.5171 0.02911 0.02100 -0.0645 0.2941 1.0000
10.750 1.5300 0.02983 0.02183 -0.0626 0.2920 1.0000
11.000 1.5436 0.03050 0.02259 -0.0607 0.2898 1.0000
11.250 1.5604 0.03110 0.02324 -0.0594 0.2877 1.0000
11.500 1.5829 0.03170 0.02386 -0.0590 0.2856 1.0000
11.750 1.6131 0.03251 0.02464 -0.0600 0.2837 1.0000
12.000 1.6342 0.03366 0.02587 -0.0597 0.2817 1.0000
12.250 1.6297 0.03454 0.02695 -0.0552 0.2801 1.0000
12.500 1.6267 0.03554 0.02812 -0.0513 0.2781 1.0000
12.750 1.6269 0.03660 0.02933 -0.0481 0.2760 1.0000
13.000 1.6304 0.03765 0.03050 -0.0456 0.2739 1.0000
13.250 1.6392 0.03844 0.03137 -0.0437 0.2716 1.0000
13.500 1.6649 0.03865 0.03153 -0.0437 0.2689 1.0000
13.750 1.6972 0.03920 0.03201 -0.0448 0.2657 1.0000
14.000 1.6726 0.04089 0.03396 -0.0395 0.2636 1.0000
14.250 1.6567 0.04282 0.03608 -0.0358 0.2612 1.0000
14.500 1.6462 0.04465 0.03808 -0.0330 0.2586 1.0000
14.750 1.6517 0.04553 0.03900 -0.0316 0.2556 1.0000
15.000 1.6989 0.04416 0.03745 -0.0329 0.2520 1.0000
15.250 1.6990 0.04580 0.03919 -0.0312 0.2494 1.0000
15.500 1.6642 0.04947 0.04315 -0.0279 0.2473 1.0000
15.750 1.6289 0.05388 0.04780 -0.0256 0.2449 1.0000
16.000 1.6023 0.05807 0.05217 -0.0243 0.2422 1.0000
16.250 1.6144 0.05886 0.05300 -0.0239 0.2395 1.0000
16.500 1.7074 0.05294 0.04673 -0.0251 0.2353 1.0000
16.750 1.6499 0.05968 0.05382 -0.0232 0.2336 1.0000
17.750 1.6153 0.07190 0.06645 -0.0228 0.2216 1.0000
18.000 0.9943 0.18127 0.17633 -0.0676 0.1706 1.0000
18.250 1.0126 0.18134 0.17644 -0.0674 0.1689 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 647 AIRFOIL (goe647-il)