GOE 673 AIRFOIL (goe673-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 673 AIRFOIL (goe673-il) Reynolds number: 1,000,000 Max Cl/Cd: 118.4 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe673-il-1000000.txt Download as CSV file: xf-goe673-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 673 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3388 0.09516 0.09365 -0.0414 1.0000 0.0096
-10.500 -0.4612 0.09802 0.09642 -0.0350 1.0000 0.0089
-10.250 -0.4617 0.09368 0.09210 -0.0371 1.0000 0.0090
-10.000 -0.4616 0.08944 0.08788 -0.0392 1.0000 0.0091
-9.750 -0.4627 0.08495 0.08340 -0.0416 1.0000 0.0092
-9.500 -0.4658 0.08007 0.07855 -0.0443 1.0000 0.0092
-9.250 -0.4725 0.07481 0.07332 -0.0471 1.0000 0.0093
-9.000 -0.5206 0.05580 0.05428 -0.0618 0.9980 0.0089
-8.750 -0.5175 0.04680 0.04496 -0.0740 0.9920 0.0090
-8.500 -0.5039 0.04196 0.03991 -0.0789 0.9879 0.0092
-8.250 -0.4838 0.03872 0.03650 -0.0823 0.9852 0.0094
-8.000 -0.4635 0.03586 0.03345 -0.0847 0.9817 0.0098
-7.250 -0.4205 0.02092 0.01703 -0.0838 0.9613 0.0079
-7.000 -0.4001 0.01805 0.01377 -0.0830 0.9561 0.0079
-6.750 -0.3765 0.01678 0.01227 -0.0823 0.9500 0.0083
-6.500 -0.3522 0.01572 0.01099 -0.0817 0.9446 0.0086
-6.250 -0.3305 0.01339 0.00843 -0.0810 0.9391 0.0091
-6.000 -0.3057 0.01250 0.00746 -0.0806 0.9332 0.0095
-5.750 -0.2802 0.01185 0.00674 -0.0803 0.9282 0.0099
-5.500 -0.2539 0.01131 0.00614 -0.0801 0.9222 0.0104
-5.250 -0.2274 0.01089 0.00566 -0.0799 0.9165 0.0112
-5.000 -0.2007 0.01039 0.00508 -0.0798 0.9108 0.0120
-4.750 -0.1735 0.01003 0.00466 -0.0797 0.9045 0.0126
-4.500 -0.1474 0.00922 0.00370 -0.0795 0.8987 0.0137
-4.250 -0.1200 0.00877 0.00319 -0.0795 0.8919 0.0151
-4.000 -0.0924 0.00852 0.00286 -0.0795 0.8858 0.0168
-3.750 -0.0643 0.00832 0.00264 -0.0796 0.8786 0.0186
-3.250 -0.0082 0.00772 0.00193 -0.0797 0.8648 0.0341
-3.000 0.0183 0.00694 0.00162 -0.0799 0.8580 0.1706
-2.750 0.0460 0.00652 0.00148 -0.0803 0.8501 0.2552
-2.500 0.0739 0.00628 0.00138 -0.0805 0.8428 0.3124
-2.250 0.1019 0.00601 0.00130 -0.0808 0.8344 0.3774
-2.000 0.1298 0.00577 0.00125 -0.0810 0.8262 0.4489
-1.750 0.1579 0.00561 0.00119 -0.0812 0.8173 0.5012
-1.500 0.1862 0.00549 0.00115 -0.0814 0.8077 0.5372
-1.250 0.2143 0.00536 0.00112 -0.0815 0.7979 0.5813
-1.000 0.2423 0.00526 0.00110 -0.0817 0.7876 0.6308
-0.750 0.2704 0.00520 0.00110 -0.0818 0.7760 0.6698
-0.500 0.2986 0.00517 0.00110 -0.0819 0.7638 0.6983
-0.250 0.3266 0.00516 0.00110 -0.0820 0.7508 0.7227
0.000 0.3545 0.00515 0.00110 -0.0820 0.7368 0.7476
0.250 0.3821 0.00514 0.00112 -0.0820 0.7219 0.7737
0.500 0.4094 0.00514 0.00114 -0.0819 0.7063 0.8015
0.750 0.4357 0.00511 0.00118 -0.0816 0.6901 0.8398
1.000 0.4598 0.00507 0.00124 -0.0807 0.6730 0.8897
1.250 0.4820 0.00508 0.00129 -0.0792 0.6562 0.9431
1.500 0.5114 0.00513 0.00131 -0.0795 0.6403 0.9806
1.750 0.5467 0.00521 0.00134 -0.0813 0.6255 1.0000
2.000 0.5742 0.00532 0.00140 -0.0814 0.6121 1.0000
2.250 0.6019 0.00543 0.00146 -0.0815 0.5993 1.0000
2.500 0.6296 0.00555 0.00155 -0.0816 0.5859 1.0000
2.750 0.6573 0.00567 0.00163 -0.0817 0.5729 1.0000
3.000 0.6848 0.00581 0.00173 -0.0818 0.5576 1.0000
3.250 0.7116 0.00601 0.00183 -0.0817 0.5328 1.0000
3.500 0.7379 0.00625 0.00197 -0.0816 0.4997 1.0000
3.750 0.7635 0.00657 0.00212 -0.0814 0.4604 1.0000
4.000 0.7886 0.00694 0.00231 -0.0811 0.4164 1.0000
4.250 0.8128 0.00742 0.00255 -0.0807 0.3621 1.0000
4.500 0.8351 0.00809 0.00288 -0.0800 0.2887 1.0000
4.750 0.8490 0.00970 0.00365 -0.0781 0.1242 1.0000
5.000 0.8663 0.01098 0.00444 -0.0765 0.0220 1.0000
5.250 0.8909 0.01143 0.00490 -0.0760 0.0162 1.0000
5.500 0.9163 0.01176 0.00526 -0.0757 0.0147 1.0000
5.750 0.9410 0.01216 0.00570 -0.0753 0.0132 1.0000
6.000 0.9639 0.01275 0.00635 -0.0745 0.0117 1.0000
6.250 0.9850 0.01351 0.00721 -0.0734 0.0107 1.0000
6.500 1.0093 0.01387 0.00760 -0.0730 0.0100 1.0000
6.750 1.0318 0.01441 0.00819 -0.0722 0.0093 1.0000
7.000 1.0534 0.01501 0.00885 -0.0712 0.0088 1.0000
7.250 1.0743 0.01564 0.00952 -0.0702 0.0083 1.0000
7.500 1.0935 0.01640 0.01032 -0.0689 0.0078 1.0000
7.750 1.1058 0.01777 0.01179 -0.0665 0.0074 1.0000
8.000 1.1156 0.01957 0.01371 -0.0636 0.0070 1.0000
8.250 1.1357 0.02021 0.01442 -0.0624 0.0068 1.0000
8.500 1.1532 0.02123 0.01552 -0.0609 0.0066 1.0000
8.750 1.1707 0.02229 0.01667 -0.0593 0.0063 1.0000
9.000 1.1879 0.02333 0.01781 -0.0579 0.0060 1.0000
9.250 1.2045 0.02441 0.01899 -0.0563 0.0057 1.0000
9.500 1.2201 0.02592 0.02062 -0.0546 0.0056 1.0000
9.750 1.2351 0.02756 0.02239 -0.0528 0.0055 1.0000
10.000 1.2485 0.02922 0.02420 -0.0509 0.0053 1.0000
10.250 1.2598 0.03094 0.02607 -0.0487 0.0052 1.0000
10.500 1.2685 0.03288 0.02818 -0.0463 0.0052 1.0000
10.750 1.2744 0.03451 0.02995 -0.0436 0.0050 1.0000
11.000 1.2781 0.03628 0.03186 -0.0409 0.0050 1.0000
11.250 1.2799 0.03800 0.03371 -0.0382 0.0049 1.0000
11.500 1.2784 0.04026 0.03614 -0.0355 0.0048 1.0000
11.750 1.2718 0.04321 0.03931 -0.0326 0.0047 1.0000
12.000 1.2614 0.04666 0.04299 -0.0298 0.0047 1.0000
12.250 1.2427 0.05109 0.04767 -0.0273 0.0046 1.0000
12.500 1.2229 0.05579 0.05261 -0.0254 0.0046 1.0000
12.750 1.2061 0.06030 0.05731 -0.0246 0.0046 1.0000
13.000 1.1882 0.06528 0.06249 -0.0245 0.0046 1.0000
13.250 1.1669 0.07107 0.06847 -0.0254 0.0046 1.0000
13.500 1.1507 0.07662 0.07419 -0.0269 0.0046 1.0000
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Polar data table (+)
Polar graphs
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