GOE 507 AIRFOIL (goe507-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 507 AIRFOIL (goe507-il) Reynolds number: 200,000 Max Cl/Cd: 67.39 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe507-il-200000-n5.txt Download as CSV file: xf-goe507-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 507 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6993 0.03626 0.03224 -0.0586 1.0000 0.0210
-9.000 -0.7078 0.03104 0.02645 -0.0568 0.9993 0.0218
-8.750 -0.6853 0.02720 0.02204 -0.0589 0.9940 0.0235
-8.500 -0.6534 0.02734 0.02220 -0.0602 0.9902 0.0251
-8.250 -0.6227 0.02637 0.02104 -0.0617 0.9865 0.0270
-8.000 -0.5973 0.02397 0.01812 -0.0627 0.9814 0.0301
-7.750 -0.5640 0.02434 0.01854 -0.0640 0.9781 0.0321
-7.500 -0.5319 0.02422 0.01832 -0.0653 0.9749 0.0349
-7.250 -0.5040 0.02285 0.01649 -0.0659 0.9701 0.0394
-7.000 -0.4724 0.02205 0.01550 -0.0672 0.9670 0.0422
-6.750 -0.4420 0.02187 0.01529 -0.0681 0.9629 0.0448
-6.500 -0.4120 0.02131 0.01450 -0.0689 0.9588 0.0485
-6.250 -0.3802 0.02068 0.01355 -0.0700 0.9558 0.0519
-6.000 -0.3523 0.01982 0.01245 -0.0703 0.9509 0.0543
-5.750 -0.3230 0.01902 0.01160 -0.0710 0.9467 0.0568
-5.500 -0.2913 0.01842 0.01089 -0.0720 0.9436 0.0594
-5.250 -0.2636 0.01781 0.01014 -0.0721 0.9386 0.0615
-5.000 -0.2342 0.01727 0.00942 -0.0725 0.9342 0.0634
-4.750 -0.2027 0.01672 0.00872 -0.0732 0.9309 0.0647
-4.500 -0.1750 0.01621 0.00810 -0.0732 0.9256 0.0653
-4.250 -0.1467 0.01558 0.00740 -0.0733 0.9206 0.0661
-4.000 -0.1164 0.01486 0.00665 -0.0738 0.9163 0.0676
-3.750 -0.0907 0.01440 0.00618 -0.0734 0.9086 0.0693
-3.500 -0.0616 0.01396 0.00571 -0.0735 0.9023 0.0707
-3.250 -0.0354 0.01359 0.00531 -0.0730 0.8943 0.0718
-3.000 -0.0078 0.01325 0.00492 -0.0728 0.8874 0.0731
-2.750 0.0190 0.01296 0.00460 -0.0724 0.8810 0.0747
-2.500 0.0454 0.01273 0.00432 -0.0720 0.8745 0.0762
-2.250 0.0736 0.01250 0.00404 -0.0719 0.8696 0.0778
-2.000 0.0988 0.01229 0.00383 -0.0712 0.8621 0.0800
-1.750 0.1260 0.01207 0.00360 -0.0709 0.8562 0.0831
-1.500 0.1522 0.01190 0.00344 -0.0705 0.8493 0.0870
-1.250 0.1788 0.01173 0.00329 -0.0701 0.8424 0.0930
-1.000 0.2050 0.01148 0.00318 -0.0696 0.8357 0.1168
-0.750 0.2306 0.01125 0.00317 -0.0691 0.8280 0.1764
-0.500 0.2572 0.01110 0.00310 -0.0687 0.8209 0.2115
-0.250 0.2829 0.01095 0.00306 -0.0682 0.8124 0.2412
0.000 0.3092 0.01079 0.00300 -0.0677 0.8044 0.2689
0.250 0.3340 0.01052 0.00296 -0.0671 0.7952 0.3357
0.750 0.4231 0.00886 0.00305 -0.0732 0.7769 0.9707
1.000 0.4717 0.00884 0.00295 -0.0775 0.7614 1.0000
1.250 0.4951 0.00888 0.00291 -0.0763 0.7445 1.0000
1.500 0.5189 0.00893 0.00290 -0.0752 0.7281 1.0000
1.750 0.5428 0.00900 0.00291 -0.0742 0.7120 1.0000
2.000 0.5662 0.00908 0.00292 -0.0731 0.6901 1.0000
2.250 0.5890 0.00919 0.00290 -0.0718 0.6598 1.0000
2.500 0.6109 0.00936 0.00290 -0.0703 0.6217 1.0000
2.750 0.6324 0.00959 0.00294 -0.0688 0.5832 1.0000
3.000 0.6541 0.00985 0.00304 -0.0674 0.5531 1.0000
3.250 0.6758 0.01013 0.00319 -0.0660 0.5267 1.0000
3.500 0.6974 0.01044 0.00337 -0.0647 0.4988 1.0000
3.750 0.7194 0.01073 0.00357 -0.0635 0.4734 1.0000
4.000 0.7414 0.01103 0.00379 -0.0623 0.4493 1.0000
4.250 0.7635 0.01133 0.00402 -0.0611 0.4224 1.0000
4.500 0.7846 0.01169 0.00426 -0.0598 0.3902 1.0000
4.750 0.8056 0.01208 0.00454 -0.0585 0.3581 1.0000
5.000 0.8260 0.01251 0.00485 -0.0572 0.3239 1.0000
5.250 0.8459 0.01301 0.00520 -0.0558 0.2865 1.0000
5.500 0.8648 0.01360 0.00561 -0.0543 0.2454 1.0000
5.750 0.8833 0.01427 0.00606 -0.0528 0.2004 1.0000
6.000 0.8998 0.01514 0.00661 -0.0511 0.1352 1.0000
6.250 0.9159 0.01609 0.00727 -0.0493 0.0948 1.0000
6.500 0.9348 0.01678 0.00790 -0.0479 0.0829 1.0000
6.750 0.9550 0.01733 0.00849 -0.0466 0.0759 1.0000
7.000 0.9739 0.01798 0.00915 -0.0452 0.0701 1.0000
7.250 0.9939 0.01850 0.00976 -0.0440 0.0649 1.0000
7.500 1.0110 0.01927 0.01053 -0.0424 0.0595 1.0000
7.750 1.0316 0.01972 0.01110 -0.0412 0.0546 1.0000
8.000 1.0493 0.02039 0.01181 -0.0397 0.0489 1.0000
8.250 1.0680 0.02096 0.01247 -0.0384 0.0440 1.0000
8.500 1.0865 0.02154 0.01314 -0.0369 0.0382 1.0000
8.750 1.1017 0.02236 0.01394 -0.0351 0.0333 1.0000
9.000 1.1174 0.02308 0.01477 -0.0333 0.0301 1.0000
9.250 1.1313 0.02384 0.01557 -0.0313 0.0274 1.0000
9.500 1.1413 0.02483 0.01657 -0.0288 0.0250 1.0000
9.750 1.1517 0.02584 0.01769 -0.0264 0.0232 1.0000
10.000 1.1623 0.02688 0.01883 -0.0242 0.0219 1.0000
10.250 1.1726 0.02794 0.01999 -0.0220 0.0207 1.0000
10.500 1.1816 0.02912 0.02123 -0.0199 0.0197 1.0000
10.750 1.1870 0.03067 0.02283 -0.0176 0.0187 1.0000
11.000 1.1967 0.03195 0.02427 -0.0158 0.0179 1.0000
11.250 1.2048 0.03339 0.02585 -0.0140 0.0169 1.0000
11.500 1.2118 0.03495 0.02755 -0.0122 0.0164 1.0000
11.750 1.2179 0.03659 0.02931 -0.0106 0.0158 1.0000
12.000 1.2224 0.03839 0.03120 -0.0090 0.0153 1.0000
12.250 1.2257 0.04037 0.03327 -0.0076 0.0149 1.0000
12.500 1.2281 0.04264 0.03567 -0.0063 0.0145 1.0000
12.750 1.2311 0.04485 0.03811 -0.0052 0.0140 1.0000
13.000 1.2314 0.04746 0.04093 -0.0041 0.0136 1.0000
13.250 1.2308 0.05015 0.04381 -0.0034 0.0133 1.0000
13.500 1.2291 0.05300 0.04683 -0.0030 0.0129 1.0000
13.750 1.2256 0.05621 0.05022 -0.0029 0.0127 1.0000
14.000 1.2212 0.05960 0.05376 -0.0032 0.0124 1.0000
14.250 1.2150 0.06344 0.05777 -0.0038 0.0122 1.0000
14.500 1.2072 0.06769 0.06219 -0.0048 0.0121 1.0000
14.750 1.1987 0.07223 0.06688 -0.0064 0.0120 1.0000
15.000 1.1888 0.07723 0.07205 -0.0083 0.0119 1.0000
15.250 1.1767 0.08295 0.07795 -0.0109 0.0118 1.0000
15.500 1.1626 0.08933 0.08452 -0.0141 0.0118 1.0000
15.750 1.1488 0.09591 0.09126 -0.0176 0.0117 1.0000
16.000 1.1292 0.10421 0.09978 -0.0223 0.0118 1.0000
16.250 1.1079 0.11334 0.10911 -0.0278 0.0118 1.0000
16.500 1.0445 0.13488 0.13114 -0.0413 0.0124 1.0000
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