GOE 515 AIRFOIL (goe515-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 515 AIRFOIL (goe515-il) Reynolds number: 200,000 Max Cl/Cd: 77.22 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe515-il-200000.txt Download as CSV file: xf-goe515-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 515 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4382 0.10443 0.10067 -0.0302 1.0000 0.0595
-9.750 -0.4240 0.10234 0.09857 -0.0284 1.0000 0.0623
-9.500 -0.4216 0.09932 0.09557 -0.0293 1.0000 0.0648
-9.250 -0.4278 0.09597 0.09228 -0.0321 1.0000 0.0679
-9.000 -0.4502 0.09248 0.08890 -0.0384 1.0000 0.0692
-8.750 -0.4651 0.08905 0.08557 -0.0409 1.0000 0.0693
-8.500 -0.4768 0.08486 0.08141 -0.0440 1.0000 0.0695
-8.250 -0.4633 0.08106 0.07764 -0.0375 1.0000 0.0709
-8.000 -0.4556 0.07904 0.07564 -0.0349 1.0000 0.0722
-7.750 -0.4575 0.07658 0.07322 -0.0340 1.0000 0.0734
-7.500 -0.4613 0.07418 0.07086 -0.0333 1.0000 0.0754
-7.250 -0.5281 0.05676 0.05321 -0.0428 1.0000 0.0587
-7.000 -0.5234 0.05554 0.05201 -0.0396 1.0000 0.0566
-6.750 -0.5292 0.05131 0.04767 -0.0381 1.0000 0.0557
-6.500 -0.5358 0.04507 0.04118 -0.0373 1.0000 0.0543
-6.250 -0.5441 0.03299 0.02814 -0.0376 0.9989 0.0522
-6.000 -0.5183 0.02757 0.02192 -0.0394 0.9950 0.0544
-5.750 -0.4890 0.02473 0.01836 -0.0405 0.9909 0.0574
-5.500 -0.4594 0.02219 0.01543 -0.0417 0.9870 0.0599
-5.250 -0.4230 0.02173 0.01492 -0.0438 0.9837 0.0639
-5.000 -0.3924 0.02050 0.01336 -0.0445 0.9790 0.0672
-4.750 -0.3597 0.01922 0.01175 -0.0456 0.9748 0.0704
-4.500 -0.3232 0.01873 0.01130 -0.0476 0.9715 0.0756
-4.250 -0.2901 0.01812 0.01045 -0.0486 0.9670 0.0809
-4.000 -0.2587 0.01737 0.00969 -0.0495 0.9619 0.0870
-3.750 -0.2220 0.01688 0.00909 -0.0513 0.9583 0.0977
-3.500 -0.1828 0.01645 0.00861 -0.0537 0.9557 0.1117
-3.250 -0.1566 0.01618 0.00831 -0.0534 0.9484 0.1232
-3.000 -0.1205 0.01578 0.00795 -0.0552 0.9445 0.1356
-2.750 -0.0810 0.01541 0.00760 -0.0577 0.9418 0.1473
-2.500 -0.0525 0.01518 0.00732 -0.0578 0.9352 0.1562
-2.250 -0.0165 0.01498 0.00713 -0.0595 0.9307 0.1680
-2.000 0.0237 0.01460 0.00682 -0.0621 0.9280 0.1797
-1.750 0.0656 0.01416 0.00644 -0.0649 0.9259 0.1899
-1.500 0.0897 0.01398 0.00628 -0.0640 0.9171 0.2001
-1.250 0.1286 0.01352 0.00591 -0.0661 0.9133 0.2158
-1.000 0.1597 0.01304 0.00560 -0.0664 0.9051 0.2392
-0.750 0.1946 0.01229 0.00517 -0.0674 0.8983 0.3013
-0.500 0.3307 0.00995 0.00490 -0.0895 0.9075 1.0000
-0.250 0.3591 0.00986 0.00471 -0.0892 0.8987 1.0000
0.000 0.3844 0.00982 0.00460 -0.0884 0.8889 1.0000
0.250 0.4138 0.00972 0.00442 -0.0883 0.8814 1.0000
0.500 0.4361 0.00974 0.00440 -0.0869 0.8702 1.0000
0.750 0.4606 0.00974 0.00435 -0.0859 0.8603 1.0000
1.000 0.4874 0.00970 0.00425 -0.0853 0.8517 1.0000
1.250 0.5099 0.00974 0.00427 -0.0839 0.8401 1.0000
1.500 0.5336 0.00978 0.00429 -0.0827 0.8292 1.0000
1.750 0.5577 0.00977 0.00424 -0.0815 0.8171 1.0000
2.000 0.5813 0.00976 0.00418 -0.0801 0.8029 1.0000
2.250 0.6044 0.00975 0.00415 -0.0786 0.7877 1.0000
2.500 0.6279 0.00977 0.00414 -0.0773 0.7733 1.0000
2.750 0.6513 0.00980 0.00416 -0.0760 0.7581 1.0000
3.000 0.6742 0.00983 0.00418 -0.0746 0.7414 1.0000
3.250 0.6972 0.00986 0.00421 -0.0732 0.7232 1.0000
3.500 0.7204 0.00990 0.00420 -0.0718 0.7030 1.0000
3.750 0.7428 0.00998 0.00427 -0.0703 0.6804 1.0000
4.000 0.7656 0.01010 0.00435 -0.0689 0.6581 1.0000
4.250 0.7873 0.01025 0.00445 -0.0673 0.6303 1.0000
4.500 0.8077 0.01046 0.00456 -0.0655 0.5952 1.0000
4.750 0.8272 0.01075 0.00470 -0.0635 0.5541 1.0000
5.000 0.8461 0.01108 0.00491 -0.0615 0.5108 1.0000
5.250 0.8638 0.01152 0.00516 -0.0593 0.4659 1.0000
5.500 0.8808 0.01204 0.00549 -0.0571 0.4213 1.0000
5.750 0.8969 0.01263 0.00588 -0.0549 0.3724 1.0000
6.000 0.9115 0.01334 0.00633 -0.0525 0.3138 1.0000
6.250 0.9243 0.01421 0.00683 -0.0499 0.2369 1.0000
6.500 0.9366 0.01523 0.00746 -0.0474 0.1783 1.0000
6.750 0.9506 0.01618 0.00820 -0.0451 0.1450 1.0000
7.000 0.9654 0.01708 0.00896 -0.0430 0.1139 1.0000
7.250 0.9800 0.01804 0.00978 -0.0408 0.0832 1.0000
7.500 0.9915 0.01924 0.01080 -0.0381 0.0640 1.0000
7.750 1.0047 0.02032 0.01187 -0.0356 0.0542 1.0000
8.000 1.0173 0.02155 0.01316 -0.0330 0.0485 1.0000
8.250 1.0310 0.02267 0.01428 -0.0308 0.0440 1.0000
8.500 1.0445 0.02419 0.01582 -0.0286 0.0406 1.0000
8.750 1.0620 0.02531 0.01706 -0.0269 0.0380 1.0000
9.000 1.0795 0.02658 0.01838 -0.0254 0.0359 1.0000
9.250 1.0989 0.02830 0.02009 -0.0243 0.0341 1.0000
9.500 1.1212 0.03086 0.02280 -0.0238 0.0324 1.0000
9.750 1.1370 0.03202 0.02419 -0.0219 0.0308 1.0000
10.000 1.1542 0.03391 0.02631 -0.0205 0.0297 1.0000
10.250 1.1694 0.03601 0.02865 -0.0188 0.0289 1.0000
10.500 1.1818 0.03840 0.03131 -0.0169 0.0284 1.0000
10.750 1.1903 0.04101 0.03420 -0.0146 0.0281 1.0000
11.000 1.1929 0.04389 0.03741 -0.0118 0.0280 1.0000
11.250 1.1901 0.04658 0.04038 -0.0083 0.0278 1.0000
11.500 1.1764 0.04992 0.04411 -0.0039 0.0281 1.0000
11.750 1.1836 0.05141 0.04557 -0.0024 0.0270 1.0000
12.000 1.1481 0.05632 0.05105 0.0023 0.0281 1.0000
12.250 1.1613 0.05737 0.05197 0.0029 0.0267 1.0000
12.500 1.1487 0.06069 0.05547 0.0046 0.0265 1.0000
12.750 1.1282 0.06475 0.05978 0.0058 0.0266 1.0000
13.000 1.0480 0.07667 0.07242 0.0033 0.0294 1.0000
13.250 1.0176 0.08437 0.08032 -0.0002 0.0305 1.0000
13.500 0.9904 0.09243 0.08852 -0.0049 0.0311 1.0000
13.750 0.9637 0.10158 0.09778 -0.0107 0.0317 1.0000
14.000 0.9447 0.11011 0.10633 -0.0149 0.0326 1.0000
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Polar data table (+)
Polar graphs
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