GOE 374 AIRFOIL (goe374-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 374 AIRFOIL (goe374-il) Reynolds number: 100,000 Max Cl/Cd: 57.08 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe374-il-100000-n5.txt Download as CSV file: xf-goe374-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 374 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.000 -0.3113 0.09029 0.08588 -0.0220 1.0000 0.0227
-6.750 -0.3076 0.08730 0.08293 -0.0218 1.0000 0.0231
-6.500 -0.3039 0.08461 0.08031 -0.0221 1.0000 0.0234
-6.250 -0.3005 0.08195 0.07771 -0.0223 1.0000 0.0239
-6.000 -0.2969 0.07938 0.07519 -0.0226 1.0000 0.0247
-5.750 -0.2927 0.07689 0.07274 -0.0231 1.0000 0.0252
-5.500 -0.2878 0.07436 0.07026 -0.0236 1.0000 0.0258
-5.250 -0.2815 0.07184 0.06776 -0.0243 1.0000 0.0267
-5.000 -0.2737 0.06928 0.06520 -0.0250 1.0000 0.0272
-4.750 -0.2300 0.06478 0.06056 -0.0342 0.9925 0.0289
-4.500 -0.1767 0.06067 0.05617 -0.0445 0.9835 0.0297
-4.250 -0.1292 0.05672 0.05192 -0.0518 0.9744 0.0299
-4.000 -0.0886 0.05229 0.04730 -0.0571 0.9664 0.0302
-3.750 -0.0673 0.04724 0.04232 -0.0593 0.9570 0.0313
-3.500 -0.0346 0.04372 0.03867 -0.0626 0.9485 0.0336
-3.250 0.0159 0.04124 0.03565 -0.0674 0.9408 0.0393
-3.000 0.0395 0.03740 0.03187 -0.0692 0.9306 0.0431
-2.750 0.0792 0.03499 0.02901 -0.0717 0.9209 0.0515
-2.500 0.1147 0.03282 0.02651 -0.0738 0.9114 0.0652
-2.250 0.1418 0.03003 0.02364 -0.0750 0.8987 0.0810
-1.750 0.2138 0.02549 0.01813 -0.0750 0.8736 0.0484
-1.500 0.2454 0.02313 0.01544 -0.0752 0.8596 0.0445
-1.250 0.2776 0.02181 0.01369 -0.0751 0.8443 0.0478
-1.000 0.3086 0.02025 0.01178 -0.0749 0.8281 0.0466
-0.750 0.3385 0.01930 0.01046 -0.0745 0.8099 0.0488
-0.500 0.3670 0.01830 0.00918 -0.0739 0.7897 0.0508
-0.250 0.3955 0.01740 0.00808 -0.0734 0.7687 0.0509
0.000 0.4224 0.01661 0.00715 -0.0727 0.7453 0.0514
0.250 0.4488 0.01597 0.00643 -0.0720 0.7212 0.0544
0.500 0.4751 0.01558 0.00591 -0.0713 0.6970 0.0588
0.750 0.5011 0.01526 0.00541 -0.0705 0.6733 0.0587
1.000 0.5273 0.01510 0.00501 -0.0697 0.6513 0.0588
1.250 0.5537 0.01505 0.00474 -0.0692 0.6297 0.0591
1.500 0.5800 0.01508 0.00454 -0.0686 0.6096 0.0597
1.750 0.6061 0.01515 0.00443 -0.0681 0.5899 0.0608
2.000 0.6320 0.01525 0.00437 -0.0675 0.5708 0.0625
2.250 0.6579 0.01536 0.00437 -0.0670 0.5525 0.0664
2.500 0.6837 0.01546 0.00442 -0.0665 0.5355 0.0840
3.000 0.7474 0.01418 0.00468 -0.0684 0.5016 1.0000
3.250 0.7723 0.01446 0.00482 -0.0678 0.4873 1.0000
3.500 0.7970 0.01475 0.00501 -0.0672 0.4731 1.0000
3.750 0.8214 0.01506 0.00521 -0.0665 0.4585 1.0000
4.000 0.8457 0.01539 0.00544 -0.0658 0.4433 1.0000
4.250 0.8697 0.01572 0.00573 -0.0651 0.4275 1.0000
4.500 0.8935 0.01606 0.00601 -0.0644 0.4118 1.0000
4.750 0.9175 0.01641 0.00633 -0.0638 0.3979 1.0000
5.000 0.9416 0.01677 0.00669 -0.0631 0.3858 1.0000
5.250 0.9655 0.01713 0.00710 -0.0625 0.3739 1.0000
5.500 0.9894 0.01748 0.00750 -0.0618 0.3611 1.0000
5.750 1.0131 0.01783 0.00793 -0.0612 0.3479 1.0000
6.000 1.0366 0.01819 0.00842 -0.0605 0.3348 1.0000
6.250 1.0599 0.01857 0.00890 -0.0597 0.3217 1.0000
6.500 1.0829 0.01898 0.00941 -0.0590 0.3083 1.0000
6.750 1.1056 0.01941 0.00995 -0.0582 0.2944 1.0000
7.000 1.1268 0.01986 0.01045 -0.0572 0.2750 1.0000
7.250 1.1459 0.02040 0.01097 -0.0560 0.2432 1.0000
7.500 1.1650 0.02107 0.01158 -0.0549 0.2129 1.0000
7.750 1.1836 0.02184 0.01229 -0.0538 0.1822 1.0000
8.000 1.1991 0.02295 0.01314 -0.0524 0.1404 1.0000
8.250 1.2143 0.02418 0.01419 -0.0509 0.1081 1.0000
8.750 1.2297 0.02818 0.01766 -0.0463 0.0226 1.0000
9.000 1.2404 0.02974 0.01941 -0.0442 0.0197 1.0000
9.250 1.2497 0.03135 0.02126 -0.0419 0.0175 1.0000
9.500 1.2582 0.03288 0.02308 -0.0397 0.0164 1.0000
9.750 1.2629 0.03452 0.02499 -0.0370 0.0159 1.0000
10.000 1.2634 0.03633 0.02703 -0.0341 0.0154 1.0000
10.250 1.2619 0.03832 0.02927 -0.0314 0.0150 1.0000
10.500 1.2579 0.04065 0.03182 -0.0291 0.0147 1.0000
10.750 1.2522 0.04332 0.03470 -0.0274 0.0144 1.0000
11.000 1.2464 0.04625 0.03783 -0.0264 0.0141 1.0000
11.250 1.2390 0.04966 0.04144 -0.0260 0.0137 1.0000
11.500 1.2306 0.05352 0.04548 -0.0262 0.0133 1.0000
11.750 1.2226 0.05763 0.04977 -0.0270 0.0129 1.0000
12.000 1.2151 0.06192 0.05422 -0.0280 0.0127 1.0000
12.250 1.2088 0.06619 0.05864 -0.0291 0.0126 1.0000
12.500 1.2007 0.07081 0.06340 -0.0304 0.0121 1.0000
12.750 1.1941 0.07529 0.06807 -0.0316 0.0118 1.0000
13.000 1.1896 0.07946 0.07236 -0.0324 0.0116 1.0000
13.250 1.1871 0.08340 0.07645 -0.0331 0.0115 1.0000
13.500 1.1846 0.08739 0.08057 -0.0337 0.0114 1.0000
13.750 1.1821 0.09156 0.08489 -0.0346 0.0113 1.0000
14.000 1.1785 0.09609 0.08959 -0.0359 0.0112 1.0000
14.250 1.1735 0.10107 0.09474 -0.0377 0.0112 1.0000
14.500 1.1671 0.10642 0.10027 -0.0400 0.0112 1.0000
14.750 1.1597 0.11218 0.10621 -0.0427 0.0112 1.0000
15.000 1.1511 0.11839 0.11259 -0.0459 0.0112 1.0000
15.250 1.1424 0.12487 0.11926 -0.0495 0.0112 1.0000
15.500 1.1333 0.13170 0.12626 -0.0535 0.0113 1.0000
15.750 1.1236 0.13896 0.13368 -0.0578 0.0114 1.0000
16.000 1.1135 0.14670 0.14157 -0.0625 0.0114 1.0000
16.250 1.1038 0.15478 0.14980 -0.0675 0.0116 1.0000
16.500 1.0936 0.16342 0.15857 -0.0728 0.0117 1.0000
16.750 1.0835 0.17259 0.16785 -0.0783 0.0120 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 374 AIRFOIL (goe374-il)