GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 101 AIRFOIL (goe101-il) Reynolds number: 100,000 Max Cl/Cd: 58.66 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe101-il-100000.txt Download as CSV file: xf-goe101-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 101 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3914 0.10754 0.10259 -0.0150 1.0000 0.0509
-8.500 -0.3927 0.10613 0.10126 -0.0179 1.0000 0.0514
-8.250 -0.3962 0.10475 0.09998 -0.0209 1.0000 0.0517
-8.000 -0.3947 0.10265 0.09794 -0.0251 1.0000 0.0519
-7.750 -0.3822 0.09516 0.09047 -0.0202 1.0000 0.0534
-7.500 -0.3712 0.09112 0.08644 -0.0187 1.0000 0.0555
-7.250 -0.3675 0.08822 0.08360 -0.0190 1.0000 0.0575
-7.000 -0.3627 0.08536 0.08079 -0.0205 1.0000 0.0593
-6.750 -0.3571 0.08243 0.07791 -0.0231 1.0000 0.0617
-6.500 -0.3473 0.08004 0.07553 -0.0305 1.0000 0.0645
-6.250 -0.3303 0.07821 0.07352 -0.0403 1.0000 0.0654
-6.000 -0.3310 0.07227 0.06783 -0.0338 1.0000 0.0672
-5.750 -0.3249 0.06931 0.06491 -0.0310 1.0000 0.0711
-5.500 -0.3115 0.06636 0.06193 -0.0341 1.0000 0.0764
-5.250 -0.2938 0.06262 0.05805 -0.0399 1.0000 0.0805
-5.000 -0.2890 0.05958 0.05513 -0.0367 1.0000 0.0836
-4.750 -0.2592 0.05844 0.05347 -0.0440 1.0000 0.0933
-4.500 -0.2585 0.05340 0.04875 -0.0405 1.0000 0.0954
-4.250 -0.2502 0.05091 0.04633 -0.0387 1.0000 0.1008
-4.000 -0.2328 0.04801 0.04324 -0.0406 1.0000 0.1100
-3.750 -0.2154 0.04607 0.04106 -0.0417 1.0000 0.1226
-3.500 -0.2104 0.04347 0.03865 -0.0392 1.0000 0.1293
-3.250 -0.1981 0.04153 0.03662 -0.0390 1.0000 0.1450
-3.000 -0.1572 0.03853 0.03338 -0.0445 0.9930 0.1807
-2.750 -0.1148 0.03544 0.03028 -0.0490 0.9834 0.2249
-2.500 -0.0786 0.03299 0.02792 -0.0516 0.9720 0.2960
-2.250 -0.0508 0.03058 0.02577 -0.0510 0.9599 0.3836
-2.000 -0.0185 0.02821 0.02352 -0.0508 0.9478 0.4669
-1.750 0.1122 0.02294 0.01516 -0.0690 0.9428 0.1170
-1.500 0.1620 0.02126 0.01291 -0.0718 0.9303 0.1055
-1.250 0.2074 0.01951 0.01092 -0.0744 0.9165 0.1002
-1.000 0.2504 0.01824 0.00943 -0.0762 0.9010 0.0979
-0.750 0.2888 0.01723 0.00836 -0.0774 0.8834 0.1009
-0.500 0.3205 0.01656 0.00762 -0.0773 0.8614 0.1058
-0.250 0.3511 0.01587 0.00687 -0.0768 0.8418 0.1078
0.000 0.3773 0.01534 0.00634 -0.0758 0.8192 0.1128
0.250 0.4049 0.01505 0.00589 -0.0749 0.7997 0.1221
0.500 0.4316 0.01469 0.00558 -0.0741 0.7796 0.1564
0.750 0.4688 0.01266 0.00529 -0.0752 0.7622 1.0000
1.000 0.4946 0.01292 0.00525 -0.0744 0.7455 1.0000
1.250 0.5204 0.01319 0.00531 -0.0736 0.7301 1.0000
1.500 0.5461 0.01349 0.00544 -0.0730 0.7159 1.0000
1.750 0.5719 0.01382 0.00565 -0.0725 0.7027 1.0000
2.000 0.5978 0.01415 0.00588 -0.0720 0.6904 1.0000
2.250 0.6237 0.01450 0.00615 -0.0715 0.6793 1.0000
2.500 0.6499 0.01484 0.00640 -0.0711 0.6695 1.0000
2.750 0.6757 0.01520 0.00676 -0.0707 0.6587 1.0000
3.000 0.7012 0.01556 0.00711 -0.0702 0.6476 1.0000
3.250 0.7265 0.01586 0.00738 -0.0695 0.6356 1.0000
3.500 0.7516 0.01612 0.00761 -0.0688 0.6225 1.0000
3.750 0.7769 0.01638 0.00788 -0.0680 0.6100 1.0000
4.000 0.8028 0.01667 0.00813 -0.0674 0.5995 1.0000
4.250 0.8278 0.01698 0.00852 -0.0667 0.5874 1.0000
4.500 0.8527 0.01730 0.00890 -0.0661 0.5752 1.0000
4.750 0.8778 0.01762 0.00933 -0.0654 0.5634 1.0000
5.000 0.9030 0.01794 0.00972 -0.0647 0.5516 1.0000
5.250 0.9280 0.01822 0.01008 -0.0640 0.5388 1.0000
5.500 0.9527 0.01847 0.01042 -0.0631 0.5245 1.0000
5.750 0.9772 0.01874 0.01079 -0.0622 0.5097 1.0000
6.000 1.0012 0.01893 0.01114 -0.0612 0.4925 1.0000
6.250 1.0240 0.01882 0.01105 -0.0596 0.4675 1.0000
6.500 1.0449 0.01859 0.01087 -0.0578 0.4354 1.0000
6.750 1.0659 0.01852 0.01089 -0.0562 0.4035 1.0000
7.000 1.0853 0.01850 0.01099 -0.0545 0.3577 1.0000
7.250 1.1004 0.01904 0.01120 -0.0524 0.2622 1.0000
7.750 1.1049 0.02496 0.01550 -0.0463 0.0729 1.0000
8.000 1.1162 0.02675 0.01730 -0.0441 0.0607 1.0000
8.250 1.1301 0.02826 0.01893 -0.0420 0.0544 1.0000
8.500 1.1410 0.03027 0.02088 -0.0399 0.0507 1.0000
8.750 1.1584 0.03224 0.02293 -0.0381 0.0486 1.0000
9.000 1.1789 0.03432 0.02519 -0.0368 0.0462 1.0000
9.250 1.1977 0.03632 0.02728 -0.0357 0.0430 1.0000
9.500 1.2193 0.03919 0.03019 -0.0350 0.0410 1.0000
9.750 1.2422 0.04302 0.03426 -0.0344 0.0405 1.0000
10.000 1.2591 0.04656 0.03816 -0.0332 0.0405 1.0000
10.250 1.2700 0.04989 0.04195 -0.0313 0.0410 1.0000
10.500 1.2738 0.05363 0.04627 -0.0289 0.0421 1.0000
10.750 1.2634 0.05807 0.05135 -0.0258 0.0436 1.0000
11.000 1.2491 0.06239 0.05611 -0.0231 0.0448 1.0000
11.250 1.2307 0.06632 0.06035 -0.0205 0.0456 1.0000
11.500 1.2115 0.07043 0.06470 -0.0192 0.0463 1.0000
11.750 1.1912 0.07500 0.06949 -0.0192 0.0469 1.0000
12.000 1.1700 0.08015 0.07484 -0.0205 0.0474 1.0000
12.250 1.1481 0.08596 0.08082 -0.0231 0.0478 1.0000
12.500 1.1253 0.09254 0.08755 -0.0267 0.0482 1.0000
12.750 1.1015 0.10006 0.09520 -0.0315 0.0485 1.0000
13.000 1.0791 0.10820 0.10344 -0.0367 0.0490 1.0000
13.250 0.9098 0.11566 0.11137 -0.0388 0.0482 1.0000
13.500 0.8818 0.12381 0.11957 -0.0439 0.0492 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 101 AIRFOIL (goe101-il)