GOE 432 AIRFOIL (goe432-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 432 AIRFOIL (goe432-il) Reynolds number: 100,000 Max Cl/Cd: 53.34 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe432-il-100000.txt Download as CSV file: xf-goe432-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 432 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3628 0.09900 0.09419 -0.0238 1.0000 0.0700
-7.750 -0.3777 0.09835 0.09367 -0.0251 1.0000 0.0707
-7.500 -0.3820 0.09659 0.09197 -0.0293 1.0000 0.0711
-7.250 -0.3760 0.09103 0.08648 -0.0250 1.0000 0.0725
-7.000 -0.3651 0.08763 0.08309 -0.0207 1.0000 0.0758
-6.750 -0.3639 0.08508 0.08059 -0.0204 1.0000 0.0786
-6.500 -0.3636 0.08255 0.07811 -0.0216 1.0000 0.0815
-6.250 -0.3598 0.08062 0.07615 -0.0316 1.0000 0.0850
-6.000 -0.3553 0.07631 0.07186 -0.0328 1.0000 0.0861
-5.750 -0.3525 0.07297 0.06862 -0.0274 1.0000 0.0879
-5.500 -0.3465 0.07031 0.06600 -0.0255 1.0000 0.0909
-5.250 -0.3367 0.06741 0.06308 -0.0271 1.0000 0.0960
-5.000 -0.3114 0.05256 0.04864 -0.0258 1.0000 0.1020
-4.750 -0.3071 0.04991 0.04604 -0.0230 1.0000 0.1054
-4.250 -0.2856 0.05445 0.04994 -0.0329 1.0000 0.1190
-4.000 -0.2671 0.05149 0.04683 -0.0350 1.0000 0.1323
-3.750 -0.2491 0.04895 0.04416 -0.0362 1.0000 0.1465
-3.500 -0.2381 0.04668 0.04198 -0.0340 1.0000 0.1534
-3.250 -0.2203 0.04428 0.03950 -0.0344 1.0000 0.1685
-3.000 -0.1998 0.04203 0.03709 -0.0355 1.0000 0.1918
-2.750 -0.1856 0.04010 0.03520 -0.0341 1.0000 0.2103
-2.250 -0.0897 0.02809 0.02106 -0.0425 1.0000 0.1033
-2.000 -0.0552 0.02636 0.01858 -0.0431 0.9978 0.0995
-1.750 -0.0041 0.02426 0.01592 -0.0471 0.9889 0.1001
-1.500 0.0438 0.02235 0.01390 -0.0512 0.9796 0.1075
-1.250 0.0962 0.02133 0.01248 -0.0554 0.9697 0.1155
-1.000 0.1437 0.02010 0.01128 -0.0589 0.9579 0.1239
-0.750 0.1901 0.01916 0.01039 -0.0622 0.9449 0.1376
-0.500 0.2365 0.01823 0.00956 -0.0653 0.9312 0.1543
-0.250 0.2830 0.01722 0.00888 -0.0683 0.9170 0.2075
0.000 0.3534 0.01462 0.00812 -0.0756 0.9096 1.0000
0.250 0.4020 0.01417 0.00746 -0.0787 0.8909 1.0000
0.500 0.4498 0.01365 0.00677 -0.0814 0.8716 1.0000
0.750 0.4913 0.01319 0.00616 -0.0828 0.8456 1.0000
1.000 0.5318 0.01284 0.00564 -0.0841 0.8164 1.0000
1.250 0.5685 0.01270 0.00529 -0.0847 0.7841 1.0000
1.500 0.6004 0.01276 0.00512 -0.0847 0.7516 1.0000
1.750 0.6277 0.01295 0.00510 -0.0839 0.7194 1.0000
2.000 0.6541 0.01321 0.00515 -0.0830 0.6910 1.0000
2.250 0.6793 0.01350 0.00527 -0.0821 0.6649 1.0000
2.500 0.7043 0.01382 0.00543 -0.0812 0.6416 1.0000
2.750 0.7292 0.01415 0.00560 -0.0802 0.6198 1.0000
3.000 0.7535 0.01450 0.00581 -0.0792 0.5984 1.0000
3.250 0.7779 0.01487 0.00606 -0.0783 0.5789 1.0000
3.500 0.8027 0.01527 0.00635 -0.0774 0.5615 1.0000
3.750 0.8275 0.01569 0.00670 -0.0766 0.5450 1.0000
4.000 0.8521 0.01613 0.00706 -0.0758 0.5289 1.0000
4.250 0.8763 0.01657 0.00745 -0.0749 0.5124 1.0000
4.500 0.8998 0.01699 0.00787 -0.0739 0.4951 1.0000
4.750 0.9229 0.01740 0.00831 -0.0729 0.4767 1.0000
5.000 0.9457 0.01779 0.00868 -0.0717 0.4569 1.0000
5.250 0.9678 0.01817 0.00902 -0.0704 0.4352 1.0000
5.500 0.9890 0.01854 0.00935 -0.0689 0.4107 1.0000
5.750 1.0095 0.01897 0.00976 -0.0673 0.3840 1.0000
6.000 1.0300 0.01947 0.01017 -0.0659 0.3586 1.0000
6.250 1.0513 0.02004 0.01063 -0.0646 0.3370 1.0000
6.500 1.0739 0.02073 0.01123 -0.0636 0.3202 1.0000
6.750 1.0967 0.02144 0.01195 -0.0627 0.3057 1.0000
7.000 1.1193 0.02213 0.01270 -0.0617 0.2925 1.0000
7.250 1.1417 0.02283 0.01350 -0.0607 0.2804 1.0000
7.500 1.1641 0.02358 0.01432 -0.0598 0.2694 1.0000
7.750 1.1873 0.02439 0.01517 -0.0590 0.2598 1.0000
8.000 1.2100 0.02520 0.01608 -0.0581 0.2499 1.0000
8.250 1.2316 0.02616 0.01721 -0.0571 0.2403 1.0000
8.500 1.2540 0.02713 0.01823 -0.0563 0.2307 1.0000
8.750 1.2762 0.02801 0.01908 -0.0554 0.2196 1.0000
9.000 1.2926 0.02860 0.01973 -0.0538 0.2053 1.0000
9.250 1.3064 0.02918 0.02040 -0.0518 0.1901 1.0000
9.500 1.3218 0.03001 0.02127 -0.0502 0.1765 1.0000
9.750 1.3374 0.03101 0.02229 -0.0487 0.1633 1.0000
10.000 1.3473 0.03176 0.02327 -0.0462 0.1495 1.0000
10.250 1.3512 0.03233 0.02411 -0.0429 0.1336 1.0000
10.500 1.3483 0.03265 0.02476 -0.0385 0.1120 1.0000
10.750 1.3426 0.03380 0.02585 -0.0341 0.0912 1.0000
11.000 1.3375 0.03593 0.02789 -0.0301 0.0777 1.0000
11.250 1.3354 0.03828 0.03022 -0.0270 0.0695 1.0000
11.500 1.3345 0.04039 0.03238 -0.0244 0.0630 1.0000
11.750 1.3349 0.04286 0.03497 -0.0221 0.0583 1.0000
12.000 1.3372 0.04525 0.03752 -0.0201 0.0548 1.0000
12.250 1.3402 0.04774 0.04004 -0.0185 0.0520 1.0000
12.500 1.3444 0.05089 0.04331 -0.0171 0.0498 1.0000
12.750 1.3424 0.05393 0.04666 -0.0157 0.0481 1.0000
13.000 1.3371 0.05719 0.05017 -0.0147 0.0465 1.0000
13.250 1.3316 0.06054 0.05373 -0.0143 0.0449 1.0000
13.500 1.3270 0.06416 0.05751 -0.0140 0.0438 1.0000
13.750 1.3222 0.06807 0.06157 -0.0141 0.0428 1.0000
14.000 1.3151 0.07253 0.06620 -0.0145 0.0422 1.0000
14.250 1.3027 0.07770 0.07159 -0.0157 0.0420 1.0000
14.500 1.2857 0.08354 0.07769 -0.0179 0.0419 1.0000
14.750 1.2626 0.09046 0.08490 -0.0216 0.0422 1.0000
15.000 1.2247 0.10061 0.09543 -0.0284 0.0433 1.0000
15.250 1.1554 0.11939 0.11463 -0.0416 0.0473 1.0000
15.500 1.1184 0.13298 0.12828 -0.0501 0.0496 1.0000
15.750 1.0923 0.14488 0.14021 -0.0570 0.0507 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 432 AIRFOIL (goe432-il)