GOE 429 AIRFOIL (goe429-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 429 AIRFOIL (goe429-il) Reynolds number: 100,000 Max Cl/Cd: 37.57 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe429-il-100000-n5.txt Download as CSV file: xf-goe429-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 429 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
------ -------- --------- --------- -------- -------- --------
-14.250 -1.0197 0.09564 0.09016 0.0180 1.0000 0.0322
-14.000 -1.0454 0.08653 0.08085 0.0123 1.0000 0.0321
-13.750 -1.0675 0.07852 0.07262 0.0074 1.0000 0.0322
-13.500 -1.0816 0.07234 0.06631 0.0037 1.0000 0.0325
-13.250 -1.0911 0.06717 0.06104 0.0006 1.0000 0.0330
-13.000 -1.0995 0.06235 0.05610 -0.0022 1.0000 0.0336
-12.750 -1.1067 0.05790 0.05152 -0.0047 1.0000 0.0343
-12.500 -1.1125 0.05381 0.04727 -0.0068 1.0000 0.0353
-12.250 -1.1170 0.05004 0.04329 -0.0084 1.0000 0.0364
-12.000 -1.1200 0.04655 0.03953 -0.0094 1.0000 0.0378
-11.750 -1.1205 0.04365 0.03646 -0.0096 1.0000 0.0391
-11.500 -1.1181 0.04158 0.03433 -0.0091 1.0000 0.0405
-11.250 -1.1108 0.03961 0.03222 -0.0087 1.0000 0.0427
-11.000 -1.1009 0.03752 0.02985 -0.0081 1.0000 0.0462
-10.750 -1.0884 0.03606 0.02836 -0.0077 1.0000 0.0490
-10.500 -1.0735 0.03450 0.02659 -0.0072 1.0000 0.0534
-10.250 -1.0573 0.03309 0.02509 -0.0068 1.0000 0.0575
-10.000 -1.0388 0.03191 0.02375 -0.0064 1.0000 0.0627
-9.750 -1.0196 0.03071 0.02242 -0.0061 1.0000 0.0676
-9.500 -0.9989 0.02973 0.02129 -0.0058 1.0000 0.0733
-9.250 -0.9778 0.02866 0.02007 -0.0055 1.0000 0.0789
-9.000 -0.9557 0.02781 0.01913 -0.0053 1.0000 0.0854
-8.750 -0.9333 0.02681 0.01800 -0.0051 1.0000 0.0918
-8.500 -0.9103 0.02588 0.01699 -0.0049 1.0000 0.0979
-8.250 -0.8861 0.02496 0.01588 -0.0046 1.0000 0.1035
-8.000 -0.8632 0.02387 0.01485 -0.0043 1.0000 0.1088
-7.750 -0.8387 0.02300 0.01389 -0.0041 1.0000 0.1147
-7.500 -0.8145 0.02208 0.01295 -0.0039 1.0000 0.1207
-7.250 -0.7898 0.02126 0.01210 -0.0037 1.0000 0.1276
-7.000 -0.7647 0.02046 0.01131 -0.0035 1.0000 0.1352
-6.750 -0.7392 0.01974 0.01059 -0.0034 1.0000 0.1452
-6.500 -0.7136 0.01904 0.00994 -0.0034 1.0000 0.1574
-6.250 -0.6873 0.01841 0.00935 -0.0034 1.0000 0.1742
-6.000 -0.6610 0.01779 0.00881 -0.0034 1.0000 0.1956
-5.750 -0.6343 0.01724 0.00827 -0.0035 1.0000 0.2196
-5.500 -0.6076 0.01669 0.00781 -0.0035 1.0000 0.2429
-5.250 -0.5805 0.01621 0.00739 -0.0036 1.0000 0.2671
-5.000 -0.5533 0.01577 0.00703 -0.0037 1.0000 0.2903
-4.750 -0.5259 0.01539 0.00672 -0.0038 1.0000 0.3139
-4.500 -0.4983 0.01504 0.00645 -0.0039 1.0000 0.3377
-4.250 -0.4705 0.01474 0.00620 -0.0040 1.0000 0.3617
-4.000 -0.4427 0.01445 0.00600 -0.0041 1.0000 0.3851
-3.750 -0.4149 0.01419 0.00582 -0.0042 1.0000 0.4113
-3.500 -0.3874 0.01394 0.00570 -0.0043 1.0000 0.4370
-3.250 -0.3602 0.01374 0.00560 -0.0044 1.0000 0.4634
-3.000 -0.3305 0.01358 0.00555 -0.0050 0.9877 0.4895
-2.750 -0.2924 0.01347 0.00552 -0.0071 0.9604 0.5144
-2.500 -0.2564 0.01341 0.00551 -0.0085 0.9363 0.5363
-2.250 -0.2243 0.01340 0.00550 -0.0090 0.9133 0.5560
-2.000 -0.1963 0.01341 0.00551 -0.0085 0.8910 0.5743
-1.750 -0.1701 0.01343 0.00553 -0.0076 0.8708 0.5913
-1.500 -0.1444 0.01346 0.00554 -0.0066 0.8521 0.6080
-1.250 -0.1191 0.01349 0.00555 -0.0055 0.8339 0.6250
-1.000 -0.0936 0.01350 0.00554 -0.0045 0.8148 0.6416
-0.750 -0.0681 0.01351 0.00552 -0.0035 0.7950 0.6578
-0.500 -0.0426 0.01351 0.00549 -0.0024 0.7754 0.6737
-0.250 -0.0171 0.01351 0.00547 -0.0013 0.7566 0.6901
0.000 0.0082 0.01351 0.00547 -0.0002 0.7384 0.7083
0.250 0.0334 0.01351 0.00546 0.0010 0.7202 0.7274
0.500 0.0589 0.01348 0.00545 0.0020 0.7009 0.7456
0.750 0.0849 0.01345 0.00543 0.0029 0.6795 0.7610
1.000 0.1109 0.01341 0.00539 0.0038 0.6569 0.7753
1.250 0.1372 0.01336 0.00535 0.0046 0.6315 0.7900
1.500 0.1632 0.01333 0.00530 0.0055 0.6030 0.8052
1.750 0.1892 0.01330 0.00526 0.0064 0.5697 0.8213
2.000 0.2150 0.01330 0.00522 0.0073 0.5305 0.8390
2.250 0.2407 0.01336 0.00518 0.0082 0.4872 0.8590
2.500 0.2675 0.01351 0.00519 0.0088 0.4435 0.8826
2.750 0.2978 0.01373 0.00528 0.0086 0.4030 0.9126
3.000 0.3333 0.01402 0.00544 0.0070 0.3661 0.9520
3.250 0.3687 0.01435 0.00563 0.0053 0.3350 1.0000
3.500 0.3965 0.01473 0.00587 0.0049 0.3120 1.0000
3.750 0.4243 0.01511 0.00615 0.0047 0.2930 1.0000
4.000 0.4521 0.01551 0.00645 0.0045 0.2762 1.0000
4.250 0.4797 0.01591 0.00677 0.0043 0.2591 1.0000
4.500 0.5072 0.01630 0.00707 0.0041 0.2415 1.0000
4.750 0.5347 0.01669 0.00739 0.0040 0.2259 1.0000
5.000 0.5622 0.01709 0.00776 0.0039 0.2147 1.0000
5.250 0.5894 0.01750 0.00813 0.0038 0.2036 1.0000
5.500 0.6170 0.01785 0.00853 0.0037 0.1915 1.0000
5.750 0.6444 0.01823 0.00897 0.0037 0.1818 1.0000
6.000 0.6715 0.01864 0.00939 0.0036 0.1711 1.0000
6.250 0.6986 0.01903 0.00981 0.0036 0.1558 1.0000
6.500 0.7256 0.01946 0.01029 0.0036 0.1379 1.0000
6.750 0.7519 0.02002 0.01081 0.0035 0.1149 1.0000
7.000 0.7778 0.02070 0.01146 0.0036 0.0912 1.0000
7.250 0.8026 0.02157 0.01221 0.0037 0.0770 1.0000
7.500 0.8270 0.02248 0.01309 0.0038 0.0711 1.0000
7.750 0.8505 0.02351 0.01412 0.0040 0.0672 1.0000
8.000 0.8739 0.02450 0.01518 0.0042 0.0644 1.0000
8.250 0.8966 0.02555 0.01633 0.0045 0.0616 1.0000
8.500 0.9182 0.02671 0.01756 0.0048 0.0594 1.0000
8.750 0.9386 0.02799 0.01891 0.0052 0.0578 1.0000
9.000 0.9581 0.02941 0.02034 0.0057 0.0565 1.0000
9.250 0.9789 0.03065 0.02175 0.0063 0.0555 1.0000
9.500 0.9989 0.03200 0.02325 0.0068 0.0543 1.0000
9.750 1.0180 0.03345 0.02486 0.0073 0.0529 1.0000
10.000 1.0360 0.03499 0.02655 0.0079 0.0514 1.0000
10.250 1.0532 0.03664 0.02832 0.0084 0.0502 1.0000
10.500 1.0696 0.03843 0.03022 0.0090 0.0492 1.0000
10.750 1.0850 0.04038 0.03224 0.0096 0.0482 1.0000
11.000 1.0995 0.04285 0.03475 0.0102 0.0470 1.0000
11.250 1.1083 0.04507 0.03729 0.0108 0.0461 1.0000
11.500 1.1130 0.04749 0.04003 0.0114 0.0453 1.0000
11.750 1.1127 0.05017 0.04300 0.0121 0.0447 1.0000
12.000 1.1072 0.05324 0.04634 0.0124 0.0442 1.0000
12.250 1.0984 0.05693 0.05028 0.0116 0.0437 1.0000
12.500 1.0860 0.06136 0.05497 0.0096 0.0434 1.0000
12.750 1.0696 0.06677 0.06063 0.0065 0.0431 1.0000
13.000 1.0472 0.07366 0.06777 0.0018 0.0431 1.0000
13.250 1.0145 0.08323 0.07762 -0.0055 0.0433 1.0000
13.500 0.9572 0.09937 0.09403 -0.0177 0.0443 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 429 AIRFOIL (goe429-il)