GOE 436 AIRFOIL (goe436-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 436 AIRFOIL (goe436-il) Reynolds number: 100,000 Max Cl/Cd: 47.78 at α=10.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe436-il-100000.txt Download as CSV file: xf-goe436-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 436 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3460 0.09296 0.08805 -0.0325 1.0000 0.1192
-8.000 -0.3773 0.09220 0.08746 -0.0322 1.0000 0.1213
-7.750 -0.4140 0.09130 0.08670 -0.0328 1.0000 0.1221
-7.500 -0.4497 0.09013 0.08550 -0.0352 1.0000 0.1227
-7.250 -0.4146 0.08469 0.08021 -0.0273 1.0000 0.1252
-7.000 -0.4093 0.08277 0.07833 -0.0235 1.0000 0.1283
-6.750 -0.4177 0.08084 0.07645 -0.0218 1.0000 0.1319
-6.500 -0.4357 0.07857 0.07419 -0.0235 1.0000 0.1366
-6.250 -0.4529 0.07486 0.07038 -0.0269 1.0000 0.1399
-6.000 -0.4455 0.07245 0.06809 -0.0223 1.0000 0.1421
-5.750 -0.4427 0.07040 0.06606 -0.0200 1.0000 0.1455
-5.500 -0.4479 0.06710 0.06245 -0.0256 1.0000 0.1560
-5.250 -0.4417 0.06425 0.05975 -0.0223 1.0000 0.1579
-5.000 -0.4356 0.06235 0.05790 -0.0199 1.0000 0.1620
-4.750 -0.4067 0.05861 0.05400 -0.0252 0.9943 0.1762
-4.500 -0.3639 0.05488 0.05011 -0.0318 0.9843 0.1928
-4.250 -0.3256 0.05165 0.04676 -0.0365 0.9739 0.2101
-4.000 -0.2898 0.04920 0.04409 -0.0412 0.9637 0.2398
-3.750 -0.2546 0.04649 0.04145 -0.0433 0.9566 0.2602
-3.500 -0.2273 0.04448 0.03938 -0.0447 0.9460 0.2915
-3.250 -0.1514 0.03388 0.02651 -0.0565 0.9411 0.1416
-3.000 -0.1179 0.03168 0.02393 -0.0575 0.9307 0.1340
-2.750 -0.0696 0.02928 0.02107 -0.0610 0.9251 0.1297
-2.500 -0.0376 0.02799 0.01951 -0.0616 0.9136 0.1316
-2.250 0.0025 0.02665 0.01788 -0.0634 0.9046 0.1334
-2.000 0.0469 0.02533 0.01629 -0.0659 0.8965 0.1350
-1.750 0.0847 0.02454 0.01525 -0.0671 0.8860 0.1389
-1.500 0.1345 0.02296 0.01374 -0.0707 0.8800 0.1452
-1.250 0.1726 0.02206 0.01281 -0.0719 0.8692 0.1507
-1.000 0.2225 0.02079 0.01163 -0.0752 0.8631 0.1612
-0.750 0.2561 0.02006 0.01090 -0.0754 0.8504 0.1714
-0.500 0.2911 0.01911 0.01011 -0.0759 0.8385 0.1867
-0.250 0.4090 0.01515 0.00878 -0.0916 0.8349 1.0000
0.000 0.4411 0.01488 0.00824 -0.0914 0.8184 1.0000
0.250 0.4712 0.01466 0.00781 -0.0909 0.8006 1.0000
0.500 0.5004 0.01447 0.00743 -0.0902 0.7815 1.0000
0.750 0.5221 0.01444 0.00725 -0.0883 0.7573 1.0000
1.000 0.5483 0.01433 0.00694 -0.0872 0.7343 1.0000
1.250 0.5709 0.01435 0.00678 -0.0855 0.7082 1.0000
1.500 0.5966 0.01438 0.00658 -0.0844 0.6847 1.0000
1.750 0.6196 0.01457 0.00657 -0.0830 0.6599 1.0000
2.000 0.6447 0.01480 0.00656 -0.0820 0.6388 1.0000
2.250 0.6678 0.01514 0.00673 -0.0808 0.6183 1.0000
2.500 0.6912 0.01550 0.00694 -0.0797 0.6005 1.0000
2.750 0.7151 0.01588 0.00718 -0.0788 0.5852 1.0000
3.000 0.7395 0.01627 0.00742 -0.0779 0.5717 1.0000
3.250 0.7631 0.01665 0.00772 -0.0770 0.5587 1.0000
3.500 0.7859 0.01706 0.00810 -0.0760 0.5467 1.0000
3.750 0.8103 0.01748 0.00845 -0.0753 0.5367 1.0000
4.000 0.8348 0.01789 0.00880 -0.0746 0.5272 1.0000
4.250 0.8577 0.01834 0.00927 -0.0737 0.5176 1.0000
4.500 0.8839 0.01877 0.00960 -0.0732 0.5098 1.0000
4.750 0.9057 0.01928 0.01020 -0.0722 0.5010 1.0000
5.000 0.9324 0.01974 0.01056 -0.0719 0.4939 1.0000
5.250 0.9535 0.02029 0.01123 -0.0707 0.4853 1.0000
5.500 0.9803 0.02077 0.01163 -0.0705 0.4783 1.0000
5.750 1.0005 0.02132 0.01231 -0.0692 0.4691 1.0000
6.000 1.0264 0.02177 0.01268 -0.0687 0.4606 1.0000
6.250 1.0471 0.02221 0.01321 -0.0674 0.4500 1.0000
6.500 1.0690 0.02269 0.01372 -0.0663 0.4399 1.0000
6.750 1.0955 0.02308 0.01398 -0.0659 0.4301 1.0000
7.000 1.1134 0.02360 0.01469 -0.0642 0.4199 1.0000
7.250 1.1379 0.02413 0.01521 -0.0636 0.4115 1.0000
7.500 1.1578 0.02468 0.01591 -0.0623 0.4027 1.0000
7.750 1.1806 0.02524 0.01650 -0.0614 0.3941 1.0000
8.000 1.2019 0.02565 0.01700 -0.0603 0.3843 1.0000
8.250 1.2205 0.02616 0.01764 -0.0587 0.3741 1.0000
8.500 1.2456 0.02648 0.01790 -0.0581 0.3646 1.0000
8.750 1.2610 0.02694 0.01859 -0.0561 0.3542 1.0000
9.000 1.2806 0.02738 0.01915 -0.0547 0.3446 1.0000
9.250 1.3039 0.02750 0.01921 -0.0537 0.3338 1.0000
9.500 1.3158 0.02778 0.01972 -0.0511 0.3215 1.0000
9.750 1.3296 0.02803 0.02012 -0.0487 0.3091 1.0000
10.000 1.3429 0.02818 0.02040 -0.0463 0.2961 1.0000
10.250 1.3535 0.02833 0.02065 -0.0434 0.2820 1.0000
10.500 1.3593 0.02851 0.02096 -0.0398 0.2656 1.0000
10.750 1.3599 0.02877 0.02129 -0.0355 0.2464 1.0000
11.000 1.3534 0.02929 0.02178 -0.0303 0.2255 1.0000
11.250 1.3438 0.03034 0.02274 -0.0254 0.2015 1.0000
11.500 1.3338 0.03201 0.02427 -0.0212 0.1787 1.0000
11.750 1.3240 0.03408 0.02618 -0.0177 0.1609 1.0000
12.000 1.3165 0.03632 0.02831 -0.0149 0.1469 1.0000
12.250 1.3115 0.03863 0.03060 -0.0126 0.1353 1.0000
12.500 1.3083 0.04099 0.03292 -0.0108 0.1262 1.0000
12.750 1.3064 0.04335 0.03518 -0.0092 0.1181 1.0000
13.000 1.3047 0.04584 0.03781 -0.0079 0.1105 1.0000
13.250 1.3047 0.04828 0.04016 -0.0068 0.1035 1.0000
13.500 1.3020 0.05105 0.04309 -0.0059 0.0970 1.0000
13.750 1.3027 0.05364 0.04564 -0.0051 0.0908 1.0000
14.000 1.3003 0.05659 0.04874 -0.0046 0.0854 1.0000
14.250 1.3048 0.05906 0.05111 -0.0038 0.0797 1.0000
14.500 1.3034 0.06220 0.05443 -0.0035 0.0754 1.0000
14.750 1.3205 0.06394 0.05597 -0.0022 0.0699 1.0000
15.000 1.3134 0.06772 0.06004 -0.0023 0.0673 1.0000
15.250 1.3120 0.07112 0.06360 -0.0023 0.0646 1.0000
15.500 1.3445 0.07285 0.06514 -0.0006 0.0601 1.0000
15.750 1.3302 0.07731 0.06991 -0.0012 0.0595 1.0000
16.000 1.3153 0.08217 0.07507 -0.0023 0.0589 1.0000
16.250 1.2990 0.08746 0.08065 -0.0039 0.0584 1.0000
16.500 1.2797 0.09348 0.08694 -0.0062 0.0581 1.0000
16.750 1.2577 0.10017 0.09390 -0.0093 0.0579 1.0000
17.000 1.2329 0.10772 0.10170 -0.0133 0.0579 1.0000
17.250 1.2056 0.11630 0.11051 -0.0181 0.0583 1.0000
17.500 1.1735 0.12635 0.12078 -0.0242 0.0589 1.0000
17.750 1.1415 0.13723 0.13182 -0.0309 0.0597 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 436 AIRFOIL (goe436-il)