GOE 366 AIRFOIL (goe366-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 366 AIRFOIL (goe366-il) Reynolds number: 500,000 Max Cl/Cd: 97.49 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe366-il-500000-n5.txt Download as CSV file: xf-goe366-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 366 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 0.0403 0.09192 0.08866 -0.0942 0.7966 0.0202
-8.750 0.0483 0.08930 0.08600 -0.0958 0.7874 0.0203
-8.500 0.0561 0.08670 0.08334 -0.0975 0.7779 0.0203
-8.250 0.0640 0.08407 0.08068 -0.0994 0.7690 0.0203
-8.000 0.0715 0.08144 0.07799 -0.1016 0.7605 0.0204
-7.750 0.0793 0.07875 0.07528 -0.1039 0.7532 0.0204
-7.500 0.0891 0.07616 0.07267 -0.1051 0.7463 0.0204
-7.250 0.0997 0.07384 0.07031 -0.1055 0.7407 0.0205
-7.000 0.1102 0.07154 0.06801 -0.1068 0.7354 0.0206
-6.750 0.1234 0.06909 0.06554 -0.1091 0.7300 0.0207
-6.500 0.1378 0.06666 0.06307 -0.1117 0.7248 0.0209
-6.250 0.1538 0.06411 0.06051 -0.1148 0.7201 0.0212
-6.000 0.1715 0.06143 0.05781 -0.1183 0.7150 0.0215
-5.750 0.1906 0.05866 0.05500 -0.1220 0.7099 0.0218
-5.250 0.2346 0.05280 0.04903 -0.1301 0.7003 0.0226
-5.000 0.2629 0.04932 0.04545 -0.1355 0.6949 0.0229
-4.750 0.2923 0.04585 0.04183 -0.1404 0.6896 0.0230
-4.500 0.3201 0.04275 0.03858 -0.1436 0.6845 0.0230
-4.250 0.3433 0.04023 0.03600 -0.1453 0.6785 0.0231
-4.000 0.3635 0.03846 0.03419 -0.1459 0.6718 0.0232
-3.750 0.3862 0.03671 0.03237 -0.1469 0.6655 0.0235
-3.500 0.4112 0.03483 0.03041 -0.1481 0.6574 0.0237
-3.250 0.4369 0.03293 0.02837 -0.1492 0.6489 0.0240
-3.000 0.4635 0.03103 0.02635 -0.1502 0.6386 0.0244
-2.750 0.4904 0.02918 0.02433 -0.1509 0.6275 0.0250
-2.500 0.5230 0.02665 0.02142 -0.1514 0.6138 0.0259
-2.250 0.5498 0.02484 0.01928 -0.1513 0.5953 0.0260
-2.000 0.5721 0.02345 0.01772 -0.1510 0.5736 0.0261
-1.750 0.5940 0.02251 0.01661 -0.1507 0.5540 0.0263
-1.500 0.6173 0.02171 0.01568 -0.1504 0.5388 0.0266
-1.250 0.6417 0.02092 0.01472 -0.1501 0.5262 0.0273
-0.750 0.6966 0.01908 0.01215 -0.1488 0.5051 0.0292
-0.500 0.7211 0.01805 0.01100 -0.1486 0.4961 0.0293
-0.250 0.7467 0.01729 0.01019 -0.1485 0.4884 0.0296
0.000 0.7738 0.01625 0.00886 -0.1478 0.4812 0.0258
0.250 0.7998 0.01567 0.00818 -0.1476 0.4746 0.0255
0.500 0.8259 0.01518 0.00755 -0.1472 0.4673 0.0254
0.750 0.8513 0.01485 0.00704 -0.1466 0.4604 0.0256
1.500 0.9284 0.01415 0.00600 -0.1453 0.4423 0.0260
1.750 0.9546 0.01390 0.00573 -0.1450 0.4371 0.0260
2.250 1.0040 0.01358 0.00535 -0.1440 0.4252 0.0263
2.500 1.0296 0.01347 0.00526 -0.1436 0.4201 0.0266
2.750 1.0546 0.01343 0.00521 -0.1432 0.4147 0.0270
3.000 1.0788 0.01345 0.00521 -0.1426 0.4097 0.0274
3.250 1.1032 0.01350 0.00524 -0.1421 0.4053 0.0285
3.500 1.1281 0.01354 0.00529 -0.1416 0.4006 0.0292
3.750 1.1522 0.01357 0.00533 -0.1411 0.3955 0.0296
4.000 1.1749 0.01370 0.00543 -0.1403 0.3905 0.0301
4.250 1.1987 0.01377 0.00553 -0.1396 0.3863 0.0308
4.500 1.2213 0.01388 0.00565 -0.1388 0.3819 0.0316
4.750 1.2424 0.01404 0.00579 -0.1376 0.3776 0.0326
5.000 1.2628 0.01424 0.00597 -0.1364 0.3735 0.0348
5.250 1.2858 0.01440 0.00615 -0.1357 0.3694 0.0363
5.500 1.3082 0.01461 0.00635 -0.1348 0.3642 0.0385
5.750 1.3291 0.01488 0.00661 -0.1338 0.3588 0.0447
6.000 1.3502 0.01385 0.00718 -0.1331 0.3547 1.0000
6.250 1.3723 0.01411 0.00742 -0.1323 0.3507 1.0000
6.500 1.3934 0.01441 0.00771 -0.1314 0.3463 1.0000
6.750 1.4131 0.01477 0.00804 -0.1302 0.3419 1.0000
7.000 1.4335 0.01511 0.00837 -0.1292 0.3383 1.0000
7.250 1.4550 0.01541 0.00868 -0.1284 0.3349 1.0000
7.500 1.4753 0.01576 0.00905 -0.1274 0.3308 1.0000
7.750 1.4941 0.01619 0.00945 -0.1262 0.3264 1.0000
8.000 1.5126 0.01664 0.00989 -0.1250 0.3225 1.0000
8.250 1.5331 0.01699 0.01029 -0.1242 0.3190 1.0000
8.500 1.5527 0.01741 0.01072 -0.1232 0.3153 1.0000
8.750 1.5707 0.01790 0.01122 -0.1221 0.3113 1.0000
9.000 1.5873 0.01847 0.01178 -0.1207 0.3073 1.0000
9.250 1.6067 0.01891 0.01227 -0.1199 0.3036 1.0000
9.500 1.6242 0.01946 0.01284 -0.1187 0.2978 1.0000
9.750 1.6387 0.02017 0.01353 -0.1173 0.2920 1.0000
10.000 1.6562 0.02073 0.01413 -0.1162 0.2873 1.0000
10.250 1.6726 0.02137 0.01480 -0.1151 0.2816 1.0000
10.500 1.6849 0.02226 0.01567 -0.1135 0.2752 1.0000
10.750 1.7015 0.02291 0.01637 -0.1124 0.2688 1.0000
11.000 1.7136 0.02385 0.01730 -0.1109 0.2617 1.0000
11.250 1.7276 0.02469 0.01816 -0.1097 0.2558 1.0000
11.500 1.7401 0.02564 0.01914 -0.1083 0.2488 1.0000
11.750 1.7498 0.02681 0.02030 -0.1067 0.2394 1.0000
12.000 1.7557 0.02828 0.02173 -0.1048 0.2262 1.0000
12.250 1.7588 0.03000 0.02340 -0.1028 0.2104 1.0000
12.500 1.7587 0.03205 0.02538 -0.1007 0.1924 1.0000
12.750 1.7550 0.03448 0.02773 -0.0984 0.1737 1.0000
13.000 1.7496 0.03716 0.03033 -0.0962 0.1555 1.0000
13.250 1.7439 0.03998 0.03311 -0.0942 0.1405 1.0000
13.500 1.7403 0.04270 0.03581 -0.0925 0.1304 1.0000
13.750 1.7381 0.04536 0.03849 -0.0910 0.1229 1.0000
14.000 1.7373 0.04797 0.04114 -0.0898 0.1174 1.0000
14.250 1.7362 0.05067 0.04389 -0.0886 0.1123 1.0000
14.500 1.7323 0.05373 0.04698 -0.0875 0.1075 1.0000
14.750 1.7324 0.05641 0.04973 -0.0867 0.1036 1.0000
15.000 1.7292 0.05954 0.05293 -0.0859 0.0993 1.0000
15.250 1.7230 0.06311 0.05655 -0.0852 0.0953 1.0000
15.500 1.7204 0.06633 0.05985 -0.0847 0.0911 1.0000
15.750 1.7131 0.07018 0.06375 -0.0842 0.0857 1.0000
16.000 1.7058 0.07411 0.06774 -0.0839 0.0816 1.0000
16.250 1.6974 0.07827 0.07196 -0.0838 0.0758 1.0000
16.500 1.6855 0.08295 0.07668 -0.0838 0.0703 1.0000
17.000 1.6577 0.09297 0.08679 -0.0842 0.0584 1.0000
17.250 1.6433 0.09816 0.09203 -0.0846 0.0531 1.0000
17.500 1.6278 0.10354 0.09745 -0.0852 0.0473 1.0000
17.750 1.6121 0.10900 0.10295 -0.0860 0.0406 1.0000
18.000 1.5881 0.11576 0.10970 -0.0872 0.0281 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 366 AIRFOIL (goe366-il)