GOE 505 AIRFOIL (goe505-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 505 AIRFOIL (goe505-il) Reynolds number: 500,000 Max Cl/Cd: 73.5 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe505-il-500000.txt Download as CSV file: xf-goe505-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 505 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 0.0029 0.07545 0.07296 -0.1319 0.9524 0.0441
-9.500 -0.1658 0.03711 0.03392 -0.1598 0.9189 0.0367
-9.250 -0.1671 0.03371 0.03039 -0.1581 0.9125 0.0358
-9.000 -0.1848 0.03011 0.02645 -0.1537 0.9038 0.0354
-8.750 -0.1872 0.02732 0.02330 -0.1505 0.8979 0.0353
-8.500 -0.1799 0.02515 0.02079 -0.1482 0.8938 0.0353
-8.250 -0.1745 0.02379 0.01919 -0.1448 0.8881 0.0356
-8.000 -0.1597 0.02259 0.01774 -0.1429 0.8842 0.0359
-7.750 -0.1405 0.02145 0.01635 -0.1417 0.8810 0.0362
-7.500 -0.1238 0.02053 0.01523 -0.1398 0.8768 0.0363
-7.250 -0.1069 0.01978 0.01430 -0.1378 0.8721 0.0365
-7.000 -0.0866 0.01836 0.01266 -0.1367 0.8678 0.0367
-6.750 -0.0620 0.01700 0.01110 -0.1363 0.8643 0.0370
-6.500 -0.0452 0.01609 0.01011 -0.1342 0.8596 0.0374
-6.250 -0.0230 0.01536 0.00932 -0.1332 0.8549 0.0381
-6.000 0.0037 0.01478 0.00865 -0.1330 0.8509 0.0388
-5.750 0.0287 0.01428 0.00808 -0.1324 0.8468 0.0393
-5.500 0.0495 0.01386 0.00761 -0.1309 0.8418 0.0398
-5.250 0.0750 0.01341 0.00708 -0.1303 0.8369 0.0404
-5.000 0.1031 0.01299 0.00656 -0.1302 0.8319 0.0410
-4.750 0.1226 0.01265 0.00620 -0.1283 0.8253 0.0416
-4.500 0.1483 0.01232 0.00579 -0.1278 0.8204 0.0422
-4.250 0.1769 0.01203 0.00542 -0.1278 0.8168 0.0429
-4.000 0.1974 0.01173 0.00511 -0.1262 0.8125 0.0439
-3.750 0.2202 0.01141 0.00479 -0.1251 0.8081 0.0457
-3.500 0.2465 0.01117 0.00452 -0.1247 0.8038 0.0476
-3.250 0.2754 0.01098 0.00427 -0.1248 0.7998 0.0497
-3.000 0.2954 0.01079 0.00410 -0.1230 0.7945 0.0525
-2.750 0.3196 0.01055 0.00390 -0.1221 0.7890 0.0609
-2.500 0.3468 0.01026 0.00370 -0.1219 0.7845 0.0926
-2.250 0.3671 0.01008 0.00364 -0.1202 0.7789 0.1169
-2.000 0.3900 0.00995 0.00356 -0.1191 0.7730 0.1374
-1.750 0.4175 0.00984 0.00347 -0.1189 0.7678 0.1578
-1.500 0.4366 0.00978 0.00346 -0.1169 0.7608 0.1725
-1.250 0.4604 0.00972 0.00338 -0.1159 0.7539 0.1857
-1.000 0.4814 0.00965 0.00334 -0.1143 0.7461 0.1973
-0.750 0.5032 0.00958 0.00326 -0.1129 0.7374 0.2074
-0.500 0.5222 0.00954 0.00321 -0.1109 0.7267 0.2171
-0.250 0.5413 0.00946 0.00315 -0.1089 0.7152 0.2269
0.250 0.5751 0.00940 0.00302 -0.1040 0.6841 0.2452
0.500 0.5889 0.00943 0.00296 -0.1008 0.6628 0.2549
0.750 0.5973 0.00949 0.00294 -0.0966 0.6382 0.2642
1.000 0.6032 0.00963 0.00295 -0.0918 0.6112 0.2728
1.250 0.6094 0.00982 0.00303 -0.0872 0.5830 0.2829
1.500 0.6153 0.01009 0.00318 -0.0826 0.5541 0.2938
1.750 0.6209 0.01040 0.00337 -0.0780 0.5246 0.3061
2.000 0.6260 0.01075 0.00360 -0.0734 0.4942 0.3228
2.250 0.6341 0.01104 0.00383 -0.0696 0.4680 0.3459
2.500 0.6425 0.01133 0.00408 -0.0658 0.4454 0.3772
2.750 0.6525 0.01150 0.00429 -0.0624 0.4270 0.4246
3.000 0.8177 0.01148 0.00539 -0.0931 0.3905 0.9901
3.250 0.8687 0.01195 0.00572 -0.0984 0.3786 0.9980
3.500 0.8996 0.01224 0.00593 -0.0995 0.3699 1.0000
3.750 0.9093 0.01247 0.00610 -0.0961 0.3641 1.0000
4.000 0.9222 0.01266 0.00626 -0.0932 0.3585 1.0000
4.250 0.9338 0.01291 0.00646 -0.0902 0.3530 1.0000
4.500 0.9467 0.01319 0.00668 -0.0875 0.3488 1.0000
4.750 0.9628 0.01337 0.00686 -0.0853 0.3439 1.0000
5.000 0.9780 0.01361 0.00707 -0.0831 0.3401 1.0000
5.250 0.9927 0.01390 0.00732 -0.0808 0.3361 1.0000
5.500 1.0082 0.01421 0.00759 -0.0786 0.3325 1.0000
5.750 1.0262 0.01442 0.00782 -0.0770 0.3295 1.0000
6.000 1.0438 0.01466 0.00807 -0.0752 0.3263 1.0000
6.250 1.0609 0.01495 0.00833 -0.0735 0.3230 1.0000
6.500 1.0773 0.01528 0.00864 -0.0717 0.3196 1.0000
6.750 1.0943 0.01563 0.00896 -0.0700 0.3161 1.0000
7.000 1.1128 0.01586 0.00923 -0.0685 0.3125 1.0000
7.250 1.1304 0.01615 0.00953 -0.0670 0.3079 1.0000
7.500 1.1457 0.01656 0.00989 -0.0651 0.3027 1.0000
7.750 1.1633 0.01690 0.01024 -0.0637 0.2988 1.0000
8.000 1.1820 0.01717 0.01056 -0.0624 0.2944 1.0000
8.250 1.1990 0.01755 0.01093 -0.0609 0.2899 1.0000
8.500 1.2138 0.01805 0.01139 -0.0592 0.2848 1.0000
8.750 1.2333 0.01832 0.01173 -0.0581 0.2806 1.0000
9.000 1.2503 0.01872 0.01214 -0.0568 0.2746 1.0000
9.250 1.2651 0.01926 0.01264 -0.0551 0.2687 1.0000
9.500 1.2836 0.01961 0.01304 -0.0541 0.2613 1.0000
9.750 1.2974 0.02022 0.01361 -0.0524 0.2539 1.0000
10.000 1.3148 0.02068 0.01408 -0.0513 0.2436 1.0000
10.250 1.3287 0.02133 0.01470 -0.0498 0.2308 1.0000
10.500 1.3389 0.02221 0.01547 -0.0479 0.2117 1.0000
10.750 1.3455 0.02335 0.01645 -0.0456 0.1882 1.0000
11.000 1.3469 0.02487 0.01777 -0.0428 0.1627 1.0000
11.250 1.3494 0.02637 0.01913 -0.0403 0.1423 1.0000
11.500 1.3365 0.02898 0.02140 -0.0364 0.0905 1.0000
11.750 1.3098 0.03278 0.02492 -0.0316 0.0326 1.0000
12.000 1.3099 0.03474 0.02688 -0.0295 0.0223 1.0000
12.250 1.3158 0.03632 0.02852 -0.0279 0.0204 1.0000
12.500 1.3217 0.03793 0.03021 -0.0265 0.0192 1.0000
12.750 1.3277 0.03957 0.03194 -0.0252 0.0186 1.0000
13.000 1.3336 0.04127 0.03373 -0.0240 0.0182 1.0000
13.250 1.3365 0.04331 0.03587 -0.0228 0.0175 1.0000
13.500 1.3380 0.04556 0.03822 -0.0216 0.0170 1.0000
13.750 1.3353 0.04829 0.04106 -0.0205 0.0165 1.0000
14.000 1.3295 0.05150 0.04439 -0.0195 0.0161 1.0000
14.250 1.3226 0.05498 0.04799 -0.0187 0.0159 1.0000
14.500 1.3197 0.05813 0.05125 -0.0183 0.0157 1.0000
14.750 1.3156 0.06153 0.05476 -0.0180 0.0154 1.0000
15.000 1.3094 0.06530 0.05864 -0.0180 0.0152 1.0000
15.250 1.3002 0.06958 0.06305 -0.0181 0.0152 1.0000
15.500 1.2915 0.07393 0.06753 -0.0185 0.0151 1.0000
15.750 1.2796 0.07880 0.07253 -0.0191 0.0150 1.0000
16.000 1.2676 0.08381 0.07766 -0.0199 0.0148 1.0000
16.250 1.2547 0.08907 0.08304 -0.0209 0.0148 1.0000
16.500 1.2409 0.09452 0.08861 -0.0221 0.0146 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 505 AIRFOIL (goe505-il)