GOE 15K AIRFOIL (goe15k-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 15K AIRFOIL (goe15k-il) Reynolds number: 200,000 Max Cl/Cd: 49.42 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe15k-il-200000.txt Download as CSV file: xf-goe15k-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 15K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6169 0.08130 0.07760 -0.0860 0.9625 0.0540
-10.500 -0.6671 0.07752 0.07367 -0.0817 0.9604 0.0542
-10.250 -0.6731 0.07495 0.07115 -0.0790 0.9597 0.0532
-10.000 -0.7072 0.07257 0.06869 -0.0727 0.9601 0.0529
-9.750 -0.7349 0.07050 0.06656 -0.0663 0.9595 0.0522
-9.500 -0.7808 0.07243 0.06867 -0.0522 0.9791 0.0516
-9.250 -0.8194 0.07000 0.06604 -0.0442 0.9769 0.0520
-9.000 -0.8523 0.06681 0.06262 -0.0365 0.9760 0.0515
-8.750 -0.8809 0.06394 0.05949 -0.0285 0.9755 0.0519
-8.500 -0.9345 0.05744 0.05246 -0.0158 0.9744 0.0486
-8.250 -0.9390 0.05589 0.05078 -0.0107 0.9725 0.0495
-8.000 -0.9420 0.05401 0.04865 -0.0058 0.9692 0.0512
-7.750 -0.9479 0.05114 0.04533 -0.0001 0.9666 0.0519
-7.500 -0.9479 0.04917 0.04288 0.0049 0.9651 0.0530
-7.250 -0.9575 0.04711 0.04046 0.0121 0.9641 0.0536
-7.000 -0.9627 0.04600 0.03897 0.0187 0.9606 0.0545
-6.750 -0.9566 0.04287 0.03561 0.0221 0.9576 0.0563
-6.500 -0.9367 0.04181 0.03444 0.0232 0.9554 0.0576
-6.250 -0.9142 0.04155 0.03408 0.0237 0.9540 0.0600
-6.000 -0.9088 0.04041 0.03270 0.0281 0.9529 0.0616
-5.750 -0.9047 0.03893 0.03095 0.0329 0.9487 0.0630
-5.500 -0.8897 0.03824 0.02991 0.0355 0.9453 0.0655
-5.250 -0.8607 0.03683 0.02847 0.0346 0.9436 0.0685
-5.000 -0.8300 0.03686 0.02840 0.0335 0.9420 0.0714
-4.750 -0.8183 0.03601 0.02741 0.0365 0.9407 0.0734
-4.500 -0.8073 0.03516 0.02638 0.0398 0.9356 0.0755
-4.250 -0.7742 0.03413 0.02517 0.0383 0.9332 0.0782
-4.000 -0.7387 0.03369 0.02478 0.0360 0.9312 0.0815
-3.750 -0.7005 0.03394 0.02496 0.0335 0.9296 0.0852
-3.500 -0.6845 0.03365 0.02459 0.0354 0.9280 0.0877
-3.250 -0.6662 0.03286 0.02370 0.0370 0.9213 0.0892
-3.000 -0.6254 0.03227 0.02321 0.0335 0.9191 0.0944
-2.750 -0.5868 0.03256 0.02351 0.0308 0.9169 0.0985
-2.500 -0.5479 0.03317 0.02407 0.0282 0.9155 0.1023
-2.250 -0.5439 0.03183 0.02276 0.0324 0.9076 0.1054
-2.000 -0.4976 0.03204 0.02302 0.0282 0.9034 0.1133
-1.750 -0.4366 0.03109 0.02216 0.0220 0.8810 0.1317
-1.500 -0.3926 0.03054 0.02216 0.0179 0.8767 0.2139
-1.250 -0.0846 0.03248 0.02723 -0.0433 0.8876 0.9126
0.000 0.1635 0.03640 0.03067 -0.0687 0.8745 1.0000
0.250 0.1988 0.03655 0.03074 -0.0705 0.8728 1.0000
0.500 0.1951 0.03673 0.03090 -0.0652 0.8639 1.0000
0.750 0.2289 0.03663 0.03074 -0.0665 0.8600 1.0000
1.000 0.2700 0.03653 0.03059 -0.0693 0.8578 1.0000
1.250 0.2726 0.03666 0.03070 -0.0650 0.8481 1.0000
1.500 0.3059 0.03655 0.03055 -0.0662 0.8446 1.0000
1.750 0.3435 0.03644 0.03042 -0.0682 0.8426 1.0000
2.000 0.3469 0.03670 0.03067 -0.0641 0.8331 1.0000
2.250 0.3809 0.03647 0.03043 -0.0654 0.8294 1.0000
2.500 0.4182 0.03627 0.03023 -0.0673 0.8273 1.0000
2.750 0.4583 0.03603 0.02999 -0.0696 0.8258 1.0000
3.000 0.4562 0.03632 0.03029 -0.0644 0.8140 1.0000
3.250 0.4938 0.03599 0.02999 -0.0662 0.8118 1.0000
3.500 0.5332 0.03566 0.02968 -0.0684 0.8104 1.0000
3.750 0.5296 0.03620 0.03025 -0.0631 0.7988 1.0000
4.000 0.5733 0.03543 0.02952 -0.0656 0.7963 1.0000
4.250 0.6278 0.03393 0.02806 -0.0696 0.7938 1.0000
4.500 0.6383 0.03379 0.02798 -0.0663 0.7822 1.0000
4.750 0.6920 0.03215 0.02640 -0.0700 0.7801 1.0000
5.000 0.7091 0.03162 0.02592 -0.0675 0.7687 1.0000
5.250 0.7614 0.02973 0.02411 -0.0708 0.7657 1.0000
5.500 0.8153 0.02740 0.02187 -0.0741 0.7616 1.0000
5.750 0.8314 0.02693 0.02147 -0.0713 0.7506 1.0000
6.000 0.8771 0.02504 0.01968 -0.0734 0.7471 1.0000
6.500 0.8974 0.02461 0.01939 -0.0657 0.7173 1.0000
6.750 0.8937 0.02528 0.02011 -0.0599 0.6931 1.0000
7.000 0.9239 0.02413 0.01899 -0.0592 0.6663 1.0000
7.250 1.0126 0.02049 0.01436 -0.0675 0.5020 1.0000
7.500 0.9954 0.02194 0.01530 -0.0593 0.4333 1.0000
7.750 0.9868 0.02324 0.01630 -0.0530 0.3836 1.0000
8.000 0.9814 0.02451 0.01729 -0.0475 0.3296 1.0000
8.250 0.9734 0.02599 0.01839 -0.0418 0.2648 1.0000
8.750 0.9707 0.02869 0.02052 -0.0331 0.1792 1.0000
9.000 0.9740 0.02990 0.02151 -0.0297 0.1518 1.0000
9.250 0.9815 0.03091 0.02243 -0.0268 0.1323 1.0000
9.500 0.9899 0.03188 0.02331 -0.0241 0.1184 1.0000
9.750 0.9986 0.03287 0.02424 -0.0215 0.1082 1.0000
10.000 1.0101 0.03368 0.02509 -0.0194 0.0986 1.0000
10.250 1.0185 0.03477 0.02612 -0.0169 0.0914 1.0000
10.500 1.0294 0.03566 0.02703 -0.0148 0.0839 1.0000
10.750 1.0397 0.03673 0.02809 -0.0125 0.0785 1.0000
11.000 1.0511 0.03766 0.02905 -0.0105 0.0729 1.0000
11.250 1.0632 0.03878 0.03013 -0.0086 0.0682 1.0000
11.500 1.0761 0.03968 0.03113 -0.0068 0.0639 1.0000
11.750 1.0883 0.04068 0.03208 -0.0052 0.0602 1.0000
12.000 1.1067 0.04167 0.03311 -0.0040 0.0570 1.0000
12.250 1.1203 0.04262 0.03417 -0.0024 0.0542 1.0000
12.500 1.1339 0.04361 0.03517 -0.0009 0.0516 1.0000
12.750 1.1560 0.04477 0.03633 -0.0004 0.0487 1.0000
13.000 1.1643 0.04592 0.03766 0.0017 0.0465 1.0000
13.250 1.1780 0.04706 0.03891 0.0032 0.0449 1.0000
13.500 1.1944 0.04828 0.04021 0.0043 0.0437 1.0000
13.750 1.2185 0.04971 0.04156 0.0041 0.0415 1.0000
14.000 1.2195 0.05138 0.04352 0.0071 0.0403 1.0000
14.250 1.2288 0.05319 0.04556 0.0089 0.0396 1.0000
14.500 1.2318 0.05504 0.04760 0.0113 0.0382 1.0000
14.750 1.2357 0.05669 0.04939 0.0133 0.0372 1.0000
15.000 1.2387 0.05831 0.05107 0.0153 0.0363 1.0000
15.250 1.2435 0.06054 0.05347 0.0171 0.0359 1.0000
15.500 1.2521 0.06325 0.05623 0.0182 0.0348 1.0000
15.750 1.2368 0.06623 0.05953 0.0215 0.0344 1.0000
16.000 1.2248 0.06952 0.06308 0.0241 0.0343 1.0000
16.250 1.2090 0.07319 0.06703 0.0265 0.0341 1.0000
16.500 1.1902 0.07703 0.07114 0.0287 0.0338 1.0000
16.750 1.1656 0.08198 0.07639 0.0304 0.0336 1.0000
17.000 1.1436 0.08706 0.08171 0.0314 0.0337 1.0000
17.250 1.1234 0.09194 0.08681 0.0317 0.0337 1.0000
17.500 1.0986 0.09798 0.09306 0.0313 0.0339 1.0000
17.750 1.0675 0.10531 0.10063 0.0298 0.0340 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 15K AIRFOIL (goe15k-il)