GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il) Reynolds number: 200,000 Max Cl/Cd: 60.81 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe235-il-200000.txt Download as CSV file: xf-goe235-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 235 (SCHTTE-LANZ) 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4676 0.11464 0.11088 0.0058 1.0000 0.0368
-9.000 -0.4680 0.11246 0.10873 0.0009 1.0000 0.0371
-8.750 -0.4679 0.10993 0.10624 -0.0042 1.0000 0.0372
-8.500 -0.4590 0.10542 0.10175 -0.0042 1.0000 0.0374
-8.250 -0.4454 0.10108 0.09740 -0.0019 1.0000 0.0378
-8.000 -0.4358 0.09772 0.09407 -0.0019 1.0000 0.0383
-7.750 -0.4278 0.09467 0.09104 -0.0029 1.0000 0.0389
-7.500 -0.4206 0.09167 0.08807 -0.0045 1.0000 0.0396
-7.250 -0.4125 0.08855 0.08497 -0.0070 1.0000 0.0403
-7.000 -0.4013 0.08517 0.08161 -0.0105 1.0000 0.0412
-6.750 -0.3874 0.08163 0.07808 -0.0155 1.0000 0.0423
-6.500 -0.3566 0.07739 0.07370 -0.0319 1.0000 0.0435
-6.250 -0.3398 0.07291 0.06919 -0.0357 1.0000 0.0438
-6.000 -0.3311 0.06940 0.06576 -0.0338 1.0000 0.0441
-5.750 -0.3193 0.06641 0.06280 -0.0333 1.0000 0.0446
-5.500 -0.3059 0.06363 0.06004 -0.0337 1.0000 0.0454
-5.250 -0.2924 0.06098 0.05740 -0.0347 1.0000 0.0465
-5.000 -0.2810 0.05850 0.05491 -0.0354 1.0000 0.0477
-4.750 -0.2733 0.05633 0.05271 -0.0354 1.0000 0.0490
-4.500 -0.2335 0.05334 0.04918 -0.0433 0.9985 0.0513
-4.250 -0.2079 0.04865 0.04464 -0.0459 0.9946 0.0520
-4.000 -0.1762 0.04551 0.04151 -0.0492 0.9896 0.0531
-3.750 -0.1372 0.04249 0.03840 -0.0538 0.9851 0.0552
-3.500 -0.0804 0.03930 0.03457 -0.0610 0.9794 0.0606
-3.250 -0.0449 0.03561 0.03104 -0.0649 0.9739 0.0623
-3.000 -0.0047 0.03323 0.02858 -0.0686 0.9657 0.0658
-2.750 0.0419 0.03063 0.02551 -0.0726 0.9571 0.0724
-2.500 0.0709 0.02848 0.02340 -0.0738 0.9438 0.0748
-2.000 0.1190 0.01010 0.00500 -0.0691 0.8828 0.0880
-1.750 0.1442 0.00927 0.00383 -0.0683 0.8678 0.0997
-1.500 0.1658 0.00818 0.00278 -0.0677 0.8539 0.1046
-1.250 0.1900 0.00738 0.00183 -0.0672 0.8415 0.1189
-1.000 0.2138 0.00673 0.00108 -0.0667 0.8301 0.1376
-0.750 0.2378 0.00610 0.00039 -0.0664 0.8165 0.1662
-0.500 0.2738 0.01907 0.01305 -0.0699 0.8254 0.1892
0.250 0.3685 0.01585 0.00860 -0.0666 0.7856 0.0945
0.500 0.3972 0.01511 0.00763 -0.0656 0.7709 0.0815
0.750 0.4245 0.01433 0.00677 -0.0648 0.7545 0.0794
1.000 0.4520 0.01366 0.00600 -0.0640 0.7373 0.0754
1.250 0.4793 0.01318 0.00544 -0.0633 0.7198 0.0736
1.500 0.5062 0.01273 0.00497 -0.0625 0.7007 0.0736
1.750 0.5326 0.01236 0.00457 -0.0617 0.6741 0.0750
2.000 0.5584 0.01217 0.00428 -0.0608 0.6406 0.0794
2.250 0.5844 0.01212 0.00406 -0.0600 0.6037 0.0821
2.500 0.6111 0.01203 0.00387 -0.0595 0.5582 0.0856
2.750 0.6352 0.01234 0.00368 -0.0586 0.4421 0.0935
3.000 0.6588 0.01303 0.00385 -0.0580 0.3392 0.1157
3.250 0.6802 0.01174 0.00405 -0.0569 0.2881 1.0000
3.500 0.7062 0.01224 0.00427 -0.0565 0.2664 1.0000
3.750 0.7325 0.01266 0.00453 -0.0562 0.2529 1.0000
4.000 0.7589 0.01304 0.00480 -0.0560 0.2400 1.0000
4.250 0.7852 0.01340 0.00509 -0.0557 0.2298 1.0000
4.500 0.8113 0.01379 0.00540 -0.0554 0.2174 1.0000
4.750 0.8375 0.01413 0.00567 -0.0551 0.2020 1.0000
5.000 0.8642 0.01437 0.00588 -0.0549 0.1832 1.0000
5.250 0.8908 0.01465 0.00607 -0.0546 0.1567 1.0000
5.500 0.9134 0.01571 0.00658 -0.0541 0.0750 1.0000
5.750 0.9380 0.01644 0.00729 -0.0535 0.0671 1.0000
6.000 0.9626 0.01711 0.00801 -0.0529 0.0628 1.0000
6.250 0.9860 0.01797 0.00888 -0.0522 0.0594 1.0000
6.500 1.0103 0.01861 0.00960 -0.0516 0.0579 1.0000
6.750 1.0337 0.01935 0.01040 -0.0509 0.0566 1.0000
7.000 1.0564 0.02018 0.01128 -0.0502 0.0554 1.0000
7.250 1.0784 0.02110 0.01222 -0.0493 0.0544 1.0000
7.500 1.1000 0.02210 0.01322 -0.0484 0.0534 1.0000
7.750 1.1211 0.02323 0.01432 -0.0475 0.0522 1.0000
8.000 1.1419 0.02467 0.01572 -0.0466 0.0510 1.0000
8.250 1.1639 0.02644 0.01745 -0.0458 0.0502 1.0000
8.500 1.1873 0.02782 0.01888 -0.0451 0.0499 1.0000
8.750 1.2107 0.02917 0.02033 -0.0444 0.0496 1.0000
9.000 1.2337 0.03059 0.02188 -0.0436 0.0493 1.0000
9.250 1.2558 0.03200 0.02346 -0.0428 0.0487 1.0000
9.500 1.2771 0.03355 0.02520 -0.0419 0.0481 1.0000
9.750 1.2977 0.03548 0.02733 -0.0410 0.0479 1.0000
10.000 1.3169 0.03768 0.02978 -0.0400 0.0479 1.0000
10.250 1.3341 0.04014 0.03251 -0.0388 0.0481 1.0000
10.500 1.3486 0.04270 0.03538 -0.0374 0.0481 1.0000
10.750 1.3602 0.04540 0.03838 -0.0359 0.0479 1.0000
11.000 1.3688 0.04850 0.04181 -0.0342 0.0481 1.0000
11.250 1.3745 0.05256 0.04618 -0.0326 0.0489 1.0000
11.500 1.3917 0.05716 0.05088 -0.0322 0.0506 1.0000
11.750 1.3402 0.06180 0.05666 -0.0258 0.0562 1.0000
12.000 1.2472 0.05462 0.04958 -0.0137 0.0529 1.0000
12.250 1.2018 0.06012 0.05544 -0.0115 0.0541 1.0000
12.500 1.1631 0.06635 0.06193 -0.0113 0.0548 1.0000
12.750 1.1243 0.07342 0.06921 -0.0127 0.0552 1.0000
13.000 1.0824 0.08166 0.07765 -0.0157 0.0555 1.0000
13.250 1.0330 0.09186 0.08803 -0.0210 0.0556 1.0000
13.500 0.9651 0.10690 0.10325 -0.0299 0.0575 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il)