GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 415 AIRFOIL (goe415-il) Reynolds number: 100,000 Max Cl/Cd: 52.03 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe415-il-100000.txt Download as CSV file: xf-goe415-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 415 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4272 0.10556 0.10094 -0.0034 1.0000 0.0606
-8.250 -0.4288 0.10319 0.09866 -0.0083 1.0000 0.0620
-8.000 -0.4309 0.10074 0.09629 -0.0144 1.0000 0.0625
-7.750 -0.4251 0.09716 0.09274 -0.0235 1.0000 0.0628
-7.500 -0.4105 0.09143 0.08704 -0.0099 1.0000 0.0670
-7.250 -0.4046 0.08823 0.08389 -0.0112 1.0000 0.0700
-7.000 -0.3980 0.08466 0.08036 -0.0161 1.0000 0.0737
-6.750 -0.3857 0.08045 0.07606 -0.0331 1.0000 0.0764
-6.500 -0.3805 0.07613 0.07191 -0.0254 1.0000 0.0791
-6.250 -0.3697 0.07312 0.06889 -0.0248 1.0000 0.0839
-6.000 -0.3528 0.06806 0.06361 -0.0383 1.0000 0.0908
-5.750 -0.3460 0.06520 0.06097 -0.0317 1.0000 0.0946
-5.500 -0.3296 0.06096 0.05651 -0.0392 1.0000 0.1049
-5.250 -0.3213 0.05851 0.05425 -0.0346 1.0000 0.1100
-5.000 -0.3078 0.05499 0.05065 -0.0369 1.0000 0.1212
-4.750 -0.2943 0.05203 0.04762 -0.0377 1.0000 0.1346
-4.500 -0.2817 0.04942 0.04500 -0.0373 1.0000 0.1492
-4.250 -0.2695 0.04710 0.04270 -0.0362 1.0000 0.1656
-4.000 -0.2572 0.04496 0.04056 -0.0354 1.0000 0.1920
-3.750 -0.2470 0.04333 0.03900 -0.0329 1.0000 0.2215
-2.750 -0.1348 0.02886 0.02293 -0.0409 1.0000 0.1663
-2.500 -0.0891 0.02435 0.01695 -0.0408 1.0000 0.0843
-2.250 -0.0311 0.02191 0.01384 -0.0451 0.9900 0.0728
-2.000 0.0262 0.02005 0.01160 -0.0499 0.9770 0.0745
-1.750 0.0819 0.01827 0.00969 -0.0543 0.9629 0.0747
-1.500 0.1357 0.01665 0.00819 -0.0588 0.9474 0.0838
-1.250 0.1856 0.01552 0.00700 -0.0622 0.9285 0.0968
-1.000 0.2245 0.01366 0.00633 -0.0638 0.9037 0.3782
-0.750 0.2704 0.01181 0.00575 -0.0650 0.8766 1.0000
-0.500 0.2985 0.01180 0.00538 -0.0639 0.8413 1.0000
-0.250 0.3239 0.01182 0.00507 -0.0624 0.8049 1.0000
0.000 0.3490 0.01188 0.00480 -0.0610 0.7724 1.0000
0.250 0.3740 0.01204 0.00464 -0.0597 0.7416 1.0000
0.500 0.3992 0.01228 0.00457 -0.0586 0.7138 1.0000
0.750 0.4246 0.01258 0.00455 -0.0576 0.6885 1.0000
1.000 0.4501 0.01289 0.00463 -0.0569 0.6637 1.0000
1.250 0.4758 0.01321 0.00474 -0.0562 0.6413 1.0000
1.500 0.5015 0.01354 0.00486 -0.0555 0.6205 1.0000
1.750 0.5274 0.01389 0.00503 -0.0549 0.6009 1.0000
2.000 0.5534 0.01424 0.00523 -0.0544 0.5816 1.0000
2.250 0.5793 0.01463 0.00545 -0.0539 0.5639 1.0000
2.500 0.6054 0.01504 0.00570 -0.0534 0.5476 1.0000
2.750 0.6315 0.01547 0.00605 -0.0530 0.5318 1.0000
3.000 0.6576 0.01592 0.00641 -0.0526 0.5167 1.0000
3.250 0.6837 0.01636 0.00678 -0.0522 0.5025 1.0000
3.500 0.7099 0.01682 0.00716 -0.0519 0.4893 1.0000
3.750 0.7361 0.01727 0.00764 -0.0516 0.4769 1.0000
4.000 0.7622 0.01774 0.00814 -0.0513 0.4648 1.0000
4.250 0.7881 0.01820 0.00861 -0.0509 0.4528 1.0000
4.500 0.8139 0.01862 0.00905 -0.0505 0.4409 1.0000
4.750 0.8396 0.01900 0.00947 -0.0501 0.4290 1.0000
5.000 0.8652 0.01933 0.00979 -0.0496 0.4169 1.0000
5.250 0.8909 0.01968 0.01012 -0.0492 0.4057 1.0000
5.500 0.9162 0.02013 0.01071 -0.0488 0.3947 1.0000
5.750 0.9417 0.02065 0.01138 -0.0484 0.3848 1.0000
6.000 0.9676 0.02114 0.01189 -0.0481 0.3761 1.0000
6.250 0.9928 0.02166 0.01256 -0.0476 0.3659 1.0000
6.500 1.0175 0.02218 0.01325 -0.0472 0.3548 1.0000
6.750 1.0412 0.02199 0.01293 -0.0462 0.3360 1.0000
7.000 1.0631 0.02194 0.01306 -0.0452 0.3156 1.0000
7.250 1.0864 0.02219 0.01343 -0.0445 0.3007 1.0000
7.500 1.1082 0.02211 0.01339 -0.0435 0.2806 1.0000
7.750 1.1298 0.02219 0.01370 -0.0426 0.2600 1.0000
8.000 1.1512 0.02230 0.01392 -0.0416 0.2372 1.0000
8.250 1.1723 0.02253 0.01431 -0.0407 0.1990 1.0000
8.500 1.1907 0.02354 0.01507 -0.0396 0.1347 1.0000
8.750 1.1990 0.02615 0.01706 -0.0379 0.0594 1.0000
9.000 1.2094 0.02839 0.01919 -0.0361 0.0472 1.0000
9.250 1.2211 0.03024 0.02126 -0.0344 0.0429 1.0000
9.500 1.2273 0.03245 0.02365 -0.0324 0.0394 1.0000
9.750 1.2302 0.03471 0.02609 -0.0303 0.0374 1.0000
10.000 1.2336 0.03673 0.02832 -0.0281 0.0363 1.0000
10.250 1.2328 0.03883 0.03060 -0.0256 0.0357 1.0000
10.500 1.2321 0.04114 0.03310 -0.0237 0.0353 1.0000
10.750 1.2329 0.04356 0.03563 -0.0220 0.0349 1.0000
11.000 1.2369 0.04594 0.03813 -0.0204 0.0346 1.0000
11.250 1.2452 0.04826 0.04055 -0.0187 0.0343 1.0000
11.500 1.2567 0.05070 0.04312 -0.0170 0.0343 1.0000
11.750 1.2681 0.05350 0.04612 -0.0154 0.0345 1.0000
12.000 1.2743 0.05671 0.04958 -0.0142 0.0345 1.0000
12.250 1.2741 0.06028 0.05340 -0.0135 0.0343 1.0000
12.500 1.2690 0.06421 0.05759 -0.0135 0.0341 1.0000
12.750 1.2605 0.06857 0.06222 -0.0141 0.0341 1.0000
13.000 1.2488 0.07344 0.06735 -0.0155 0.0341 1.0000
13.250 1.2346 0.07886 0.07303 -0.0177 0.0343 1.0000
13.500 1.2182 0.08485 0.07926 -0.0205 0.0345 1.0000
13.750 1.1999 0.09144 0.08608 -0.0241 0.0349 1.0000
14.000 1.1801 0.09856 0.09340 -0.0282 0.0353 1.0000
14.250 1.1592 0.10623 0.10125 -0.0328 0.0358 1.0000
14.500 1.1378 0.11445 0.10962 -0.0379 0.0363 1.0000
14.750 1.1167 0.12314 0.11844 -0.0433 0.0369 1.0000
15.000 1.0975 0.13200 0.12739 -0.0484 0.0376 1.0000
15.250 1.0811 0.14074 0.13622 -0.0534 0.0382 1.0000
15.500 0.8772 0.12799 0.12368 -0.0320 0.0384 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 415 AIRFOIL (goe415-il)