GOE 240 (KOLLER) AIRFOIL (goe240-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 240 (KOLLER) AIRFOIL (goe240-il) Reynolds number: 500,000 Max Cl/Cd: 107.64 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe240-il-500000.txt Download as CSV file: xf-goe240-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 240 (KOLLER) 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3182 0.10184 0.09959 -0.0255 1.0000 0.0178
-8.500 -0.3165 0.09929 0.09708 -0.0274 1.0000 0.0180
-8.250 -0.3194 0.09710 0.09493 -0.0280 1.0000 0.0181
-8.000 -0.2479 0.08062 0.07868 -0.0325 0.9977 0.0184
-7.750 -0.2298 0.07630 0.07435 -0.0333 0.9960 0.0187
-7.500 -0.2142 0.07222 0.07027 -0.0366 0.9932 0.0191
-7.250 -0.3108 0.08433 0.08230 -0.0340 0.9957 0.0184
-7.000 -0.2911 0.08067 0.07863 -0.0353 0.9939 0.0187
-6.750 -0.2662 0.07689 0.07484 -0.0404 0.9897 0.0192
-6.500 -0.2358 0.07278 0.07071 -0.0475 0.9860 0.0202
-6.250 -0.1975 0.06781 0.06568 -0.0578 0.9827 0.0222
-6.000 -0.1459 0.06144 0.05919 -0.0736 0.9741 0.0231
-5.750 -0.1100 0.05625 0.05391 -0.0812 0.9669 0.0232
-5.500 -0.0931 0.05110 0.04872 -0.0841 0.9574 0.0238
-5.250 -0.0739 0.04884 0.04643 -0.0848 0.9441 0.0244
-5.000 -0.0519 0.04646 0.04398 -0.0863 0.9286 0.0255
-4.750 -0.0246 0.04326 0.04066 -0.0893 0.9108 0.0275
-4.500 0.0199 0.03888 0.03594 -0.0947 0.8933 0.0296
-4.250 0.0424 0.03312 0.02994 -0.0974 0.8775 0.0306
-4.000 0.0648 0.03160 0.02830 -0.0977 0.8639 0.0315
-3.750 0.0899 0.03005 0.02661 -0.0983 0.8525 0.0332
-3.500 0.1255 0.02800 0.02415 -0.0992 0.8428 0.0380
-3.250 0.1506 0.02359 0.01944 -0.1008 0.8338 0.0398
-3.000 0.1759 0.02261 0.01835 -0.1010 0.8250 0.0414
-2.750 0.2033 0.02142 0.01698 -0.1012 0.8160 0.0450
-2.500 0.2318 0.01951 0.01474 -0.1015 0.8080 0.0509
-2.250 0.2583 0.01877 0.01391 -0.1016 0.7999 0.0539
-2.000 0.2872 0.01752 0.01234 -0.1016 0.7923 0.0619
-1.750 0.3138 0.01675 0.01152 -0.1016 0.7843 0.0648
-1.500 0.3437 0.01801 0.01254 -0.1009 0.7762 0.0722
-1.250 0.3742 0.01184 0.00564 -0.1008 0.7704 0.0434
-1.000 0.4024 0.01071 0.00429 -0.1004 0.7638 0.0409
-0.750 0.4302 0.01020 0.00370 -0.1002 0.7566 0.0406
-0.500 0.4579 0.00982 0.00326 -0.1000 0.7495 0.0407
-0.250 0.4855 0.00950 0.00291 -0.0998 0.7408 0.0410
0.000 0.5129 0.00925 0.00264 -0.0996 0.7305 0.0422
0.250 0.5404 0.00905 0.00240 -0.0993 0.7196 0.0429
0.500 0.5678 0.00889 0.00218 -0.0990 0.7070 0.0435
0.750 0.5952 0.00876 0.00201 -0.0988 0.6928 0.0446
1.000 0.6227 0.00867 0.00188 -0.0985 0.6776 0.0464
1.250 0.6501 0.00862 0.00179 -0.0983 0.6607 0.0498
1.500 0.6772 0.00859 0.00178 -0.0980 0.6373 0.0690
1.750 0.7035 0.00865 0.00176 -0.0976 0.6060 0.0952
2.000 0.7291 0.00863 0.00183 -0.0973 0.5726 0.1999
2.250 0.7524 0.00720 0.00202 -0.0966 0.5489 1.0000
2.500 0.7786 0.00744 0.00211 -0.0963 0.5285 1.0000
2.750 0.8051 0.00765 0.00220 -0.0960 0.5111 1.0000
3.000 0.8316 0.00785 0.00231 -0.0957 0.4947 1.0000
3.250 0.8581 0.00804 0.00243 -0.0955 0.4776 1.0000
3.500 0.8844 0.00825 0.00255 -0.0952 0.4589 1.0000
3.750 0.9106 0.00846 0.00268 -0.0949 0.4376 1.0000
4.000 0.9363 0.00874 0.00283 -0.0945 0.4085 1.0000
4.250 0.9613 0.00908 0.00302 -0.0941 0.3765 1.0000
4.500 0.9862 0.00943 0.00324 -0.0937 0.3500 1.0000
4.750 1.0111 0.00979 0.00348 -0.0932 0.3266 1.0000
5.000 1.0361 0.01013 0.00372 -0.0928 0.3058 1.0000
5.250 1.0605 0.01054 0.00401 -0.0923 0.2795 1.0000
5.500 1.0845 0.01098 0.00429 -0.0918 0.2459 1.0000
5.750 1.1079 0.01149 0.00461 -0.0912 0.2104 1.0000
6.000 1.1310 0.01202 0.00500 -0.0906 0.1821 1.0000
6.250 1.1544 0.01250 0.00537 -0.0900 0.1636 1.0000
6.500 1.1780 0.01295 0.00574 -0.0895 0.1481 1.0000
6.750 1.2018 0.01335 0.00611 -0.0889 0.1368 1.0000
7.000 1.2253 0.01378 0.00650 -0.0884 0.1244 1.0000
7.250 1.2485 0.01422 0.00688 -0.0877 0.1081 1.0000
7.500 1.2662 0.01526 0.00758 -0.0864 0.0606 1.0000
7.750 1.2877 0.01586 0.00819 -0.0855 0.0496 1.0000
8.000 1.3092 0.01645 0.00880 -0.0845 0.0383 1.0000
8.250 1.3273 0.01738 0.00958 -0.0831 0.0200 1.0000
8.500 1.3473 0.01807 0.01035 -0.0820 0.0177 1.0000
8.750 1.3660 0.01888 0.01124 -0.0806 0.0162 1.0000
9.000 1.3828 0.01982 0.01229 -0.0790 0.0151 1.0000
9.250 1.3976 0.02088 0.01348 -0.0770 0.0144 1.0000
9.500 1.4132 0.02178 0.01449 -0.0753 0.0140 1.0000
9.750 1.4265 0.02280 0.01562 -0.0732 0.0136 1.0000
10.000 1.4364 0.02386 0.01678 -0.0706 0.0131 1.0000
10.250 1.4439 0.02500 0.01802 -0.0678 0.0126 1.0000
10.500 1.4497 0.02630 0.01940 -0.0649 0.0121 1.0000
10.750 1.4532 0.02780 0.02100 -0.0621 0.0117 1.0000
11.000 1.4545 0.02957 0.02286 -0.0592 0.0114 1.0000
11.250 1.4538 0.03161 0.02502 -0.0565 0.0112 1.0000
11.500 1.4511 0.03402 0.02752 -0.0540 0.0110 1.0000
11.750 1.4473 0.03676 0.03038 -0.0517 0.0108 1.0000
12.000 1.4439 0.03973 0.03345 -0.0495 0.0106 1.0000
12.250 1.4473 0.04197 0.03581 -0.0480 0.0105 1.0000
12.500 1.4514 0.04412 0.03809 -0.0469 0.0104 1.0000
12.750 1.4544 0.04647 0.04058 -0.0459 0.0102 1.0000
13.000 1.4562 0.04901 0.04326 -0.0449 0.0101 1.0000
13.250 1.4568 0.05176 0.04614 -0.0439 0.0100 1.0000
13.500 1.4563 0.05471 0.04925 -0.0430 0.0100 1.0000
13.750 1.4542 0.05788 0.05258 -0.0424 0.0099 1.0000
14.000 1.4510 0.06127 0.05614 -0.0419 0.0098 1.0000
14.250 1.4463 0.06491 0.05994 -0.0418 0.0097 1.0000
14.500 1.4400 0.06886 0.06406 -0.0420 0.0096 1.0000
14.750 1.4330 0.07306 0.06843 -0.0426 0.0095 1.0000
15.000 1.4246 0.07764 0.07317 -0.0435 0.0094 1.0000
15.250 1.4150 0.08257 0.07827 -0.0448 0.0093 1.0000
15.500 1.4046 0.08784 0.08371 -0.0466 0.0092 1.0000
15.750 1.3925 0.09362 0.08967 -0.0487 0.0092 1.0000
16.000 1.3796 0.09972 0.09593 -0.0512 0.0092 1.0000
16.250 1.3647 0.10639 0.10279 -0.0541 0.0093 1.0000
16.500 1.3498 0.11338 0.10995 -0.0576 0.0093 1.0000
16.750 1.3339 0.12084 0.11758 -0.0615 0.0094 1.0000
17.000 1.3181 0.12857 0.12548 -0.0658 0.0095 1.0000
17.250 1.3016 0.13676 0.13382 -0.0706 0.0096 1.0000
17.500 1.2856 0.14525 0.14247 -0.0759 0.0097 1.0000
17.750 1.2692 0.15423 0.15158 -0.0815 0.0098 1.0000
18.000 1.2524 0.16367 0.16116 -0.0876 0.0099 1.0000
18.250 1.2351 0.17376 0.17138 -0.0942 0.0101 1.0000
18.500 1.2356 0.17746 0.17510 -0.0954 0.0108 1.0000
18.750 1.2247 0.18723 0.18499 -0.1024 0.0109 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 240 (KOLLER) AIRFOIL (goe240-il)