GOE 675 AIRFOIL (goe675-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 675 AIRFOIL (goe675-il) Reynolds number: 200,000 Max Cl/Cd: 75.32 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe675-il-200000.txt Download as CSV file: xf-goe675-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 675 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 -0.0472 0.09449 0.09054 -0.1015 0.9560 0.0830
-9.000 -0.0386 0.09111 0.08715 -0.1057 0.9519 0.0860
-8.750 -0.0684 0.08531 0.08139 -0.1163 0.9405 0.0880
-8.500 -0.0394 0.08252 0.07859 -0.1152 0.9398 0.0887
-8.250 -0.0145 0.08001 0.07606 -0.1159 0.9382 0.0896
-8.000 -0.0124 0.07856 0.07464 -0.1141 0.9295 0.0905
-7.750 -0.1162 0.05363 0.04936 -0.1370 0.9073 0.0707
-7.500 -0.1117 0.04594 0.04135 -0.1424 0.9036 0.0701
-7.250 -0.1373 0.03954 0.03442 -0.1403 0.8910 0.0699
-7.000 -0.1270 0.03255 0.02648 -0.1426 0.8873 0.0713
-6.750 -0.0932 0.03212 0.02614 -0.1439 0.8857 0.0728
-6.500 -0.0907 0.03198 0.02597 -0.1399 0.8760 0.0738
-6.250 -0.0645 0.02986 0.02342 -0.1406 0.8724 0.0770
-6.000 -0.0324 0.02813 0.02141 -0.1420 0.8699 0.0806
-5.750 0.0040 0.02746 0.02072 -0.1434 0.8678 0.0843
-5.500 0.0157 0.02646 0.01932 -0.1409 0.8592 0.0885
-5.250 0.0470 0.02592 0.01889 -0.1415 0.8549 0.0921
-5.000 0.0806 0.02490 0.01755 -0.1423 0.8519 0.0987
-4.750 0.1156 0.02437 0.01713 -0.1434 0.8496 0.1034
-4.500 0.1357 0.02419 0.01676 -0.1419 0.8422 0.1084
-4.250 0.1637 0.02337 0.01592 -0.1419 0.8369 0.1127
-4.000 0.1970 0.02285 0.01537 -0.1426 0.8336 0.1177
-3.750 0.2314 0.02227 0.01453 -0.1433 0.8311 0.1226
-3.500 0.2523 0.02189 0.01426 -0.1423 0.8245 0.1267
-3.250 0.2788 0.02167 0.01403 -0.1419 0.8191 0.1317
-3.000 0.3112 0.02128 0.01346 -0.1423 0.8157 0.1360
-2.750 0.3447 0.02020 0.01232 -0.1431 0.8130 0.1392
-2.500 0.3670 0.01995 0.01210 -0.1421 0.8061 0.1421
-2.250 0.3947 0.01953 0.01165 -0.1418 0.8007 0.1451
-2.000 0.4274 0.01894 0.01098 -0.1423 0.7971 0.1482
-1.750 0.4618 0.01840 0.01032 -0.1429 0.7942 0.1517
-1.500 0.4819 0.01809 0.01008 -0.1415 0.7854 0.1550
-1.250 0.5126 0.01753 0.00956 -0.1417 0.7806 0.1593
-1.000 0.5461 0.01699 0.00898 -0.1423 0.7770 0.1648
-0.750 0.5690 0.01686 0.00886 -0.1412 0.7687 0.1701
-0.500 0.5985 0.01631 0.00841 -0.1412 0.7628 0.1793
-0.250 0.6325 0.01573 0.00784 -0.1419 0.7584 0.1923
0.000 0.6540 0.01557 0.00781 -0.1406 0.7487 0.2097
0.250 0.6854 0.01502 0.00744 -0.1410 0.7428 0.2528
0.500 0.7097 0.01462 0.00738 -0.1403 0.7339 0.3407
0.750 0.7359 0.01377 0.00718 -0.1398 0.7261 0.5341
1.000 0.7714 0.01276 0.00711 -0.1403 0.7168 1.0000
1.250 0.8003 0.01264 0.00683 -0.1401 0.7076 1.0000
1.500 0.8245 0.01267 0.00677 -0.1391 0.6958 1.0000
1.750 0.8521 0.01262 0.00658 -0.1387 0.6845 1.0000
2.000 0.8789 0.01260 0.00642 -0.1382 0.6717 1.0000
2.250 0.9027 0.01266 0.00640 -0.1372 0.6565 1.0000
2.500 0.9270 0.01275 0.00638 -0.1363 0.6406 1.0000
2.750 0.9511 0.01288 0.00637 -0.1354 0.6235 1.0000
3.000 0.9747 0.01306 0.00641 -0.1344 0.6058 1.0000
3.250 0.9977 0.01328 0.00649 -0.1334 0.5883 1.0000
3.500 1.0206 0.01355 0.00663 -0.1324 0.5718 1.0000
3.750 1.0432 0.01385 0.00680 -0.1313 0.5564 1.0000
4.000 1.0660 0.01417 0.00699 -0.1304 0.5426 1.0000
4.500 1.1116 0.01486 0.00747 -0.1286 0.5185 1.0000
4.750 1.1352 0.01523 0.00774 -0.1278 0.5086 1.0000
5.000 1.1582 0.01559 0.00802 -0.1270 0.4988 1.0000
5.250 1.1820 0.01598 0.00835 -0.1264 0.4903 1.0000
5.500 1.2052 0.01635 0.00867 -0.1256 0.4820 1.0000
5.750 1.2303 0.01677 0.00901 -0.1253 0.4746 1.0000
6.000 1.2521 0.01715 0.00940 -0.1243 0.4668 1.0000
6.250 1.2796 0.01761 0.00971 -0.1244 0.4599 1.0000
6.500 1.2996 0.01799 0.01016 -0.1231 0.4527 1.0000
6.750 1.3227 0.01840 0.01054 -0.1225 0.4458 1.0000
7.000 1.3478 0.01887 0.01096 -0.1222 0.4392 1.0000
7.250 1.3666 0.01927 0.01141 -0.1207 0.4319 1.0000
7.500 1.3919 0.01973 0.01176 -0.1205 0.4249 1.0000
7.750 1.4091 0.02016 0.01228 -0.1188 0.4180 1.0000
8.000 1.4277 0.02057 0.01269 -0.1173 0.4109 1.0000
8.250 1.4526 0.02107 0.01312 -0.1171 0.4043 1.0000
8.500 1.4642 0.02147 0.01365 -0.1144 0.3979 1.0000
8.750 1.4821 0.02190 0.01407 -0.1129 0.3916 1.0000
9.000 1.5033 0.02241 0.01456 -0.1120 0.3855 1.0000
9.250 1.5141 0.02289 0.01515 -0.1094 0.3792 1.0000
9.500 1.5306 0.02335 0.01562 -0.1077 0.3730 1.0000
9.750 1.5470 0.02389 0.01619 -0.1061 0.3669 1.0000
10.000 1.5565 0.02446 0.01687 -0.1035 0.3604 1.0000
10.250 1.5722 0.02497 0.01735 -0.1019 0.3541 1.0000
10.500 1.5815 0.02565 0.01814 -0.0994 0.3475 1.0000
10.750 1.5890 0.02634 0.01890 -0.0968 0.3402 1.0000
11.000 1.5996 0.02705 0.01960 -0.0947 0.3329 1.0000
11.250 1.6047 0.02794 0.02062 -0.0920 0.3251 1.0000
11.500 1.6167 0.02869 0.02132 -0.0902 0.3184 1.0000
11.750 1.6228 0.02972 0.02254 -0.0879 0.3112 1.0000
12.000 1.6314 0.03069 0.02353 -0.0859 0.3043 1.0000
12.250 1.6380 0.03181 0.02475 -0.0839 0.2967 1.0000
12.500 1.6435 0.03303 0.02601 -0.0818 0.2885 1.0000
12.750 1.6494 0.03432 0.02739 -0.0799 0.2803 1.0000
13.000 1.6530 0.03577 0.02887 -0.0779 0.2717 1.0000
13.250 1.6563 0.03736 0.03055 -0.0760 0.2619 1.0000
13.500 1.6581 0.03913 0.03233 -0.0741 0.2526 1.0000
13.750 1.6599 0.04103 0.03430 -0.0724 0.2422 1.0000
14.000 1.6609 0.04309 0.03642 -0.0708 0.2314 1.0000
14.250 1.6594 0.04545 0.03881 -0.0692 0.2201 1.0000
14.500 1.6555 0.04816 0.04153 -0.0677 0.2077 1.0000
14.750 1.6492 0.05125 0.04462 -0.0663 0.1939 1.0000
15.000 1.6400 0.05475 0.04813 -0.0650 0.1785 1.0000
15.250 1.6269 0.05886 0.05222 -0.0639 0.1612 1.0000
15.500 1.6089 0.06371 0.05703 -0.0630 0.1430 1.0000
15.750 1.5877 0.06914 0.06241 -0.0624 0.1233 1.0000
16.000 1.5632 0.07523 0.06843 -0.0622 0.1060 1.0000
16.250 1.5395 0.08146 0.07463 -0.0623 0.0934 1.0000
16.500 1.5168 0.08773 0.08090 -0.0628 0.0851 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 675 AIRFOIL (goe675-il)