GOE 767 AIRFOIL (goe767-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 767 AIRFOIL (goe767-il) Reynolds number: 100,000 Max Cl/Cd: 38.58 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe767-il-100000-n5.txt Download as CSV file: xf-goe767-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 767 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.8890 0.05441 0.04794 -0.0318 1.0000 0.0537
-11.250 -0.9184 0.04983 0.04309 -0.0309 1.0000 0.0539
-11.000 -0.9326 0.04632 0.03930 -0.0292 1.0000 0.0545
-10.750 -0.9381 0.04303 0.03567 -0.0276 1.0000 0.0553
-10.500 -0.9381 0.03992 0.03218 -0.0260 1.0000 0.0562
-10.250 -0.9313 0.03733 0.02925 -0.0244 1.0000 0.0572
-10.000 -0.9150 0.03612 0.02801 -0.0233 1.0000 0.0580
-9.750 -0.8987 0.03484 0.02666 -0.0221 1.0000 0.0590
-9.500 -0.8823 0.03343 0.02511 -0.0209 1.0000 0.0603
-9.250 -0.8658 0.03187 0.02333 -0.0197 1.0000 0.0621
-9.000 -0.8483 0.03034 0.02157 -0.0185 1.0000 0.0642
-8.750 -0.8279 0.02958 0.02083 -0.0175 1.0000 0.0659
-8.500 -0.8078 0.02858 0.01973 -0.0164 1.0000 0.0682
-8.250 -0.7877 0.02721 0.01812 -0.0152 1.0000 0.0711
-8.000 -0.7664 0.02647 0.01742 -0.0142 1.0000 0.0734
-7.750 -0.7447 0.02566 0.01654 -0.0132 1.0000 0.0768
-7.500 -0.7228 0.02480 0.01558 -0.0121 1.0000 0.0811
-7.250 -0.7005 0.02421 0.01498 -0.0112 1.0000 0.0855
-7.000 -0.6780 0.02347 0.01414 -0.0101 1.0000 0.0908
-6.750 -0.6552 0.02292 0.01358 -0.0091 1.0000 0.0959
-6.500 -0.6322 0.02227 0.01282 -0.0081 1.0000 0.1016
-6.250 -0.6092 0.02181 0.01239 -0.0072 1.0000 0.1071
-6.000 -0.5858 0.02127 0.01174 -0.0062 1.0000 0.1130
-5.750 -0.5628 0.02083 0.01137 -0.0052 1.0000 0.1182
-5.500 -0.5392 0.02041 0.01083 -0.0042 1.0000 0.1248
-5.250 -0.5164 0.02000 0.01054 -0.0032 1.0000 0.1308
-5.000 -0.4930 0.01969 0.01016 -0.0022 1.0000 0.1385
-4.750 -0.4702 0.01927 0.00986 -0.0012 1.0000 0.1447
-4.500 -0.4469 0.01886 0.00946 -0.0002 1.0000 0.1506
-4.250 -0.4236 0.01839 0.00899 0.0009 1.0000 0.1553
-4.000 -0.3983 0.01791 0.00863 0.0014 0.9963 0.1601
-3.750 -0.3523 0.01741 0.00816 -0.0020 0.9564 0.1672
-3.500 -0.3010 0.01690 0.00766 -0.0063 0.8819 0.1752
-3.250 -0.2570 0.01667 0.00714 -0.0085 0.7586 0.1844
-3.000 -0.2318 0.01660 0.00674 -0.0072 0.6755 0.1918
-2.750 -0.2073 0.01652 0.00643 -0.0060 0.6282 0.2017
-2.500 -0.1827 0.01634 0.00617 -0.0050 0.5958 0.2133
-2.250 -0.1575 0.01614 0.00594 -0.0042 0.5698 0.2288
-2.000 -0.1320 0.01591 0.00576 -0.0035 0.5477 0.2516
-1.750 -0.1069 0.01565 0.00561 -0.0027 0.5280 0.2919
-1.500 -0.0824 0.01533 0.00548 -0.0019 0.5097 0.3529
-1.250 -0.0585 0.01496 0.00541 -0.0009 0.4929 0.4320
-1.000 -0.0349 0.01461 0.00539 0.0003 0.4776 0.5255
-0.750 -0.0108 0.01434 0.00545 0.0017 0.4637 0.6246
-0.500 0.0148 0.01422 0.00552 0.0031 0.4504 0.7086
-0.250 0.0439 0.01419 0.00564 0.0038 0.4364 0.7774
0.000 0.0791 0.01428 0.00576 0.0034 0.4224 0.8368
0.250 0.1191 0.01446 0.00587 0.0020 0.4086 0.8865
0.500 0.1647 0.01469 0.00601 -0.0006 0.3945 0.9293
0.750 0.2132 0.01492 0.00614 -0.0041 0.3812 0.9599
1.000 0.2609 0.01513 0.00620 -0.0077 0.3695 0.9810
1.250 0.3116 0.01526 0.00618 -0.0121 0.3572 0.9979
1.500 0.3375 0.01540 0.00624 -0.0118 0.3483 1.0000
1.750 0.3594 0.01556 0.00631 -0.0107 0.3407 1.0000
2.000 0.3816 0.01577 0.00641 -0.0095 0.3341 1.0000
2.500 0.4268 0.01619 0.00669 -0.0074 0.3200 1.0000
2.750 0.4498 0.01641 0.00689 -0.0064 0.3122 1.0000
3.000 0.4726 0.01661 0.00702 -0.0053 0.3042 1.0000
3.250 0.4957 0.01685 0.00719 -0.0043 0.2966 1.0000
3.500 0.5189 0.01708 0.00740 -0.0033 0.2887 1.0000
3.750 0.5421 0.01735 0.00756 -0.0022 0.2828 1.0000
4.000 0.5658 0.01763 0.00789 -0.0013 0.2758 1.0000
4.250 0.5892 0.01791 0.00813 -0.0004 0.2696 1.0000
4.500 0.6126 0.01823 0.00841 0.0006 0.2639 1.0000
4.750 0.6363 0.01855 0.00877 0.0015 0.2572 1.0000
5.000 0.6597 0.01886 0.00903 0.0025 0.2514 1.0000
5.250 0.6831 0.01922 0.00942 0.0034 0.2454 1.0000
5.500 0.7066 0.01957 0.00980 0.0043 0.2386 1.0000
5.750 0.7298 0.01993 0.01006 0.0052 0.2331 1.0000
6.000 0.7533 0.02035 0.01061 0.0061 0.2263 1.0000
6.250 0.7765 0.02074 0.01100 0.0070 0.2200 1.0000
6.500 0.7995 0.02118 0.01142 0.0078 0.2142 1.0000
6.750 0.8225 0.02165 0.01200 0.0087 0.2071 1.0000
7.000 0.8453 0.02208 0.01239 0.0096 0.2014 1.0000
7.250 0.8679 0.02264 0.01306 0.0105 0.1951 1.0000
7.500 0.8904 0.02315 0.01363 0.0113 0.1889 1.0000
7.750 0.9127 0.02366 0.01409 0.0122 0.1840 1.0000
8.000 0.9345 0.02430 0.01491 0.0131 0.1776 1.0000
8.250 0.9564 0.02483 0.01547 0.0140 0.1727 1.0000
8.500 0.9777 0.02548 0.01616 0.0149 0.1680 1.0000
8.750 0.9985 0.02619 0.01700 0.0158 0.1627 1.0000
9.000 1.0195 0.02675 0.01757 0.0167 0.1585 1.0000
9.250 1.0395 0.02752 0.01842 0.0177 0.1543 1.0000
9.500 1.0587 0.02831 0.01934 0.0186 0.1496 1.0000
9.750 1.0781 0.02890 0.01993 0.0196 0.1455 1.0000
10.000 1.0957 0.02977 0.02090 0.0207 0.1412 1.0000
10.250 1.1124 0.03063 0.02187 0.0218 0.1365 1.0000
10.500 1.1294 0.03131 0.02252 0.0229 0.1327 1.0000
10.750 1.1432 0.03242 0.02379 0.0241 0.1286 1.0000
11.000 1.1559 0.03351 0.02500 0.0254 0.1246 1.0000
11.250 1.1687 0.03445 0.02594 0.0266 0.1213 1.0000
11.500 1.1784 0.03570 0.02727 0.0280 0.1182 1.0000
11.750 1.1827 0.03726 0.02904 0.0296 0.1151 1.0000
12.000 1.1845 0.03871 0.03058 0.0315 0.1123 1.0000
12.250 1.1871 0.04016 0.03205 0.0330 0.1099 1.0000
12.500 1.1887 0.04191 0.03382 0.0340 0.1077 1.0000
12.750 1.1773 0.04510 0.03727 0.0339 0.1056 1.0000
13.000 1.1654 0.04890 0.04127 0.0325 0.1035 1.0000
13.250 1.1543 0.05308 0.04558 0.0304 0.1014 1.0000
13.500 1.1480 0.05682 0.04940 0.0285 0.0994 1.0000
13.750 1.1536 0.05894 0.05145 0.0281 0.0974 1.0000
14.000 1.1164 0.06766 0.06048 0.0224 0.0960 1.0000
14.250 1.0259 0.08608 0.07930 0.0103 0.0948 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 767 AIRFOIL (goe767-il)