GOE 496 AIRFOIL (goe496-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 496 AIRFOIL (goe496-il) Reynolds number: 1,000,000 Max Cl/Cd: 150.17 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe496-il-1000000.txt Download as CSV file: xf-goe496-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 496 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.2756 0.12904 0.12744 -0.0353 1.0000 0.0179
-12.500 -0.7921 0.05680 0.05483 -0.0661 1.0000 0.0175
-12.250 -0.8173 0.02818 0.02548 -0.1118 0.9930 0.0175
-12.000 -0.7881 0.02634 0.02356 -0.1155 0.9907 0.0180
-11.750 -0.7565 0.02551 0.02269 -0.1178 0.9889 0.0184
-11.500 -0.7240 0.02485 0.02198 -0.1199 0.9874 0.0189
-11.250 -0.6911 0.02394 0.02098 -0.1225 0.9861 0.0195
-11.000 -0.6574 0.02310 0.02003 -0.1250 0.9850 0.0202
-10.750 -0.6232 0.02227 0.01908 -0.1275 0.9841 0.0207
-10.500 -0.5938 0.02025 0.01680 -0.1304 0.9821 0.0213
-10.250 -0.5670 0.01915 0.01562 -0.1315 0.9785 0.0220
-10.000 -0.5346 0.01876 0.01522 -0.1329 0.9766 0.0225
-9.750 -0.5009 0.01835 0.01476 -0.1346 0.9751 0.0231
-9.500 -0.4665 0.01795 0.01430 -0.1364 0.9739 0.0238
-9.250 -0.4316 0.01730 0.01355 -0.1384 0.9728 0.0246
-9.000 -0.3957 0.01674 0.01287 -0.1405 0.9719 0.0252
-8.750 -0.3597 0.01636 0.01240 -0.1424 0.9708 0.0255
-8.500 -0.3373 0.01449 0.01035 -0.1426 0.9643 0.0267
-8.250 -0.3051 0.01402 0.00986 -0.1438 0.9610 0.0274
-8.000 -0.2722 0.01356 0.00935 -0.1450 0.9581 0.0281
-7.750 -0.2441 0.01310 0.00882 -0.1453 0.9526 0.0287
-7.500 -0.2153 0.01272 0.00837 -0.1455 0.9466 0.0295
-7.250 -0.1836 0.01227 0.00784 -0.1464 0.9421 0.0302
-7.000 -0.1575 0.01190 0.00739 -0.1460 0.9334 0.0306
-6.750 -0.1275 0.01158 0.00699 -0.1464 0.9266 0.0310
-6.500 -0.1019 0.01054 0.00582 -0.1464 0.9167 0.0320
-6.250 -0.0742 0.01004 0.00526 -0.1464 0.9074 0.0329
-5.750 -0.0189 0.00943 0.00452 -0.1462 0.8863 0.0344
-5.500 0.0087 0.00918 0.00419 -0.1461 0.8755 0.0352
-5.250 0.0362 0.00895 0.00388 -0.1459 0.8648 0.0360
-5.000 0.0636 0.00878 0.00363 -0.1457 0.8540 0.0369
-4.750 0.0911 0.00862 0.00340 -0.1455 0.8436 0.0375
-4.500 0.1185 0.00841 0.00311 -0.1453 0.8336 0.0380
-4.250 0.1457 0.00801 0.00262 -0.1451 0.8237 0.0399
-4.000 0.1733 0.00783 0.00238 -0.1450 0.8146 0.0415
-3.500 0.2285 0.00759 0.00203 -0.1446 0.7963 0.0447
-3.250 0.2561 0.00753 0.00191 -0.1444 0.7878 0.0461
-3.000 0.2839 0.00731 0.00166 -0.1442 0.7790 0.0511
-2.750 0.3115 0.00721 0.00151 -0.1441 0.7705 0.0556
-2.500 0.3391 0.00698 0.00142 -0.1440 0.7618 0.0933
-2.250 0.3668 0.00696 0.00140 -0.1438 0.7534 0.1077
-2.000 0.3944 0.00693 0.00136 -0.1436 0.7444 0.1159
-1.750 0.4222 0.00693 0.00133 -0.1434 0.7359 0.1201
-1.500 0.4496 0.00690 0.00129 -0.1432 0.7270 0.1278
-1.250 0.4773 0.00688 0.00125 -0.1430 0.7169 0.1328
-1.000 0.5047 0.00689 0.00122 -0.1428 0.7061 0.1382
-0.750 0.5318 0.00688 0.00119 -0.1425 0.6947 0.1471
-0.500 0.5592 0.00684 0.00117 -0.1423 0.6831 0.1622
-0.250 0.5865 0.00672 0.00117 -0.1421 0.6720 0.2116
0.000 0.6135 0.00664 0.00121 -0.1419 0.6610 0.2712
0.500 0.6675 0.00663 0.00128 -0.1414 0.6357 0.3360
0.750 0.6946 0.00666 0.00132 -0.1411 0.6231 0.3596
1.000 0.7215 0.00669 0.00137 -0.1408 0.6100 0.3860
1.250 0.7483 0.00674 0.00144 -0.1405 0.5974 0.4143
1.500 0.7750 0.00679 0.00151 -0.1402 0.5858 0.4424
1.750 0.8022 0.00683 0.00158 -0.1400 0.5754 0.4704
2.000 0.8292 0.00685 0.00166 -0.1397 0.5656 0.5037
2.250 0.8556 0.00686 0.00176 -0.1394 0.5550 0.5547
2.500 0.8819 0.00676 0.00191 -0.1391 0.5443 0.6659
2.750 0.9057 0.00652 0.00204 -0.1382 0.5345 0.8133
3.250 0.9599 0.00650 0.00220 -0.1375 0.5149 1.0000
3.500 0.9865 0.00664 0.00230 -0.1372 0.5046 1.0000
3.750 1.0127 0.00681 0.00242 -0.1368 0.4928 1.0000
4.000 1.0393 0.00695 0.00253 -0.1365 0.4809 1.0000
4.250 1.0657 0.00710 0.00265 -0.1361 0.4690 1.0000
4.500 1.0917 0.00727 0.00278 -0.1357 0.4556 1.0000
4.750 1.1168 0.00750 0.00294 -0.1351 0.4353 1.0000
5.000 1.1405 0.00784 0.00313 -0.1343 0.4037 1.0000
5.250 1.1617 0.00837 0.00342 -0.1331 0.3552 1.0000
5.500 1.1807 0.00909 0.00381 -0.1316 0.2950 1.0000
5.750 1.2011 0.00970 0.00418 -0.1303 0.2538 1.0000
6.000 1.2237 0.01012 0.00449 -0.1294 0.2322 1.0000
6.250 1.2471 0.01046 0.00477 -0.1286 0.2177 1.0000
6.500 1.2703 0.01081 0.00505 -0.1278 0.2040 1.0000
6.750 1.2941 0.01110 0.00532 -0.1270 0.1927 1.0000
7.000 1.3174 0.01141 0.00559 -0.1262 0.1808 1.0000
7.250 1.3399 0.01178 0.00590 -0.1253 0.1660 1.0000
7.500 1.3606 0.01227 0.00627 -0.1241 0.1452 1.0000
7.750 1.3783 0.01294 0.00676 -0.1224 0.1131 1.0000
8.000 1.3917 0.01391 0.00747 -0.1201 0.0751 1.0000
8.250 1.4072 0.01467 0.00810 -0.1180 0.0542 1.0000
8.500 1.4229 0.01539 0.00872 -0.1160 0.0397 1.0000
8.750 1.4386 0.01602 0.00930 -0.1140 0.0299 1.0000
9.000 1.4529 0.01662 0.00987 -0.1116 0.0243 1.0000
9.250 1.4663 0.01724 0.01048 -0.1091 0.0209 1.0000
9.500 1.4818 0.01775 0.01102 -0.1071 0.0193 1.0000
9.750 1.4954 0.01837 0.01166 -0.1048 0.0178 1.0000
10.000 1.5074 0.01910 0.01243 -0.1022 0.0166 1.0000
10.250 1.5220 0.01968 0.01306 -0.1002 0.0159 1.0000
10.500 1.5353 0.02035 0.01377 -0.0981 0.0152 1.0000
10.750 1.5474 0.02110 0.01456 -0.0959 0.0145 1.0000
11.000 1.5558 0.02212 0.01562 -0.0932 0.0137 1.0000
11.250 1.5644 0.02315 0.01672 -0.0907 0.0132 1.0000
11.500 1.5763 0.02398 0.01762 -0.0888 0.0128 1.0000
11.750 1.5866 0.02495 0.01866 -0.0868 0.0124 1.0000
12.000 1.5961 0.02602 0.01978 -0.0847 0.0120 1.0000
12.250 1.6048 0.02718 0.02101 -0.0828 0.0117 1.0000
12.500 1.6122 0.02848 0.02237 -0.0808 0.0113 1.0000
12.750 1.6174 0.03002 0.02397 -0.0789 0.0110 1.0000
13.000 1.6175 0.03205 0.02610 -0.0767 0.0107 1.0000
13.250 1.6112 0.03477 0.02894 -0.0743 0.0104 1.0000
13.500 1.6184 0.03640 0.03064 -0.0731 0.0102 1.0000
13.750 1.6243 0.03819 0.03251 -0.0719 0.0101 1.0000
14.000 1.6290 0.04017 0.03457 -0.0708 0.0099 1.0000
14.250 1.6327 0.04232 0.03681 -0.0698 0.0097 1.0000
14.500 1.6352 0.04469 0.03927 -0.0690 0.0095 1.0000
14.750 1.6375 0.04715 0.04181 -0.0684 0.0093 1.0000
15.000 1.6392 0.04978 0.04453 -0.0679 0.0091 1.0000
15.250 1.6394 0.05271 0.04754 -0.0675 0.0089 1.0000
15.500 1.6391 0.05579 0.05070 -0.0674 0.0087 1.0000
15.750 1.6363 0.05931 0.05432 -0.0675 0.0086 1.0000
16.000 1.6328 0.06307 0.05817 -0.0678 0.0085 1.0000
16.250 1.6259 0.06746 0.06266 -0.0685 0.0083 1.0000
16.500 1.6181 0.07213 0.06745 -0.0694 0.0082 1.0000
16.750 1.6060 0.07762 0.07305 -0.0707 0.0081 1.0000
17.000 1.5897 0.08396 0.07954 -0.0724 0.0080 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 496 AIRFOIL (goe496-il)