GOE 342 AIRFOIL (goe342-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 342 AIRFOIL (goe342-il) Reynolds number: 500,000 Max Cl/Cd: 105.76 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe342-il-500000-n5.txt Download as CSV file: xf-goe342-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 342 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2117 0.09055 0.08824 -0.0492 0.9546 0.0096
-7.750 -0.2018 0.08713 0.08479 -0.0516 0.9402 0.0097
-7.500 -0.1929 0.08376 0.08140 -0.0540 0.9252 0.0099
-7.250 -0.1793 0.08150 0.07911 -0.0558 0.9128 0.0101
-7.000 -0.1647 0.07891 0.07648 -0.0584 0.9004 0.0103
-6.750 -0.1485 0.07631 0.07383 -0.0612 0.8883 0.0109
-6.500 -0.1305 0.07320 0.07067 -0.0649 0.8762 0.0118
-6.250 -0.1102 0.06951 0.06692 -0.0698 0.8648 0.0121
-6.000 -0.0843 0.06458 0.06191 -0.0772 0.8543 0.0125
-5.500 -0.0378 0.05949 0.05669 -0.0844 0.8342 0.0133
-5.250 -0.0109 0.05673 0.05386 -0.0888 0.8246 0.0145
-5.000 0.0236 0.05244 0.04945 -0.0961 0.8159 0.0155
-4.750 0.0576 0.04872 0.04564 -0.1024 0.8086 0.0163
-4.500 0.0862 0.04646 0.04330 -0.1059 0.8012 0.0169
-4.250 0.1193 0.04375 0.04049 -0.1105 0.7944 0.0182
-4.000 0.1739 0.03817 0.03467 -0.1210 0.7880 0.0201
-3.750 0.1940 0.03730 0.03377 -0.1211 0.7823 0.0209
-3.500 0.2359 0.03538 0.03170 -0.1256 0.7770 0.0246
-3.250 0.2755 0.03240 0.02853 -0.1297 0.7719 0.0248
-3.000 0.3105 0.02957 0.02550 -0.1328 0.7673 0.0248
-2.750 0.3452 0.02697 0.02272 -0.1354 0.7618 0.0249
-2.500 0.3817 0.02320 0.01869 -0.1390 0.7566 0.0256
-2.250 0.4085 0.02224 0.01767 -0.1400 0.7520 0.0263
-2.000 0.4396 0.02080 0.01610 -0.1413 0.7472 0.0270
-1.750 0.4715 0.01926 0.01439 -0.1426 0.7422 0.0276
-1.500 0.5038 0.01771 0.01262 -0.1437 0.7364 0.0280
-1.250 0.5363 0.01608 0.01072 -0.1447 0.7277 0.0276
-1.000 0.5665 0.01510 0.00953 -0.1452 0.7186 0.0284
-0.750 0.5974 0.01399 0.00820 -0.1458 0.7110 0.0283
-0.500 0.6289 0.01266 0.00654 -0.1463 0.7053 0.0269
-0.250 0.6588 0.01189 0.00560 -0.1466 0.6994 0.0268
0.000 0.6877 0.01138 0.00496 -0.1468 0.6931 0.0271
0.250 0.7164 0.01096 0.00446 -0.1470 0.6861 0.0275
0.500 0.7447 0.01063 0.00404 -0.1470 0.6767 0.0280
0.750 0.7728 0.01034 0.00368 -0.1470 0.6630 0.0284
1.000 0.8003 0.01015 0.00339 -0.1469 0.6438 0.0291
1.250 0.8273 0.01006 0.00320 -0.1466 0.6137 0.0301
1.500 0.8522 0.01017 0.00301 -0.1460 0.5543 0.0303
1.750 0.8763 0.01045 0.00298 -0.1454 0.4933 0.0303
2.000 0.9012 0.01070 0.00300 -0.1450 0.4457 0.0304
2.250 0.9266 0.01092 0.00305 -0.1446 0.4075 0.0306
2.500 0.9524 0.01110 0.00310 -0.1444 0.3804 0.0310
2.750 0.9787 0.01123 0.00316 -0.1442 0.3627 0.0316
3.000 1.0050 0.01138 0.00327 -0.1439 0.3488 0.0323
3.500 1.0581 0.01161 0.00343 -0.1436 0.3283 0.0348
3.750 1.0842 0.01177 0.00357 -0.1433 0.3186 0.0377
4.000 1.1105 0.01192 0.00372 -0.1431 0.3114 0.0392
4.250 1.1366 0.01209 0.00388 -0.1428 0.3044 0.0407
4.750 1.1851 0.01122 0.00448 -0.1419 0.2841 1.0000
5.000 1.2106 0.01148 0.00469 -0.1415 0.2746 1.0000
5.250 1.2363 0.01169 0.00489 -0.1412 0.2656 1.0000
5.500 1.2617 0.01194 0.00513 -0.1408 0.2556 1.0000
5.750 1.2866 0.01223 0.00539 -0.1403 0.2419 1.0000
6.000 1.3103 0.01263 0.00568 -0.1397 0.2163 1.0000
6.250 1.3270 0.01381 0.00638 -0.1381 0.1365 1.0000
6.500 1.3477 0.01454 0.00696 -0.1371 0.1124 1.0000
6.750 1.3698 0.01508 0.00746 -0.1362 0.0997 1.0000
7.000 1.3927 0.01552 0.00788 -0.1355 0.0894 1.0000
7.250 1.4150 0.01599 0.00833 -0.1347 0.0784 1.0000
7.500 1.4367 0.01652 0.00881 -0.1338 0.0676 1.0000
7.750 1.4578 0.01708 0.00933 -0.1328 0.0578 1.0000
8.000 1.4789 0.01762 0.00986 -0.1318 0.0512 1.0000
8.250 1.5000 0.01812 0.01039 -0.1308 0.0449 1.0000
8.500 1.5192 0.01879 0.01101 -0.1295 0.0340 1.0000
8.750 1.5347 0.01978 0.01186 -0.1278 0.0188 1.0000
9.000 1.5520 0.02054 0.01266 -0.1262 0.0149 1.0000
9.250 1.5683 0.02135 0.01352 -0.1244 0.0128 1.0000
9.500 1.5844 0.02211 0.01436 -0.1226 0.0115 1.0000
9.750 1.5997 0.02283 0.01517 -0.1207 0.0105 1.0000
10.000 1.6119 0.02363 0.01605 -0.1183 0.0097 1.0000
10.250 1.6224 0.02457 0.01708 -0.1157 0.0090 1.0000
10.500 1.6301 0.02575 0.01836 -0.1128 0.0083 1.0000
10.750 1.6409 0.02671 0.01942 -0.1105 0.0080 1.0000
11.000 1.6504 0.02779 0.02061 -0.1081 0.0077 1.0000
11.250 1.6589 0.02898 0.02191 -0.1058 0.0074 1.0000
11.500 1.6665 0.03027 0.02330 -0.1035 0.0070 1.0000
11.750 1.6735 0.03167 0.02480 -0.1012 0.0067 1.0000
12.000 1.6797 0.03317 0.02640 -0.0991 0.0064 1.0000
12.250 1.6837 0.03491 0.02825 -0.0969 0.0061 1.0000
12.500 1.6835 0.03711 0.03056 -0.0946 0.0059 1.0000
12.750 1.6849 0.03926 0.03283 -0.0926 0.0057 1.0000
13.000 1.6877 0.04133 0.03504 -0.0909 0.0055 1.0000
13.250 1.6884 0.04369 0.03753 -0.0893 0.0054 1.0000
13.500 1.6878 0.04627 0.04025 -0.0878 0.0053 1.0000
13.750 1.6861 0.04906 0.04318 -0.0865 0.0051 1.0000
14.000 1.6829 0.05208 0.04634 -0.0854 0.0050 1.0000
14.250 1.6788 0.05533 0.04974 -0.0844 0.0049 1.0000
14.500 1.6734 0.05886 0.05342 -0.0837 0.0048 1.0000
14.750 1.6670 0.06264 0.05734 -0.0833 0.0047 1.0000
15.000 1.6599 0.06671 0.06155 -0.0831 0.0046 1.0000
15.250 1.6515 0.07117 0.06616 -0.0834 0.0046 1.0000
15.500 1.6425 0.07596 0.07110 -0.0840 0.0045 1.0000
15.750 1.6326 0.08109 0.07637 -0.0850 0.0044 1.0000
16.000 1.6218 0.08660 0.08203 -0.0865 0.0044 1.0000
16.250 1.6101 0.09248 0.08806 -0.0883 0.0043 1.0000
16.500 1.5977 0.09865 0.09438 -0.0904 0.0043 1.0000
16.750 1.5848 0.10501 0.10088 -0.0927 0.0043 1.0000
17.000 1.5717 0.11152 0.10753 -0.0953 0.0042 1.0000
17.250 1.5585 0.11815 0.11430 -0.0981 0.0042 1.0000
17.500 1.5455 0.12481 0.12109 -0.1010 0.0042 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 342 AIRFOIL (goe342-il)