FX 63-110 AIRFOIL (fx63110-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-110 AIRFOIL (fx63110-il) Reynolds number: 1,000,000 Max Cl/Cd: 117.17 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx63110-il-1000000-n5.txt Download as CSV file: xf-fx63110-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-110 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.9134 0.04576 0.04319 -0.0819 1.0000 0.0071
-13.750 -0.9653 0.03754 0.03478 -0.0857 1.0000 0.0069
-13.500 -0.9493 0.03034 0.02732 -0.1001 0.9969 0.0070
-13.250 -0.9183 0.02649 0.02325 -0.1098 0.9947 0.0073
-13.000 -0.8917 0.02441 0.02102 -0.1137 0.9925 0.0074
-12.750 -0.8628 0.02279 0.01926 -0.1169 0.9903 0.0076
-12.500 -0.8322 0.02140 0.01775 -0.1198 0.9884 0.0078
-12.250 -0.8001 0.02020 0.01643 -0.1225 0.9870 0.0081
-12.000 -0.7671 0.01909 0.01520 -0.1251 0.9858 0.0083
-11.750 -0.7409 0.01819 0.01419 -0.1260 0.9821 0.0086
-11.500 -0.7103 0.01737 0.01326 -0.1275 0.9793 0.0088
-11.250 -0.6781 0.01639 0.01219 -0.1296 0.9772 0.0092
-11.000 -0.6490 0.01562 0.01136 -0.1307 0.9735 0.0097
-10.750 -0.6208 0.01495 0.01063 -0.1314 0.9681 0.0101
-10.500 -0.5901 0.01433 0.00994 -0.1326 0.9636 0.0106
-10.250 -0.5609 0.01377 0.00930 -0.1333 0.9563 0.0111
-10.000 -0.5262 0.01321 0.00866 -0.1351 0.9510 0.0115
-9.750 -0.4885 0.01261 0.00801 -0.1377 0.9443 0.0123
-9.500 -0.4465 0.01211 0.00746 -0.1410 0.9379 0.0133
-9.250 -0.4050 0.01165 0.00693 -0.1442 0.9279 0.0142
-9.000 -0.3666 0.01126 0.00644 -0.1466 0.9140 0.0150
-8.750 -0.3340 0.01099 0.00610 -0.1478 0.8984 0.0163
-8.500 -0.3046 0.01080 0.00583 -0.1482 0.8845 0.0174
-8.250 -0.2760 0.01059 0.00553 -0.1484 0.8728 0.0183
-7.750 -0.2194 0.01013 0.00493 -0.1488 0.8526 0.0200
-7.500 -0.1912 0.00997 0.00473 -0.1489 0.8433 0.0210
-7.250 -0.1630 0.00982 0.00454 -0.1490 0.8341 0.0221
-7.000 -0.1346 0.00967 0.00433 -0.1491 0.8264 0.0230
-6.750 -0.1061 0.00951 0.00410 -0.1492 0.8182 0.0236
-6.500 -0.0777 0.00939 0.00392 -0.1493 0.8104 0.0240
-6.250 -0.0486 0.00906 0.00355 -0.1497 0.8023 0.0256
-6.000 -0.0199 0.00889 0.00333 -0.1499 0.7952 0.0266
-5.750 0.0090 0.00874 0.00315 -0.1501 0.7876 0.0277
-5.500 0.0376 0.00861 0.00296 -0.1502 0.7794 0.0287
-5.250 0.0663 0.00849 0.00278 -0.1503 0.7693 0.0296
-5.000 0.0948 0.00838 0.00261 -0.1504 0.7583 0.0303
-4.750 0.1233 0.00830 0.00245 -0.1505 0.7469 0.0307
-4.500 0.1523 0.00815 0.00223 -0.1506 0.7365 0.0318
-4.250 0.1815 0.00800 0.00204 -0.1509 0.7285 0.0339
-4.000 0.2105 0.00790 0.00191 -0.1511 0.7205 0.0358
-3.750 0.2396 0.00780 0.00178 -0.1513 0.7131 0.0376
-3.500 0.2685 0.00773 0.00168 -0.1514 0.7052 0.0395
-3.250 0.2976 0.00766 0.00158 -0.1515 0.6978 0.0436
-3.000 0.3267 0.00758 0.00149 -0.1517 0.6902 0.0515
-2.750 0.3558 0.00749 0.00143 -0.1520 0.6830 0.0653
-2.500 0.3849 0.00741 0.00137 -0.1522 0.6750 0.0797
-2.250 0.4139 0.00736 0.00132 -0.1523 0.6679 0.0920
-2.000 0.4430 0.00731 0.00128 -0.1525 0.6604 0.1028
-1.750 0.4719 0.00727 0.00124 -0.1527 0.6530 0.1157
-1.500 0.5014 0.00716 0.00120 -0.1530 0.6450 0.1464
-1.250 0.5324 0.00680 0.00115 -0.1540 0.6378 0.2693
-1.000 0.5620 0.00668 0.00117 -0.1544 0.6299 0.3292
-0.750 0.5910 0.00664 0.00119 -0.1546 0.6219 0.3658
-0.500 0.6200 0.00663 0.00121 -0.1548 0.6136 0.3895
-0.250 0.6486 0.00665 0.00124 -0.1549 0.6054 0.4050
0.250 0.7057 0.00670 0.00131 -0.1550 0.5818 0.4448
0.500 0.7335 0.00678 0.00137 -0.1550 0.5618 0.4706
0.750 0.7614 0.00686 0.00146 -0.1550 0.5392 0.5111
1.000 0.7889 0.00699 0.00156 -0.1548 0.5171 0.5410
1.250 0.8165 0.00713 0.00165 -0.1547 0.4998 0.5598
1.500 0.8438 0.00729 0.00176 -0.1546 0.4800 0.5745
1.750 0.8704 0.00752 0.00189 -0.1543 0.4546 0.5833
2.000 0.8974 0.00772 0.00201 -0.1541 0.4360 0.5903
2.250 0.9243 0.00792 0.00214 -0.1538 0.4159 0.5974
2.500 0.9509 0.00816 0.00230 -0.1536 0.3940 0.6034
2.750 0.9779 0.00835 0.00244 -0.1534 0.3773 0.6105
3.250 1.0311 0.00880 0.00276 -0.1529 0.3384 0.6232
3.500 1.0575 0.00905 0.00293 -0.1526 0.3184 0.6292
3.750 1.0831 0.00937 0.00314 -0.1522 0.2918 0.6347
4.000 1.1065 0.00993 0.00348 -0.1515 0.2449 0.6408
4.250 1.1303 0.01045 0.00381 -0.1509 0.2086 0.6476
4.500 1.1554 0.01080 0.00409 -0.1504 0.1869 0.6562
5.000 1.2064 0.01142 0.00462 -0.1497 0.1564 0.6747
5.250 1.2314 0.01177 0.00491 -0.1492 0.1390 0.6843
5.500 1.2554 0.01222 0.00526 -0.1486 0.1170 0.6956
5.750 1.2789 0.01272 0.00566 -0.1479 0.0944 0.7074
6.000 1.3031 0.01311 0.00602 -0.1474 0.0820 0.7201
6.250 1.3281 0.01342 0.00635 -0.1469 0.0753 0.7357
6.500 1.3527 0.01375 0.00670 -0.1464 0.0689 0.7540
6.750 1.3773 0.01406 0.00705 -0.1459 0.0630 0.7742
7.000 1.4012 0.01442 0.00743 -0.1453 0.0560 0.7946
7.250 1.4244 0.01482 0.00784 -0.1445 0.0484 0.8158
7.500 1.4474 0.01516 0.00824 -0.1438 0.0431 0.8447
7.750 1.4627 0.01533 0.00857 -0.1412 0.0392 1.0000
8.000 1.4861 0.01571 0.00897 -0.1406 0.0374 1.0000
8.250 1.5090 0.01613 0.00938 -0.1398 0.0358 1.0000
8.500 1.5310 0.01660 0.00986 -0.1389 0.0338 1.0000
8.750 1.5528 0.01707 0.01034 -0.1380 0.0318 1.0000
9.000 1.5748 0.01748 0.01078 -0.1372 0.0307 1.0000
9.250 1.5960 0.01794 0.01126 -0.1362 0.0297 1.0000
9.500 1.6163 0.01844 0.01178 -0.1350 0.0286 1.0000
9.750 1.6354 0.01899 0.01235 -0.1337 0.0275 1.0000
10.000 1.6533 0.01958 0.01298 -0.1322 0.0263 1.0000
10.250 1.6704 0.02008 0.01351 -0.1305 0.0258 1.0000
10.500 1.6865 0.02056 0.01405 -0.1287 0.0252 1.0000
10.750 1.7016 0.02110 0.01463 -0.1267 0.0244 1.0000
11.000 1.7159 0.02171 0.01528 -0.1247 0.0235 1.0000
11.250 1.7292 0.02240 0.01601 -0.1227 0.0225 1.0000
11.500 1.7413 0.02319 0.01683 -0.1205 0.0215 1.0000
11.750 1.7538 0.02398 0.01767 -0.1185 0.0206 1.0000
12.000 1.7665 0.02478 0.01852 -0.1166 0.0196 1.0000
12.250 1.7779 0.02570 0.01948 -0.1147 0.0181 1.0000
12.500 1.7877 0.02678 0.02057 -0.1127 0.0167 1.0000
12.750 1.7976 0.02789 0.02174 -0.1109 0.0155 1.0000
13.000 1.8061 0.02917 0.02306 -0.1091 0.0140 1.0000
13.250 1.8134 0.03060 0.02452 -0.1073 0.0127 1.0000
13.500 1.8205 0.03211 0.02610 -0.1056 0.0116 1.0000
13.750 1.8259 0.03385 0.02789 -0.1040 0.0107 1.0000
14.000 1.8311 0.03567 0.02977 -0.1025 0.0099 1.0000
14.250 1.8362 0.03756 0.03174 -0.1012 0.0093 1.0000
14.500 1.8400 0.03967 0.03392 -0.0999 0.0087 1.0000
14.750 1.8422 0.04201 0.03633 -0.0988 0.0081 1.0000
15.000 1.8440 0.04449 0.03890 -0.0978 0.0077 1.0000
15.250 1.8466 0.04697 0.04146 -0.0970 0.0074 1.0000
15.500 1.8481 0.04966 0.04424 -0.0964 0.0071 1.0000
15.750 1.8484 0.05259 0.04727 -0.0959 0.0068 1.0000
16.000 1.8478 0.05574 0.05051 -0.0956 0.0065 1.0000
16.250 1.8459 0.05918 0.05404 -0.0956 0.0063 1.0000
16.500 1.8429 0.06290 0.05786 -0.0957 0.0061 1.0000
16.750 1.8384 0.06695 0.06202 -0.0961 0.0059 1.0000
17.000 1.8338 0.07116 0.06634 -0.0967 0.0057 1.0000
17.250 1.8302 0.07536 0.07067 -0.0975 0.0056 1.0000
17.500 1.8248 0.07994 0.07537 -0.0985 0.0055 1.0000
17.750 1.8183 0.08483 0.08038 -0.0999 0.0054 1.0000
18.000 1.8111 0.08998 0.08564 -0.1014 0.0053 1.0000
18.250 1.8025 0.09552 0.09131 -0.1033 0.0052 1.0000
18.500 1.7928 0.10134 0.09726 -0.1055 0.0051 1.0000
18.750 1.7818 0.10751 0.10356 -0.1080 0.0050 1.0000
19.000 1.7706 0.11384 0.11001 -0.1108 0.0049 1.0000
19.250 1.7582 0.12049 0.11679 -0.1139 0.0048 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 63-110 AIRFOIL (fx63110-il)