Ornithopter airfoil. (s1020-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: Ornithopter airfoil. (s1020-il) Reynolds number: 500,000 Max Cl/Cd: 120 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s1020-il-500000.txt Download as CSV file: xf-s1020-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: Ornithopter 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.0460 0.09623 0.09394 -0.1038 0.9653 0.0271
-11.250 -0.0422 0.09119 0.08890 -0.1070 0.9640 0.0286
-10.000 -0.1442 0.07746 0.07513 -0.1173 0.9634 0.0306
-9.750 -0.1278 0.07419 0.07186 -0.1211 0.9620 0.0311
-9.500 -0.1221 0.07149 0.06916 -0.1221 0.9536 0.0316
-9.250 -0.1114 0.06640 0.06407 -0.1278 0.9507 0.0325
-9.000 -0.1400 0.05114 0.04872 -0.1447 0.9392 0.0327
-8.750 -0.1071 0.02821 0.02520 -0.1602 0.9128 0.0367
-8.500 -0.1276 0.02101 0.01758 -0.1623 0.9003 0.0377
-8.250 -0.1053 0.01828 0.01483 -0.1641 0.8948 0.0386
-8.000 -0.1455 0.02111 0.01585 -0.1655 0.8965 0.0196
-7.750 -0.1191 0.01988 0.01440 -0.1660 0.8901 0.0194
-7.500 -0.0942 0.01860 0.01295 -0.1660 0.8830 0.0194
-7.250 -0.0661 0.01722 0.01139 -0.1668 0.8774 0.0195
-7.000 -0.0451 0.01608 0.01015 -0.1659 0.8697 0.0197
-6.750 -0.0166 0.01506 0.00900 -0.1665 0.8642 0.0198
-6.500 0.0041 0.01433 0.00823 -0.1655 0.8568 0.0202
-6.250 0.0304 0.01369 0.00752 -0.1656 0.8508 0.0206
-6.000 0.0549 0.01323 0.00701 -0.1653 0.8445 0.0214
-5.750 0.0795 0.01276 0.00648 -0.1649 0.8380 0.0222
-5.500 0.1072 0.01229 0.00591 -0.1652 0.8326 0.0229
-5.250 0.1303 0.01189 0.00546 -0.1644 0.8259 0.0235
-5.000 0.1574 0.01149 0.00497 -0.1645 0.8203 0.0243
-4.750 0.1829 0.01107 0.00448 -0.1643 0.8145 0.0260
-4.500 0.2086 0.01080 0.00417 -0.1640 0.8083 0.0280
-4.250 0.2378 0.01053 0.00382 -0.1644 0.8032 0.0328
-4.000 0.2612 0.00996 0.00349 -0.1640 0.7970 0.0836
-3.750 0.2866 0.00934 0.00322 -0.1641 0.7912 0.1856
-3.500 0.3138 0.00891 0.00305 -0.1645 0.7860 0.2754
-3.250 0.3388 0.00863 0.00297 -0.1642 0.7798 0.3435
-3.000 0.3666 0.00849 0.00292 -0.1644 0.7744 0.3967
-2.750 0.3937 0.00844 0.00291 -0.1643 0.7688 0.4355
-2.500 0.4200 0.00840 0.00290 -0.1641 0.7627 0.4621
-2.250 0.4490 0.00842 0.00287 -0.1644 0.7575 0.4859
-2.000 0.4746 0.00841 0.00290 -0.1639 0.7514 0.5079
-1.750 0.5016 0.00841 0.00288 -0.1638 0.7456 0.5230
-1.500 0.5300 0.00845 0.00286 -0.1639 0.7402 0.5365
-1.250 0.5555 0.00845 0.00287 -0.1635 0.7339 0.5492
-1.000 0.5831 0.00846 0.00286 -0.1635 0.7282 0.5609
-0.750 0.6096 0.00848 0.00287 -0.1633 0.7222 0.5727
-0.500 0.6359 0.00849 0.00288 -0.1630 0.7160 0.5848
-0.250 0.6641 0.00854 0.00288 -0.1631 0.7105 0.5968
0.000 0.6888 0.00853 0.00292 -0.1625 0.7038 0.6079
0.250 0.7158 0.00857 0.00293 -0.1624 0.6978 0.6198
0.500 0.7420 0.00861 0.00298 -0.1621 0.6916 0.6324
0.750 0.7675 0.00863 0.00302 -0.1617 0.6849 0.6454
1.000 0.7944 0.00869 0.00307 -0.1616 0.6790 0.6588
1.250 0.8189 0.00871 0.00314 -0.1610 0.6721 0.6727
1.500 0.8455 0.00877 0.00318 -0.1608 0.6658 0.6876
1.750 0.8698 0.00880 0.00328 -0.1601 0.6589 0.7032
2.000 0.8951 0.00886 0.00334 -0.1597 0.6521 0.7197
2.250 0.9197 0.00891 0.00344 -0.1591 0.6454 0.7372
2.500 0.9439 0.00896 0.00353 -0.1584 0.6383 0.7560
2.750 0.9684 0.00903 0.00363 -0.1578 0.6317 0.7759
3.000 0.9911 0.00907 0.00374 -0.1568 0.6243 0.7972
3.250 1.0146 0.00915 0.00385 -0.1559 0.6177 0.8215
3.500 1.0351 0.00917 0.00398 -0.1544 0.6102 0.8493
3.750 1.0551 0.00923 0.00407 -0.1527 0.6034 0.8846
4.000 1.0742 0.00914 0.00414 -0.1508 0.5958 0.9638
4.250 1.1010 0.00930 0.00423 -0.1508 0.5884 1.0000
4.500 1.1235 0.00942 0.00437 -0.1499 0.5803 1.0000
4.750 1.1465 0.00961 0.00450 -0.1491 0.5728 1.0000
5.000 1.1676 0.00973 0.00465 -0.1479 0.5648 1.0000
5.250 1.1896 0.00994 0.00481 -0.1469 0.5572 1.0000
5.500 1.2102 0.01009 0.00499 -0.1456 0.5490 1.0000
5.750 1.2314 0.01031 0.00517 -0.1445 0.5414 1.0000
6.000 1.2516 0.01049 0.00537 -0.1432 0.5328 1.0000
6.250 1.2719 0.01072 0.00559 -0.1419 0.5248 1.0000
6.500 1.2912 0.01094 0.00582 -0.1404 0.5157 1.0000
6.750 1.3108 0.01117 0.00606 -0.1390 0.5070 1.0000
7.000 1.3292 0.01145 0.00631 -0.1374 0.4980 1.0000
7.250 1.3469 0.01170 0.00660 -0.1357 0.4874 1.0000
7.500 1.3641 0.01200 0.00690 -0.1339 0.4769 1.0000
7.750 1.3785 0.01237 0.00723 -0.1316 0.4633 1.0000
8.000 1.3921 0.01277 0.00761 -0.1292 0.4491 1.0000
8.250 1.4058 0.01320 0.00802 -0.1269 0.4345 1.0000
8.500 1.4185 0.01369 0.00848 -0.1245 0.4187 1.0000
8.750 1.4296 0.01425 0.00901 -0.1219 0.3996 1.0000
9.000 1.4397 0.01491 0.00961 -0.1191 0.3805 1.0000
9.250 1.4468 0.01571 0.01033 -0.1161 0.3597 1.0000
9.500 1.4551 0.01653 0.01110 -0.1133 0.3377 1.0000
9.750 1.4596 0.01758 0.01205 -0.1101 0.3142 1.0000
10.000 1.4650 0.01866 0.01304 -0.1072 0.2906 1.0000
10.250 1.4679 0.01993 0.01420 -0.1041 0.2664 1.0000
10.500 1.4694 0.02136 0.01551 -0.1009 0.2419 1.0000
10.750 1.4714 0.02284 0.01689 -0.0981 0.2174 1.0000
11.000 1.4717 0.02450 0.01843 -0.0952 0.1934 1.0000
11.500 1.4710 0.02818 0.02190 -0.0898 0.1468 1.0000
11.750 1.4701 0.03020 0.02381 -0.0874 0.1259 1.0000
12.000 1.4672 0.03246 0.02596 -0.0849 0.1039 1.0000
12.250 1.4646 0.03479 0.02819 -0.0827 0.0836 1.0000
12.500 1.4601 0.03740 0.03069 -0.0806 0.0642 1.0000
12.750 1.4520 0.04047 0.03363 -0.0784 0.0432 1.0000
13.000 1.4434 0.04371 0.03677 -0.0765 0.0266 1.0000
13.250 1.4391 0.04672 0.03975 -0.0750 0.0186 1.0000
13.500 1.4399 0.04931 0.04238 -0.0740 0.0158 1.0000
13.750 1.4418 0.05185 0.04501 -0.0731 0.0144 1.0000
14.000 1.4449 0.05433 0.04757 -0.0724 0.0136 1.0000
14.250 1.4461 0.05709 0.05041 -0.0717 0.0128 1.0000
14.500 1.4451 0.06018 0.05359 -0.0712 0.0122 1.0000
14.750 1.4431 0.06349 0.05699 -0.0707 0.0118 1.0000
15.000 1.4444 0.06646 0.06007 -0.0705 0.0115 1.0000
15.250 1.4445 0.06964 0.06335 -0.0703 0.0113 1.0000
15.500 1.4438 0.07301 0.06682 -0.0703 0.0110 1.0000
15.750 1.4429 0.07647 0.07039 -0.0705 0.0108 1.0000
16.000 1.4407 0.08018 0.07420 -0.0707 0.0105 1.0000
16.250 1.4386 0.08393 0.07805 -0.0711 0.0103 1.0000
16.500 1.4355 0.08789 0.08212 -0.0717 0.0102 1.0000
16.750 1.4323 0.09193 0.08625 -0.0723 0.0099 1.0000
17.000 1.4282 0.09617 0.09058 -0.0732 0.0097 1.0000
17.250 1.4228 0.10063 0.09513 -0.0741 0.0096 1.0000
17.500 1.4156 0.10541 0.10001 -0.0753 0.0094 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Ornithopter airfoil. (s1020-il)