Ornithopter airfoil. (s1020-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: Ornithopter airfoil. (s1020-il) Reynolds number: 200,000 Max Cl/Cd: 82.35 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s1020-il-200000.txt Download as CSV file: xf-s1020-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1807 0.09612 0.09271 -0.0889 0.9578 0.0641
-9.000 -0.2059 0.08537 0.08202 -0.1080 0.9498 0.0680
-8.750 -0.1996 0.08201 0.07869 -0.1049 0.9433 0.0692
-8.500 -0.1743 0.08033 0.07698 -0.1042 0.9408 0.0705
-8.250 -0.1541 0.07719 0.07383 -0.1070 0.9383 0.0725
-8.000 -0.1537 0.07368 0.07034 -0.1090 0.9300 0.0749
-7.000 -0.1986 0.03626 0.03098 -0.1372 0.8946 0.0423
-6.750 -0.1938 0.03300 0.02677 -0.1342 0.8845 0.0370
-6.500 -0.1609 0.03022 0.02361 -0.1362 0.8824 0.0370
-6.250 -0.1252 0.02721 0.02027 -0.1387 0.8811 0.0374
-6.000 -0.0851 0.02507 0.01786 -0.1414 0.8801 0.0374
-5.750 -0.0710 0.02403 0.01670 -0.1390 0.8721 0.0375
-5.500 -0.0359 0.02240 0.01498 -0.1405 0.8695 0.0381
-5.250 0.0036 0.02099 0.01353 -0.1430 0.8677 0.0393
-5.000 0.0447 0.01981 0.01232 -0.1457 0.8662 0.0414
-4.750 0.0866 0.01878 0.01122 -0.1486 0.8647 0.0444
-4.500 0.0996 0.01833 0.01076 -0.1460 0.8559 0.0477
-4.250 0.1371 0.01751 0.00988 -0.1481 0.8529 0.0557
-4.000 0.1762 0.01556 0.00854 -0.1514 0.8506 0.1744
-3.750 0.2036 0.01479 0.00828 -0.1521 0.8455 0.3130
-3.500 0.2296 0.01455 0.00829 -0.1520 0.8394 0.4050
-3.250 0.2676 0.01434 0.00812 -0.1539 0.8362 0.4632
-3.000 0.3079 0.01419 0.00793 -0.1562 0.8336 0.5052
-2.750 0.3232 0.01441 0.00817 -0.1536 0.8250 0.5299
-2.500 0.3586 0.01430 0.00801 -0.1549 0.8210 0.5544
-2.250 0.3971 0.01417 0.00781 -0.1568 0.8179 0.5747
-2.000 0.4135 0.01432 0.00797 -0.1545 0.8095 0.5897
-1.750 0.4482 0.01420 0.00779 -0.1557 0.8053 0.6067
-1.500 0.4764 0.01420 0.00776 -0.1557 0.7997 0.6217
-1.250 0.5016 0.01422 0.00775 -0.1551 0.7931 0.6359
-1.000 0.5368 0.01409 0.00757 -0.1565 0.7889 0.6508
-0.750 0.5591 0.01418 0.00765 -0.1554 0.7819 0.6645
-0.500 0.5881 0.01412 0.00757 -0.1556 0.7762 0.6779
-0.250 0.6244 0.01400 0.00740 -0.1571 0.7723 0.6917
0.000 0.6414 0.01412 0.00758 -0.1550 0.7641 0.7046
0.250 0.6733 0.01403 0.00746 -0.1557 0.7590 0.7191
0.500 0.6978 0.01408 0.00752 -0.1551 0.7526 0.7338
0.750 0.7232 0.01407 0.00754 -0.1545 0.7462 0.7493
1.000 0.7573 0.01398 0.00741 -0.1556 0.7418 0.7659
1.250 0.7738 0.01408 0.00761 -0.1534 0.7336 0.7831
1.500 0.8032 0.01401 0.00753 -0.1536 0.7283 0.8021
1.750 0.8247 0.01406 0.00763 -0.1522 0.7216 0.8230
2.000 0.8464 0.01401 0.00765 -0.1508 0.7150 0.8451
2.250 0.8731 0.01394 0.00757 -0.1503 0.7099 0.8716
2.500 0.8848 0.01392 0.00768 -0.1468 0.7020 0.9074
3.000 0.9533 0.01391 0.00765 -0.1499 0.6888 1.0000
3.250 0.9869 0.01397 0.00763 -0.1515 0.6827 1.0000
3.500 1.0128 0.01416 0.00779 -0.1514 0.6754 1.0000
3.750 1.0414 0.01427 0.00787 -0.1518 0.6685 1.0000
4.000 1.0693 0.01443 0.00800 -0.1521 0.6619 1.0000
4.250 1.0940 0.01459 0.00815 -0.1516 0.6543 1.0000
4.500 1.1238 0.01472 0.00825 -0.1522 0.6478 1.0000
4.750 1.1454 0.01491 0.00847 -0.1512 0.6397 1.0000
5.000 1.1757 0.01504 0.00855 -0.1518 0.6331 1.0000
5.250 1.1956 0.01525 0.00882 -0.1504 0.6245 1.0000
5.500 1.2243 0.01539 0.00893 -0.1507 0.6174 1.0000
5.750 1.2445 0.01559 0.00918 -0.1493 0.6087 1.0000
6.000 1.2703 0.01578 0.00935 -0.1491 0.6011 1.0000
6.250 1.2919 0.01597 0.00959 -0.1480 0.5924 1.0000
6.500 1.3143 0.01619 0.00984 -0.1471 0.5842 1.0000
6.750 1.3381 0.01637 0.01002 -0.1464 0.5756 1.0000
7.000 1.3561 0.01663 0.01035 -0.1447 0.5664 1.0000
7.250 1.3835 0.01680 0.01048 -0.1447 0.5579 1.0000
7.500 1.3959 0.01707 0.01085 -0.1419 0.5477 1.0000
7.750 1.4149 0.01732 0.01113 -0.1403 0.5381 1.0000
8.000 1.4359 0.01753 0.01135 -0.1391 0.5283 1.0000
8.250 1.4441 0.01783 0.01174 -0.1355 0.5176 1.0000
8.500 1.4580 0.01812 0.01207 -0.1330 0.5069 1.0000
8.750 1.4732 0.01842 0.01235 -0.1308 0.4957 1.0000
9.000 1.4813 0.01878 0.01276 -0.1273 0.4832 1.0000
9.250 1.4870 0.01921 0.01325 -0.1234 0.4696 1.0000
9.500 1.4917 0.01970 0.01375 -0.1196 0.4543 1.0000
9.750 1.4956 0.02028 0.01434 -0.1157 0.4381 1.0000
10.000 1.4976 0.02100 0.01503 -0.1117 0.4203 1.0000
10.250 1.4991 0.02185 0.01590 -0.1079 0.4024 1.0000
10.500 1.5004 0.02283 0.01689 -0.1042 0.3835 1.0000
10.750 1.5015 0.02395 0.01799 -0.1008 0.3645 1.0000
11.000 1.5012 0.02525 0.01925 -0.0973 0.3449 1.0000
11.250 1.5004 0.02670 0.02068 -0.0941 0.3234 1.0000
11.500 1.4962 0.02846 0.02236 -0.0907 0.3011 1.0000
11.750 1.4925 0.03035 0.02420 -0.0876 0.2777 1.0000
12.000 1.4862 0.03257 0.02632 -0.0846 0.2549 1.0000
12.250 1.4820 0.03481 0.02850 -0.0821 0.2328 1.0000
12.500 1.4758 0.03734 0.03095 -0.0796 0.2125 1.0000
12.750 1.4700 0.04000 0.03355 -0.0775 0.1896 1.0000
13.000 1.4629 0.04291 0.03638 -0.0755 0.1697 1.0000
13.250 1.4571 0.04587 0.03929 -0.0739 0.1483 1.0000
13.750 1.4400 0.05277 0.04602 -0.0711 0.1063 1.0000
14.000 1.4291 0.05672 0.04987 -0.0700 0.0845 1.0000
14.250 1.4132 0.06141 0.05440 -0.0690 0.0614 1.0000
14.500 1.3961 0.06649 0.05936 -0.0683 0.0459 1.0000
14.750 1.3821 0.07143 0.06428 -0.0680 0.0378 1.0000
15.000 1.3748 0.07567 0.06861 -0.0679 0.0331 1.0000
15.250 1.3647 0.08038 0.07338 -0.0681 0.0309 1.0000
15.500 1.3601 0.08445 0.07758 -0.0683 0.0290 1.0000
15.750 1.3571 0.08835 0.08159 -0.0687 0.0275 1.0000
16.000 1.3535 0.09240 0.08572 -0.0693 0.0260 1.0000
16.250 1.3484 0.09664 0.09001 -0.0699 0.0250 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Ornithopter airfoil. (s1020-il)