NACA 63-415 AIRFOIL (n63415-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63-415 AIRFOIL (n63415-il) Reynolds number: 100,000 Max Cl/Cd: 53.05 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63415-il-100000.txt Download as CSV file: xf-n63415-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-415 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.3312 0.10643 0.10167 -0.0436 1.0000 0.1704
-11.000 -0.2956 0.10409 0.09929 -0.0408 1.0000 0.1765
-10.750 -0.5457 0.07393 0.06883 -0.0666 1.0000 0.0933
-10.500 -0.5648 0.06828 0.06318 -0.0685 1.0000 0.0906
-10.250 -0.6742 0.06179 0.05637 -0.0669 1.0000 0.0812
-10.000 -0.6912 0.05912 0.05365 -0.0638 1.0000 0.0807
-9.750 -0.7131 0.05710 0.05159 -0.0594 1.0000 0.0803
-9.500 -0.7321 0.05506 0.04945 -0.0552 1.0000 0.0796
-9.250 -0.7496 0.05286 0.04708 -0.0513 1.0000 0.0794
-9.000 -0.7612 0.05035 0.04436 -0.0482 1.0000 0.0790
-8.750 -0.7678 0.04772 0.04148 -0.0455 1.0000 0.0789
-8.500 -0.7694 0.04498 0.03843 -0.0432 1.0000 0.0786
-8.250 -0.7645 0.04213 0.03521 -0.0418 0.9995 0.0784
-8.000 -0.7323 0.03855 0.03105 -0.0449 0.9935 0.0794
-7.750 -0.6995 0.03625 0.02810 -0.0474 0.9866 0.0820
-7.500 -0.6643 0.03326 0.02493 -0.0501 0.9817 0.0849
-7.250 -0.6311 0.03156 0.02311 -0.0519 0.9749 0.0885
-7.000 -0.5900 0.03014 0.02136 -0.0549 0.9697 0.0941
-6.750 -0.5600 0.02843 0.01964 -0.0558 0.9626 0.0983
-6.500 -0.5210 0.02731 0.01847 -0.0583 0.9569 0.1061
-6.250 -0.4885 0.02607 0.01729 -0.0595 0.9504 0.1133
-6.000 -0.4544 0.02516 0.01635 -0.0610 0.9435 0.1243
-5.750 -0.4142 0.02389 0.01533 -0.0638 0.9392 0.1417
-5.500 -0.3933 0.02292 0.01456 -0.0634 0.9296 0.1669
-5.250 -0.3629 0.02010 0.01323 -0.0665 0.9243 0.3527
-5.000 -0.3397 0.02022 0.01376 -0.0656 0.9153 0.4909
-4.750 -0.3015 0.02080 0.01430 -0.0667 0.9093 0.5416
-4.500 -0.2717 0.02141 0.01482 -0.0665 0.9019 0.5728
-4.250 -0.2419 0.02218 0.01562 -0.0655 0.8947 0.5968
-4.000 -0.2013 0.02291 0.01629 -0.0663 0.8907 0.6241
-3.750 -0.1862 0.02361 0.01698 -0.0631 0.8804 0.6412
-3.500 -0.1516 0.02423 0.01756 -0.0626 0.8756 0.6590
-3.250 -0.1319 0.02463 0.01789 -0.0608 0.8668 0.6740
-3.000 -0.1035 0.02507 0.01830 -0.0594 0.8606 0.6866
-2.750 -0.0675 0.02532 0.01848 -0.0594 0.8569 0.6993
-2.500 -0.0561 0.02551 0.01862 -0.0570 0.8460 0.7106
-2.250 -0.0201 0.02542 0.01843 -0.0578 0.8416 0.7212
-2.000 -0.0059 0.02564 0.01862 -0.0556 0.8322 0.7295
-1.750 0.0269 0.02544 0.01831 -0.0569 0.8266 0.7409
-1.500 0.0610 0.02536 0.01817 -0.0568 0.8231 0.7477
-1.250 0.0723 0.02556 0.01834 -0.0552 0.8120 0.7582
-1.000 0.1040 0.02546 0.01819 -0.0550 0.8079 0.7652
-0.750 0.1190 0.02569 0.01838 -0.0539 0.7984 0.7753
-0.500 0.1465 0.02562 0.01829 -0.0533 0.7929 0.7825
-0.250 0.1836 0.02533 0.01792 -0.0548 0.7897 0.7919
0.000 0.1886 0.02587 0.01850 -0.0516 0.7786 0.7998
0.250 0.2233 0.02562 0.01820 -0.0527 0.7746 0.8090
0.500 0.2334 0.02609 0.01870 -0.0503 0.7653 0.8176
0.750 0.2613 0.02602 0.01861 -0.0503 0.7598 0.8266
1.000 0.2954 0.02575 0.01831 -0.0507 0.7567 0.8354
1.250 0.2990 0.02646 0.01907 -0.0479 0.7456 0.8454
1.500 0.3326 0.02624 0.01884 -0.0487 0.7416 0.8556
1.750 0.3393 0.02683 0.01948 -0.0458 0.7324 0.8657
2.000 0.3643 0.02678 0.01944 -0.0452 0.7267 0.8761
2.250 0.3999 0.02642 0.01908 -0.0458 0.7237 0.8868
2.500 0.3981 0.02734 0.02008 -0.0423 0.7122 0.9003
2.750 0.4310 0.02698 0.01974 -0.0424 0.7085 0.9115
3.000 0.4371 0.02777 0.02059 -0.0401 0.6984 0.9274
3.250 0.4716 0.02753 0.02041 -0.0409 0.6936 0.9405
3.500 0.5210 0.02693 0.01984 -0.0436 0.6909 0.9514
3.750 0.5458 0.02800 0.02103 -0.0455 0.6795 0.9681
4.000 0.6066 0.02737 0.02046 -0.0505 0.6761 0.9763
4.250 0.6526 0.02790 0.02112 -0.0552 0.6671 0.9888
4.500 0.7102 0.02744 0.02074 -0.0602 0.6613 0.9994
4.750 0.7165 0.02759 0.02091 -0.0573 0.6542 1.0000
5.000 0.7311 0.02788 0.02126 -0.0562 0.6449 1.0000
5.250 0.7689 0.02735 0.02078 -0.0573 0.6384 1.0000
5.500 0.7990 0.02719 0.02068 -0.0578 0.6285 1.0000
5.750 0.8299 0.02703 0.02060 -0.0583 0.6183 1.0000
6.000 0.8765 0.02576 0.01937 -0.0597 0.6106 1.0000
6.250 0.9052 0.02559 0.01928 -0.0598 0.5982 1.0000
6.500 0.9421 0.02475 0.01850 -0.0603 0.5862 1.0000
6.750 0.9804 0.02366 0.01743 -0.0607 0.5724 1.0000
7.000 1.0121 0.02274 0.01655 -0.0602 0.5539 1.0000
7.250 1.0424 0.02189 0.01570 -0.0597 0.5332 1.0000
7.500 1.0726 0.02119 0.01496 -0.0591 0.5120 1.0000
7.750 1.0910 0.02115 0.01506 -0.0575 0.4873 1.0000
8.000 1.1103 0.02106 0.01499 -0.0559 0.4584 1.0000
8.250 1.1247 0.02120 0.01511 -0.0536 0.4214 1.0000
8.500 1.1315 0.02171 0.01543 -0.0504 0.3694 1.0000
8.750 1.1257 0.02293 0.01622 -0.0459 0.3015 1.0000
9.000 1.1108 0.02496 0.01774 -0.0409 0.2405 1.0000
9.250 1.0961 0.02750 0.01981 -0.0369 0.1935 1.0000
9.500 1.0867 0.03008 0.02204 -0.0339 0.1610 1.0000
9.750 1.0839 0.03246 0.02412 -0.0317 0.1395 1.0000
10.000 1.0893 0.03446 0.02598 -0.0300 0.1234 1.0000
10.250 1.1000 0.03626 0.02760 -0.0287 0.1117 1.0000
10.500 1.1139 0.03784 0.02913 -0.0276 0.1021 1.0000
10.750 1.1339 0.03939 0.03070 -0.0268 0.0941 1.0000
11.000 1.1619 0.04098 0.03207 -0.0266 0.0866 1.0000
11.250 1.1791 0.04263 0.03392 -0.0259 0.0815 1.0000
11.500 1.2185 0.04460 0.03566 -0.0267 0.0752 1.0000
11.750 1.2341 0.04660 0.03799 -0.0258 0.0723 1.0000
12.000 1.2532 0.04872 0.04030 -0.0253 0.0691 1.0000
12.250 1.2792 0.05098 0.04251 -0.0256 0.0657 1.0000
12.500 1.2948 0.05421 0.04599 -0.0251 0.0636 1.0000
12.750 1.2954 0.05709 0.04922 -0.0235 0.0627 1.0000
13.000 1.2932 0.06031 0.05280 -0.0221 0.0620 1.0000
13.250 1.2871 0.06383 0.05665 -0.0207 0.0615 1.0000
13.500 1.2769 0.06765 0.06079 -0.0197 0.0612 1.0000
13.750 1.2634 0.07183 0.06527 -0.0190 0.0610 1.0000
14.000 1.2459 0.07644 0.07018 -0.0188 0.0609 1.0000
14.250 1.2250 0.08161 0.07564 -0.0193 0.0610 1.0000
14.500 1.2011 0.08737 0.08167 -0.0205 0.0613 1.0000
14.750 1.1739 0.09391 0.08847 -0.0226 0.0617 1.0000
15.000 1.1449 0.10123 0.09603 -0.0256 0.0622 1.0000
15.250 1.1156 0.10931 0.10430 -0.0296 0.0629 1.0000
15.500 1.0873 0.11816 0.11331 -0.0345 0.0636 1.0000
15.750 1.0655 0.12698 0.12224 -0.0392 0.0642 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63-415 AIRFOIL (n63415-il)