NACA 64-215 AIRFOIL (n64215-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 64-215 AIRFOIL (n64215-il) Reynolds number: 100,000 Max Cl/Cd: 44.25 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64215-il-100000-n5.txt Download as CSV file: xf-n64215-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-215 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.6572 0.09734 0.09177 -0.0384 1.0000 0.0360
-13.500 -0.6911 0.08695 0.08124 -0.0453 1.0000 0.0356
-13.250 -0.7199 0.07895 0.07309 -0.0502 1.0000 0.0353
-13.000 -0.7480 0.07210 0.06605 -0.0541 1.0000 0.0352
-12.750 -0.7713 0.06659 0.06033 -0.0566 1.0000 0.0352
-12.500 -0.7916 0.06187 0.05539 -0.0580 1.0000 0.0352
-12.250 -0.8104 0.05780 0.05107 -0.0586 1.0000 0.0355
-12.000 -0.8263 0.05429 0.04728 -0.0583 1.0000 0.0357
-11.750 -0.8383 0.05122 0.04394 -0.0574 1.0000 0.0360
-11.500 -0.8492 0.04851 0.04093 -0.0556 1.0000 0.0363
-11.250 -0.8539 0.04613 0.03832 -0.0536 1.0000 0.0366
-11.000 -0.8479 0.04398 0.03605 -0.0523 1.0000 0.0370
-10.750 -0.8394 0.04208 0.03403 -0.0510 1.0000 0.0375
-10.500 -0.8294 0.04030 0.03211 -0.0498 1.0000 0.0381
-10.250 -0.8182 0.03861 0.03028 -0.0485 1.0000 0.0389
-10.000 -0.8060 0.03703 0.02856 -0.0472 1.0000 0.0397
-9.750 -0.7939 0.03563 0.02702 -0.0457 1.0000 0.0409
-9.500 -0.7830 0.03434 0.02555 -0.0438 1.0000 0.0423
-9.250 -0.7738 0.03321 0.02422 -0.0414 1.0000 0.0435
-9.000 -0.7648 0.03210 0.02312 -0.0390 1.0000 0.0445
-8.750 -0.7404 0.03081 0.02185 -0.0395 0.9929 0.0459
-8.500 -0.7068 0.02947 0.02044 -0.0416 0.9844 0.0480
-8.250 -0.6732 0.02827 0.01908 -0.0435 0.9756 0.0511
-8.000 -0.6416 0.02702 0.01784 -0.0452 0.9665 0.0539
-7.750 -0.6101 0.02592 0.01673 -0.0469 0.9576 0.0571
-7.500 -0.5797 0.02498 0.01567 -0.0482 0.9478 0.0613
-7.250 -0.5544 0.02398 0.01471 -0.0489 0.9365 0.0657
-7.000 -0.5285 0.02316 0.01380 -0.0494 0.9261 0.0715
-6.750 -0.5055 0.02227 0.01293 -0.0494 0.9154 0.0787
-6.500 -0.4858 0.02152 0.01216 -0.0487 0.9035 0.0884
-6.250 -0.4663 0.02073 0.01140 -0.0479 0.8928 0.1031
-6.000 -0.4476 0.01985 0.01065 -0.0470 0.8829 0.1314
-5.750 -0.4329 0.01884 0.00992 -0.0457 0.8716 0.1854
-5.500 -0.4208 0.01750 0.00916 -0.0442 0.8619 0.2912
-5.250 -0.4061 0.01674 0.00913 -0.0422 0.8524 0.4278
-5.000 -0.3830 0.01667 0.00910 -0.0413 0.8437 0.4910
-4.750 -0.3586 0.01674 0.00914 -0.0403 0.8357 0.5278
-4.500 -0.3340 0.01687 0.00920 -0.0394 0.8278 0.5578
-4.250 -0.3089 0.01709 0.00940 -0.0383 0.8199 0.5811
-4.000 -0.2837 0.01732 0.00953 -0.0373 0.8130 0.6033
-3.750 -0.2591 0.01759 0.00974 -0.0362 0.8049 0.6226
-3.500 -0.2324 0.01779 0.00989 -0.0352 0.7990 0.6350
-3.250 -0.2068 0.01779 0.00980 -0.0348 0.7906 0.6429
-3.000 -0.1798 0.01772 0.00963 -0.0345 0.7843 0.6488
-2.750 -0.1534 0.01758 0.00934 -0.0345 0.7776 0.6561
-2.500 -0.1267 0.01757 0.00928 -0.0341 0.7708 0.6606
-2.250 -0.0993 0.01749 0.00910 -0.0339 0.7655 0.6665
-2.000 -0.0731 0.01739 0.00891 -0.0340 0.7578 0.6731
-1.750 -0.0460 0.01735 0.00882 -0.0337 0.7518 0.6770
-1.500 -0.0187 0.01730 0.00869 -0.0336 0.7463 0.6820
-1.250 0.0080 0.01724 0.00856 -0.0338 0.7396 0.6882
-1.000 0.0354 0.01719 0.00847 -0.0337 0.7345 0.6923
-0.750 0.0626 0.01718 0.00844 -0.0336 0.7294 0.6969
-0.500 0.0892 0.01718 0.00841 -0.0336 0.7229 0.7030
-0.250 0.1167 0.01715 0.00833 -0.0336 0.7178 0.7084
0.000 0.1437 0.01717 0.00835 -0.0334 0.7128 0.7130
0.250 0.1702 0.01721 0.00840 -0.0334 0.7067 0.7188
0.500 0.1978 0.01721 0.00839 -0.0335 0.7021 0.7247
1.000 0.2510 0.01736 0.00859 -0.0333 0.6922 0.7365
1.250 0.2778 0.01742 0.00867 -0.0332 0.6871 0.7427
1.500 0.3052 0.01746 0.00873 -0.0330 0.6832 0.7483
1.750 0.3316 0.01759 0.00891 -0.0330 0.6782 0.7555
2.000 0.3572 0.01773 0.00913 -0.0327 0.6730 0.7612
2.250 0.3840 0.01780 0.00925 -0.0326 0.6687 0.7681
2.500 0.4119 0.01786 0.00934 -0.0326 0.6651 0.7756
2.750 0.4352 0.01810 0.00972 -0.0320 0.6588 0.7825
3.000 0.4620 0.01819 0.00986 -0.0319 0.6535 0.7909
3.250 0.4890 0.01822 0.00996 -0.0315 0.6495 0.7973
3.500 0.5127 0.01847 0.01033 -0.0311 0.6427 0.8062
3.750 0.5377 0.01858 0.01056 -0.0305 0.6373 0.8135
4.000 0.5656 0.01862 0.01066 -0.0303 0.6332 0.8227
4.250 0.5872 0.01887 0.01108 -0.0294 0.6257 0.8309
4.500 0.6140 0.01881 0.01109 -0.0289 0.6184 0.8406
4.750 0.6362 0.01880 0.01122 -0.0276 0.6074 0.8496
5.000 0.6608 0.01859 0.01108 -0.0265 0.5946 0.8595
5.250 0.6843 0.01836 0.01092 -0.0252 0.5794 0.8703
5.500 0.7061 0.01817 0.01083 -0.0236 0.5624 0.8811
5.750 0.7272 0.01802 0.01077 -0.0220 0.5426 0.8934
6.000 0.7486 0.01783 0.01061 -0.0204 0.5183 0.9069
6.250 0.7683 0.01779 0.01065 -0.0187 0.4878 0.9222
6.500 0.7894 0.01787 0.01076 -0.0176 0.4441 0.9392
6.750 0.8098 0.01830 0.01077 -0.0167 0.3524 0.9599
7.000 0.8217 0.01973 0.01155 -0.0159 0.2513 1.0000
7.250 0.8245 0.02113 0.01253 -0.0135 0.1904 1.0000
7.500 0.8297 0.02245 0.01355 -0.0113 0.1456 1.0000
7.750 0.8348 0.02368 0.01456 -0.0091 0.1167 1.0000
8.000 0.8423 0.02489 0.01565 -0.0072 0.0979 1.0000
8.250 0.8506 0.02615 0.01682 -0.0056 0.0857 1.0000
8.500 0.8592 0.02744 0.01805 -0.0042 0.0768 1.0000
8.750 0.8700 0.02865 0.01930 -0.0030 0.0700 1.0000
9.000 0.8782 0.03008 0.02067 -0.0017 0.0651 1.0000
9.250 0.8904 0.03130 0.02196 -0.0007 0.0607 1.0000
9.500 0.9013 0.03264 0.02330 0.0002 0.0568 1.0000
9.750 0.9118 0.03408 0.02474 0.0012 0.0541 1.0000
10.000 0.9253 0.03538 0.02614 0.0020 0.0511 1.0000
10.250 0.9378 0.03675 0.02752 0.0028 0.0485 1.0000
10.750 0.9664 0.03956 0.03041 0.0042 0.0445 1.0000
11.000 0.9827 0.04096 0.03192 0.0048 0.0426 1.0000
11.250 0.9979 0.04241 0.03342 0.0054 0.0410 1.0000
11.500 1.0124 0.04393 0.03491 0.0059 0.0395 1.0000
11.750 1.0284 0.04559 0.03667 0.0064 0.0381 1.0000
12.000 1.0427 0.04741 0.03870 0.0069 0.0367 1.0000
12.250 1.0569 0.04936 0.04084 0.0075 0.0357 1.0000
12.500 1.0690 0.05144 0.04308 0.0079 0.0350 1.0000
12.750 1.0788 0.05367 0.04548 0.0083 0.0342 1.0000
13.000 1.0868 0.05603 0.04798 0.0087 0.0336 1.0000
13.250 1.0930 0.05848 0.05055 0.0089 0.0331 1.0000
13.500 1.0995 0.06107 0.05324 0.0091 0.0325 1.0000
13.750 1.0985 0.06433 0.05672 0.0091 0.0321 1.0000
14.000 1.0871 0.06840 0.06114 0.0088 0.0317 1.0000
14.250 1.0762 0.07270 0.06573 0.0081 0.0315 1.0000
14.500 1.0618 0.07751 0.07082 0.0069 0.0312 1.0000
14.750 1.0432 0.08308 0.07668 0.0050 0.0309 1.0000
15.000 1.0227 0.08931 0.08318 0.0023 0.0308 1.0000
15.250 1.0003 0.09634 0.09046 -0.0012 0.0309 1.0000
15.500 0.9737 0.10474 0.09909 -0.0059 0.0309 1.0000
15.750 0.9429 0.11488 0.10946 -0.0122 0.0312 1.0000
16.000 0.9082 0.12739 0.12216 -0.0204 0.0316 1.0000
16.250 0.8657 0.14429 0.13919 -0.0310 0.0321 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 64-215 AIRFOIL (n64215-il)